Repetition of limit G-factor

[ASIP]

I am an ASIP engineer and we track and help maintaining aircraft service life.

From the structural point of view, the limit load is a load that should occur only once per the aircraft lifetime. Thus, we assume that limit G-factors (say, +9g for F-16 @22,500 lb GW) are achieved rarely. Often repetition of the limit G-factor significantly (could I say dramatically) reduces the aircraft fatigue and service lives.

On the other hand, there are quite a few youtube clips with F-16 HUD reading of 9g. I realize that, during a mission, a pilot may need go that far. I looked at a couple of Flight Manuals and found limit maneuver G-factors but no limitations on their number of occurences.

My question is, do a pilot know that he should not get to the limit G-factor often. Is there a document for the pilot prescribing or limiting the number of limit G occurences?

Thank you.

[Courtney Mil]

F-15 pilots (in the days of the A and C models) were trained to use max g (9) in every fight. That's 4 to 5 times per training sortie in the combat taining phases. From what my oppos on the F-16 tell me it was the same there. That's a lot of counts on 9g in each airframe's lifetime.

[Basil]

Well, I'm just a wee truckie but, on the Jet provost, with Pilots' Notes figures of +6/-3 we frequently took it to +6 and, less frequently, to -3 *.

The JP and, later, the Bücker Jungmann, provided me with the most fun I've ever had in an aeroplane.


*Tried an outside square loop one day. The little Viper wasn't up to the huge drag of pushing -3 and I didn't make it. One did however, push -3.5 and, upon reporting same to old Chiefy in line hut, was told: "That's what it really is but they know you young gentlemen will push things a bit so they tell you it's - 3.0." :}

[Waddo Plumber]

I thought the fatigue life of each type was assessed by assumed usage spectra, which took into account the frequency of reaching different g levels per flying hour. This lead to some issues when the DA granted a life to a new operator on the basis of exchange rates for existing operators. Again, surely fatigue stats for sqns are monitored to optimise usage of individual airframes, so aren’t those stats fed back to individual pilots as in "Prune, you are becoming unaffordable"

[ASIP]

Courtney Mil,
Yes, this should drive the number of +9g counts up immensely.

This doesn't look right from the structural point of view. About F-16 :
"the aircraft was originally designed for 1,000 6G-exceedences per 1,000 flight hours, (or 1 per flight hour). "

6g once per flight meant 8000 FH of design service life. But 9g four - five times per flight... How did they deal with the dropping fatigue life? How many sortites were in those training phases? Were there required periods of low-g flying in between?

Was fatigue tracked at all at that time?

[ASIP]

Waddo Plumber,
This is exactly my impression. Something like a preflight briefing. Nevertheless, I think there should be some formal directives.

We do the stuff you described, only for transport planes. In our bussiness reaching the limit G-factor (+3g) is really rare. I am literally shocked hearing fighter pilots get +9g so often. If one of our planes did the limit G three times within, say, a week, there would be a lots of buzz.

[Bevo]

You are correct in that each aircraft has several usage spectra based on the mission type for multi-mission aircraft. This is then used with the fatigue article to validate the specified airframe life (i.e. 8,000 flight hours). This same spectra is use to define engine time between overhaul. Once the aircraft enters service, if the usage is different than originally forecast the airframe/engine life must be reevaluated and possible structural updates made if the original airframe life is to remain.

An example of this was the original P&W F100 engines were assumed to be used similarly to the engines in the F-4. Once the aircraft entered service it was found that unlike the F-4 where the pilot usually selected afterburner/reheat at the start of a dog fight and left them there; in the F-15, with much more thrust, the engines were cycled in and out of reheat more frequently. This resulted in change overhaul schedules for the engine.

As for ASIP's comment about limit load: if you folks have assumed that the aircraft will see limit load once during its operational life you are in for a very big shock. In most air combat training the pilot will go to limit load several times during a given sortie.

[BEagle]

F-15 pilots (in the days of the A and C models) were trained to use max g (9) in every fight. That's 4 to 5 times per training sortie in the combat taining phases. From what my oppos on the F-16 tell me it was the same there. That's a lot of counts on 9g in each airframe's lifetime.

No wonder the Aggressor Squadron guys at Alconbury were so astonished at the very low G limits we were permitted to use on the F-4 - to eke out the remaining fatigue life.....:\

Taking the tanks off and flying to the proper limits once a year at Deci' wasn't really an acceptable alternative - but that's all we were allowed.

When the Hawk replaced the Hunter at TWU, BwoS were surprised that the aircraft was being flown to higher G values than the Hunter. "You didn't use such high G on the Hunter, why are you doing so on the Hawk?", they asked.

"Because we're fighter pilots and because we can!" was the response. Nevertheless, the Hawk G limit was soon reduced to save fatigue consumption....:(

[Burritto]

ASIP, there is a very big difference between an engineering structural limit load (where some damage to the airframe may (will) result, but it won't fall apart) and the G-limit imposed on pilots. The G-limit is imposed to allow to aircrew to operate up to it, without damaging the airframe. I don't recall seeing a structural load limit published for fighter aircraft - generally if we needed to pull so hard that we might damage the airframe there is probably good reason (such as the ground, missile inbound etc).

[Courtney Mil]

One of the best steps forward in fatigue and load monitoring on the F-15 was the introduction of the OWS - Overload Warning System. This measured gz at valious points and calculated stress in the wings, fuselage, mass items (engine mounts and the like), etc. It took into account fuel weight and stores, rolling g and produced an audio tone to allow the pilot to adjust the amount of pull, to get to and maintain amximum allowable g for any configuration, airpeed, altitude, etc.

If the calculated g limit was exceeded Betty would chout "Warning, Over-g" and we stopped the fight (assuming we were only training). We could then call up a display if the warning and work out if we could continue or go home. (That was due to carrying simulated weapons loads.)

Introducing the system increased the stated g-limits and reduced the number of unknown overloads. Good for everyone.

[CoffmanStarter]

Mind you repeated exposure to High G can play havoc with a chaps good looks ... for example this guy is only in his early twenties :E

http://i851.photobucket.com/albums/ab71/prooner/A694F438-EFF9-4819-BCE1-78A9AC57556F-4025-0000012F49DE39E8.jpg

[Lonewolf_50]

If I may echo Burrito's point, the G limits in the Flight Manual are not the same as the structural load limit cited in the OP.

Example: back in the 80's a friend of mine was instructing in a T-34C. The G limits in the NATOPS were +4.5 / -1. Any G over 4.5 had to be reported to maintenance. His student, coming out of a split S, got a little frisky and pulled hard enough to gray/blackhimself out, and my buddy got a bit gray as he called for and took the controls.

All said and done, 7.3 G put on the aircraft. They reported it, a few bolts were replaced, some skin removed and repaired, and that bird was still flying years later. (At the time, we referred to it as "The Corsair" ... )

Did that one pull materially shorten the FLE for that airframe?

Good question. I don't have the answer to it, but I suspect an engineer at NAVAIR knows, or once knew.

[ASIP]

Thank you gentlemen,

I've got a flavour of the problem.

Still I have difficulties to match the reports of getting to +9g several times in a sortie to the design spectrum (F-15) that shows just 4 times per 1000 hours, and even a severe spectrum, which was obtained or at least matches Structural Data Recording System data, and shows about 40 times per 1000 FH.
It is interesting that there are occurences of larger than +9g (100%) acceleration. It seems like there are different definitions of the limit load factor in structural standards (like MIL-A-83444, JSSG-2006 or FAR-25) and in Flight Manuals where "limit" may refer to a human factor.

[Fox3WheresMyBanana]

Cold War: Probably because I had an engineering degree, I was nominated as Sqn "Fatigue Officer" when a new set of sortie/role restrictions linked to the Fatigue Index of the jet came in. The idea was to find the limits of the sortie FI logging procedure so we could effectively carry on flying the way we were doing. It took me about 3 weeks to come up with a few simple rules, e.g add a couple of GCAs at the end of an ACT sortie to make the sortie length over xx minutes. No lying necessary.
From my inexact understanding of the FI calcs, I reckoned this would knock about a third off the real lifespan of the jet.

The reason this was done was because the squadron was tasked with being Operational, and it would not (in the estimation of my seniors and betters) have been possible to do this and follow the life-extending procedures.

It's pretty simple really. You can't pull 9 'g' for the first time in a real war. Buy jets more often or accept they are less capable because of the restrictions. You can't have less FI usage and keep full operational capability. Same applies to flying hours.
If the politicians and REMFs aren't prepared to accept this, the operational crews will find a way around your "rules". Always have, always will. It just suits us to appear to be knuckle-dragging banana munchers.

p.s. I have a massive 1 hour in the mighty F-15B, and personally pulled 9'g' five times. Oh boy, was it fun!

p.p.s. The Bulldog entered service in 1972 with an official release to service of 4.75'g'. For some unknown reason, the Aircrew Manual and FRCs permitted 5.25 'g' until 1984. Student doing aerobatics were taught to fly most manoeuvres to 5.25 'g' until that year. That was fun too.

[Arm out the window]

I've always been under the (possibly misguided) impression that the published flight manual g-limits are around two-thirds of the calculated g at which structural failure would occur.

In my experience of military flying, flight manual g-limits are regularly flown up to as many times as are required for the training or task being done. You don't go out and go crazy with the g for no good reason (well, mostly...) but nor do you hold back when required.

This then leads to possible mismatches between the designers' assumptions and what is actually done, therefore sometimes reducing the service life of the airframe. Reduced g-limits are sometimes introduced to try and extend service life when problems (eg fatigue cracking etc) are detected.

One case in point is the Pilatus PC-9 in service with the RAAF - as an advanced trainer it naturally copped a flogging, and needed to to get the job done. However, it was soon realised that this would quickly mean the limited fatigue data available from the manufacturer wasn't sufficient to predict what might happen to the airframes long-term (bearing in mind that we Aussies keep types in service for a long time, historically).

A solution was to take a representative airframe off line, stick it on a fatigue rig which basically pushed and pulled at it repetitively for months/years to rapidly simulate the effects of the above-mentioned flogging, thus generating useful data that could then be used to work out a more accurate estimate of airframe life under real-life working conditions.

So - designers design them to a set of assumptions, then they are used within flight manual limits (barring occasional stuff-ups) and fatigue occurs at a rate governed by same.

[Brian Abraham]

Is there a document for the pilot prescribing or limiting the number of limit G occurences?How the RAAF manage the issue on legacy F-18s.

Each Hornet aircraft has a Maintenance Status Data Recording System (MSDRS), which provides information on the aircraft’s structural - fatigue consumption and the aircrew’s flying characteristics during each flight. This data is then processed by the Mission Severity Monitoring Program No. 2 (MSMP2), to provide FLEI data relevant to each mission, and to update the fatigue - life records of each aircraft. Tactical Fighter SPO uses this information to manage individual aircraft against their structural ‐ life limits and to identify aircraft requiring close structural ‐ fatigue consumption management and structural refurbishments. The information is also provided to ACG’s 81 Wing for the purpose of structural ‐ fatigue management, through mission planning and g ‐ force management by pilots. This is to ensure that Fatigue Life Accrual Rates in the Hornet fleet are maintained within agreed limits while maintaining operational effectiveness.

Hornet aircraft have over 70 structures—such as the centre barrel—which are subjected to accumulated stress. Once their accumulated stress exceeds the structure’s specified FLEI or airframe hours flown (AFHRS) limit, they require refurbishment, or the aircraft needs to be withdrawn from service or transitioned to a safety ‐ by ‐ inspection regime. At the time of the audit, the RAAF’s classic Hornets have over 20 structures that are subject to a safety ‐ by ‐ inspection regime. A register of these structures, their accumulated FLEI limits and their airframe hours flown limit is maintained in the Hornet Service Life Limit (SLL) register.

http://www.anao.gov.au/~/media/Files/Audit%20Reports/2012%202013/Audit%20Report%205/201213%20Audit%20Report%20No%205%20OCRed.pdf

[India Four Two]

Brian,

Thanks for the link. I found this paragraph particularly interesting:
4.24 In RAAF service, however, the rate of fatigue accrual has been higher than that of the US Navy. This is due to the RAAF aircraft experiencing higher g and sustained‐g loadings, which have consumed the fatigue life ....



Any idea why that would be?

[Brian Abraham]

One could only assume that the aircraft is being used in a way that the US Navy/Manufacturers never envisaged. Spending more time in nap of the earth, or air combat manoeuvring for example.

[Bevo]

I've always been under the (possibly misguided) impression that the published flight manual g-limits are around two-thirds of the calculated g at which structural failure would occur.
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. It is interesting to note that it’s not just the wings but all mounting that need to be designed to this limit. For example the shelving holding all electronic equipment has to be designed to the same ultimate load factor.

Also it is not just longitudinal load factors that have to be considered. One bad example was the MiG-23 which had a tendency to depart at high angles of attack. On departure the aircraft would rapidly yaw which imposed high lateral loading on the engine for which it was not designed. This often resulted in an aircraft that recovered from departure but with an engine which had failed due to turbine blades having failed as the engine flexed on departure.

[Basil]

It is interesting to note that it’s not just the wings but all mounting that need to be designed to this limit.
A very good point which needs to be remembered by anyone thinking that a low ZFM gives carte blanche to take liberties.

[ASIP]

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.


Technically that is correct. However the life is a little bit more complicated.
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.

[Brian Abraham]

The F-15C disintegrated in 2007 because of a missed crack at 7g only.
14th Mar 2013 21:12A 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.

[ASIP]

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.


The plane had reached higher Gs before without failure and the longeron thickness didn't affect flight performance. The safety factor was still greater than 1. This was the extended crack that reduced the aircraft residual strength well below the limit load. This is mechanics (fracture mechanics, indeed) of this failure. The crack was the immediate cause of the aircraft disintegration.

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...

[Brian Abraham]

From the structural point of view, the limit load is a load that should occur only once per the aircraft lifetime.I could only find one place that makes that statement, on Wiki.

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.

[Lonewolf_50]

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?

[Brian Abraham]

Summation of design criteria here Lonewolf for FAR 25 aircraft.

FAR Structural Design Criteria (http://adg.stanford.edu/aa241/structures/FAR301.html)

[Lonewolf_50]

Thank you Brian, that answers my question. :ok:

[Bevo]

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/pubs/monographs/2006/RAND_MG370.pdf

[ASIP]

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


Thank you Bevo,

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.

[Brian Abraham]

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".What justification can you provide that states otherwise? I've never seen any cautionary note in any flight manual, regulatory directive, or engineering justification that says otherwise. Other than military, and a few civil aerobatic types, which have the necessary instrumentation, pilots don't even know what "g" they are pulling. Any fighter pilot will take great exception to your suggestion that they have a distorted understanding of limit load, that's where they operate all day, every day, every flight - if the opportunity presents itself.

[Bevo]

What Brian said - Agree!!

[Arm out the window]

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.

[ASIP]

Without trying to be rude, it seems you may have a somewhat distorted
understanding of what happens in normal operation of aircraft, ASIP.



Oh, don't worry. I started this thread when discovered I had very little idea about what happens in normal operation of aircraft. And if initially I thought it's a way of doing business in one particular Air Force, now I see, thanks to the sincere help from all of you, gentlemen, that this is an absolutely common thing regardless of a country.

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.

[Brian Abraham]

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.Please do, look forward to you posting ASIP.

[ASIP]

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

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...

This is a correct and wise approach. This is how the system works in air and on ground.
I am upset with statements like

You may take an aircraft to its limit load as often as you wish on any flight.

This is a very typical opinion.

The structure in FAR certification must be able to withstand the ultimate
load for at least three seconds prior to failure.

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 :

This management took the form of aircrews receiving monthly
fatigue reports, which encouraged smooth and appropriate flying, while
maintaining the development of tactical skills.

However, these "reports" don't seem like mandatory requirements.
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...


I am glad we finally came to some common points.
Sorry for may be inconsistent presentation of my humble opinion.

[Arm out the window]

What is written in a Flight Manual (FM) may not be a "true" (again, from my point of view) limit g-factor.

I think this is the crux of the issue - as I mentioned, a sensible pilot (and I think that refers to the vast majority who, sensibly, don't want the aircraft they're flying to fall to pieces around them) will operate the machine to the limits set in the flight manual but not exceed them, and if accidental exceedences occur, report them and ensure appropriate maintenance action is initiated.

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.


An example of this kind of document would be such things as standing orders issued by the responsible operating authorities; for example, on a military advanced trainer I spent some years operating, the flight manual maximum g limit was +7, but command standing instructions reduced that to +4 for most of the time for fatigue management, allowing up to +6 in some special circumstances. Exceedences of those limits required a safety report to be submitted.

[ASIP]

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.

[Thelma Viaduct]

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.

[VinRouge]

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?

[Brian Abraham]

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 following is a quote taken from an aviation textThe 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 airplaneThe following is a typical V-n diagram

http://i101.photobucket.com/albums/m56/babraham227/z198_zps3c91940e.jpg

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.

[Bevo]

Unsymmetrical maneuvers (rolling maneuvers) on fighter aircraft is the one area where pilots are not a careful/aware as they might be. Since there is no “unsymmetrical g-meter” and the limits are normally listed as “over 1/2 lateral stick”. In a maneuvering air-to-air engagement most pilots do not monitor their lateral stick displacement well enough to know if the unsymmetrical limit was exceeded. Fortunately the most modern fighters have flight control computers which will keep the pilot out of trouble.

[Arm out the window]

Scene: WWI RFC maintenance tent:

Lord Flashheart enters holding bent control column.

"Right you lot, get me a new kite pronto or I'll give your wives something to hang their towels on. Woof!"

[ASIP]

The following is a quote taken from an aviation text


Brian, the text is correct. In what context did you post it?

What reference are you able to provide ASIP that operating at the limit load is a no no.
I am afraid you are missing my point. Never said I that going to the limit g-factor (limit load) "is a no-no". I did say that the number of times of reaching this load must be limited.


We all recognise that pulling "g" impacts fatigue life,

I am glad to hear this. However I have some doubts. Somehow you always point out to aircraft static strength only. And you do that incorrectly. Your V-n diagram is just a design tool and does not correspont to the actual situation with strength of the aircraft operated under the Damage Tolerance principle. You never mention fatigue and Damage Tolerance as an aircraft design criteria.


so please correct our, or I should say, mine, distorted thinking.


Under distorted thinking (again, admittedly dramatizing) I meant the conviction that the limit g-factor can be pulled up in every flight any arbitrary number of times.
Disasters are avoided not because it is intristically safe, but rather because there is a numerous army of specialists tracking the problem continuously.
As some reference I can give a link to a paper about F-16 service life problems:
http://www.dtic.mil/cgi-bin/GetTRDoc?AD=ADA407261

Pictures follow.

[ASIP]

This was original plans for F-16 service life.
http://i1339.photobucket.com/albums/o701/ASIP2000/F16_1_zpsb585147c.jpg

This is comparison of the design and actual usage. Somebody pulled high g's too often.
http://i1339.photobucket.com/albums/o701/ASIP2000/F16_2_zps9e7b8e0e.jpg

And this is the result.
http://i1339.photobucket.com/albums/o701/ASIP2000/Untitled_zpse6814b88.jpg

If the fleet would have been kept flying without premature part replacement up to the design 8000 FH, some aircraft could have disintegrated.

[Schnowzer]

Bevo,
The F-15E had an "overload warning system" that varied the amount of g you could pull depending on the speed the aircraft was flying to avoid serious damage whilst rolling in the transonic regime. Most "over-Gs" where called at very low g-load, 3-4g whilst transonic and rolling, pretty annoying.

ASIP,
Despite aiming at 9g in an F15, most pulls by competent pilots would be in the mid to high 8s, we used to pull, wait till the jet howled and then back stick it. In a DBFM training sortie there would probably be 5 or 6 splits each with 3 or 4 pulls to between 7 and 9g and I never saw any bits fall off.

JC was in the back of a clean jet that chucked a bit of the stab off at M2.4 and I lost a Winder in a CB but that was just careless:ok:

[Courtney Mil]

Schnowzer,

Agreed. See posts 2 and 10.

[ASIP]

Thank you Schnowzer for the objective information.

There is no much information available on actual +9g occurences. What I found, though, indicates that the real number is not large, at least far shy of "pulling 9g four-five times a flight".
These are data from a presentation on the F-15 Life Extension Program
http://i1339.photobucket.com/albums/o701/ASIP2000/F-15SL_zps006e5cc1.jpg

Spectrum FTA1 is the design usage.
Spectrum FTA6 is an actual or at least close to actual usage, confirmed by the on-board Structural Data Recording System.
Spectrum FTA7 is a design spectrum for future severe usage.
http://i1339.photobucket.com/albums/o701/ASIP2000/F-15Spectrum_zps07620972.jpg

It can be seen that the original design usage predicted 3 exceedances of limit load per 1000 FH. In reality, the number was 42 exceedances. This means one +9g or higher maneuver in approximately 250 - 300 flights.
Still, McAir guys consider this too many. They want you to pull 9s no often than 20 times per 1000 FH.

In a DBFM training sortie there would probably be 5 or 6 splits each with 3 or 4 pulls to between 7 and 9g and I never saw any bits fall off.
Three things here:
1. You did everything properly.
2. Guys in St. Louis new their business well. Stresses in the aircraft parts were well below FTU/1.5 plus they coldworked fastener holes and didn't accounted for that in "certification".
3. Fatigue damage is proportional to the fourth power of stress in aluminum parts (approximately). Thus 7.5g develops only one half of the fatigue damage compared to 9g.

[Brian Abraham]

Your V-n diagram is just a design toolBeg to differ, as much as it is a design tool, it is also promulgated to pilots as to the limits in which they are obliged to operate. They only cautionary note I've ever seen as to approaching any limit is that between Vno and Vne, which states "in calm air only and then with caution".I did say that the number of times of reaching this load must be limitedThere is no mention in any text as to how often a limit load may be applied, and certainly not in any flight manual. Nor does any flight manual or regulation require reporting a of limit load being reached. Not that the pilot of most non military aircraft would know in any event, non having the necessary instrumentation.

The FAA was obliged to issue an AD on the S-2 series of aircraft due to cracking in the upper longerons due to overstressing. The AD required the fitting of an accelerometer marked with the limit loads, so that pilots did not exceed those limits. And absolutely no mention of recording the number of times those limits were reached.You never mention fatigue and Damage Tolerance as an aircraft design criteria.It's not a subject pilots are interested in. They fly the aircraft in accordance with the rules, regulations, and standard operating procedures laid down. Should untoward fatigue be detected it may be because of any number things,

1. Design - King Air forward spar, modification of strap on lower spar cap available
2. King Air - Pilor performed aerobatics in non aerobatic aircraft. Repaired and returned to service.
727 - Loss of control. Pulled 6 g (limit 2.5). Repaired and returned to service.
747 - Loss of control. Data dropout precluded record of max g but 5.1 recorded. Repaired and returned to service.

Should the evidence of excessive fatigue make its self known the operator will institute procedures to contain the problem.

A 727 is used for Zero g flights in which they operate between 1.8 and 0 g (limit g 2.5 to -1) and those g are pulled multiple times per hour. Are they exceeding fatigue limits? Certainly no airline aircraft ever was subjected to those limits. What limits do you suggest they should be applying?

[Waddo Plumber]

It's a while now since I was in fatigue monitoring, but IIRC, the normal SN curve is almost asymtotic to the N axis, and as long as S (or a combination of major and minor cycles) is below the asymptotic value, it can be repeated an infinite number of times. So, the question is what are the maximum values of + or - G that can be pulled without exceeding that stress. Up to those levels, there is no need to tell the aircrew to back off. As aircraft age, it's normal for the G limits to reduce specifically for that purpose.

[ASIP]

There is no mention in any text as to how often a limit load may be applied,
and certainly not in any flight manual.

We already know this. This is a fact.
However, if you use it as a proof that the number of very high g-maneuvers is unlimited, you are not right.
I hope we have established that it is simply not the pilot's business to maintain this number within acceptable limits.
If you are not aware of a problem, this doesn't mean the problem does not exist. Maintaining the aircarft structural integrity, which can be compromized by frequent high-g maneuvers, is performed by technical personnel and relevant documentation flow through technical channels.


Nor does any flight manual or regulation require reporting a of limit load
being reached. Not that the pilot of most non military aircraft would know in any event, non having the necessary instrumentation.


See, you are now trying to deny a known procedure for mandatory inspection of an aircraft that achieved its limit g-factor. Your claim about absence of such inspection in a FM is absolutely correct. Regulations? I am not sure. Are you familiar with the documentation required for certification for FAR25.571 and FAR25.1529? All those design spectra, numbers of occurences, inspections must be there. A similar stuff exists for military aircraft.
I hardly doubt the pilot caused a post-limit load inspection remains uninformed after the buzz accompanying such an inspection and an engineering report.


It's not a subject pilots are interested in. They fly the aircraft in
accordance with the rules, regulations, and standard operating procedures laid down.


Completely agree. However this is not a reason for cancellation of natural processes in the airframe material.

Should the evidence of excessive fatigue make its self known the operator will institute procedures to contain the problem.

Great! We are talking about the same things. But this still doesn't make the number of high-g maneuvers that the aircraft can sustain unlimited.

A 727 is used for Zero g flights in which they operate between 1.8 and 0 g (limit g 2.5 to -1) and those g are pulled multiple times per hour. Are they exceeding fatigue limits? Certainly no airline aircraft ever was subjected to those limits. What limits do you suggest they should be applying?

I am not sure I understand you correctly. What do you mean under the fatigue limit? If it is a material property, then yes, at a level flight the main structural components of a transport aircraft experience stresses above the material fatigue limit. Transport aircraft can do maneuvers up to +2 g much many more times than fighters can pull to their +9g. The reason is that the transport A/C are mostly affected by an enormous number of turbulence cycles, and they are designed for that. For fighters, the turbulence normally is not a factor. Until they (mostly strike aircraft) fly low and fast.

[ASIP]

It's a while now since I was in fatigue monitoring, but IIRC, the normal SN curve is almost asymtotic to the N axis, and as long as S (or a combination of major and minor cycles) is below the asymptotic value, it can be repeated an infinite number of times.

Oh! I glad to hear this kind of words. That's correct. Such a design is called "boiler-proof." Normally it cannot be used for airplanes because the structure is way too heavy. Nevertheless, for fighters in a level flight in many components the stresses are below the material fatigue or endurance limit.


So, the question is what are the maximum values of + or - G that can be
pulled without exceeding that stress.

Absolutely! The devil is in details.
And we know the answer. The local stresses in fastener holes or notches will be many times higher than "that stress". They will be way above the material yield limit, approaching or even exceeding FTU.


Up to those levels, there is no need to tell the aircrew to back off. As
aircraft age, it's normal for the G limits to reduce specifically for that
purpose.


Unfortunately, "those levels" are exceeded in every flight. This thread was started because Flight Manual are silent about this problem. To give you an idea about stress levels I show the picture for approximately F-15 situation. I hope you know the difference between "nominal" (reference, remote) stress and "local" (peak) stress.

http://i1339.photobucket.com/albums/o701/ASIP2000/7075T6SN_zpsf47b1a36.jpg

[Waddo Plumber]

ASIP (Aircraft Structural Integrity Programme?). I've woken with a heavy cold. I'll. reply when my brain answers to the helm.

[E L Whisty]

I dimly recall a tale told by an US Marine Corps friend about a fatigue survey performed by the US Navy on their fleet of F 14s, sometime late 70's or early 80's. The story was that they put strain gauges on a representative fraction of the fleet and had some very disturbing results.

It is important to understand the difference between stress and strain in the engineering definitions.

As I understand it, there was a great deal of fatigue being accumulated under rolling G conditions that was never appreciated with simple G counters then in use.

It is also important to appreciate that metals have been characterised as those which have fatigue limits and those that have fatigue lives. As I retired from the relevant research some 10 years ago, I do not know what the current thinking is, however, there was a growing body of evidence that even steel alloys have limited fatigue lives.

A fatigue limit is, simplistically, a loading limit which, if not exceeded, will give no accumulating damage.

A fatigue life is a description of how a material (thence structure) will become progressively damaged as particular loads are applied, cyclically, over time. As an simple illustration, a single 8G pull might cause as much damage as several hundred applications of 2G. Thus, accumulation of damage is complex to predict and calculate.

Aluminium alloys, however, have always been considered to accumulate damage at even the lowest loadings. So, an in service G limit for aircrew is imposed in order to preserve the service life of an airframe. It is acknowledged that damage will be accumulated and airframe integrity issues will arise as it ages. Older aviators will recall the sorry state of our F4s after they had had 15 years of enthusiastic aerial gymnastics.

As a back seater in F4s, I can recall very few 'dogfights' when G was not applied whilst rolling. Indeed, doing so was a necessity for an aggressive pilot. The G counter did not betray the angel faced pretence of cunning pilots that they never rolled whilst pulling!

It is my opinion that we operated aircraft for many years with naivety about fatigue that probably resulted in loss of aeroplanes and of friends.

The future looks good, though, as strain gauges can be incorporated quite cheaply into composite structures and data gathered constantly will make us better able to monitor the health of airframes.