NWA,
You are right to say that changes in load can have an effect on the progress of a fatigue crack. And you rightly note that the maximum axial load on the engine is at the front where the fan is since it is the part which is producing most of the thrust. I had a feeling these big fans which are really ducted propellors would have a big influence on the stresses at the front of the engine. That fan is literally dragging the whole aeroplane behind it. We had it easier on the low bypass engines like the Conway (0.6 bypass ratio). But what do I know - it is a long time since I designed the turbines and thrust reversers on the Conway and I left RR before the RB211 debacle. Stewart Miller the man who rescued RR from that catastrophe worked on the next board to me in those long gone days.

scfi
Unless you have one of these bolts in your hand you cannot know if it failed in tension or from fatigue. The fan assembly will have been tested rigorously on a development rig over many hundreds of cycles before the design was signed off for production. Now we know this engine is a 'bitser' (bits of one design combined with bits from another design) but I have no doubt that the combination was also tested through many cycles before being signed off. What I am getting at here is that this failure seems to me to be a unique event, and may never occur again. So all the AD's in the world are not going to make much difference. But as I have already said we have no information at all right now on this forum about which bits are actually available for inspection. Till we have....

NWA,
Components can be designed to operate at below fatigue inducing stress levels and, therefore, not fatigue fail. However, where circumstances require, components that have to operate inside their fatigue vulnerability are given a "safe life" that requires replacement before likely failure. Unfortunately, the combination of defects or unplanned loads can corrupt both strategies of component integrity. Consequently, the fact that the Fan hub has obviously failed on the remains of this engine does not confirm the type of failure or, even that the Fan hub is the cause of the incident.:8

@scifi and @olympia463 :

the attachment ring you see is actually the front of the low pressure compressor which hangs onto the fan as far as i understand it.
the part in the middle that is sheared off is the connection of the shaft connecting to the fan.

so the shaft connection there needs to carry the torque and pull from the low pressure compressor as well as the fan.

see various depictions / diagrams in the thread.

The thrust loads across a fan disk are insignificant to the loads caused by spinning.

Vibratory stresses are typically way below any measurable contribution to the tangential stresses in a fan disk.

The materials used in the fan disk and the design stresses are very resistant to crack growth in a few cycles.

Whatever happened is likely way outside simple assumptions.

When you design a bolt flange attachment in a high stress area you have to avoid concentrating stress fields around only a few bolt holes. You do this by adding extra holes (typically not used) to smooth out the stress field thus avoiding fatigue fatigue cracking from these holes.

lomapaseo,

You are confusing 'steady' stress such as that due to centrifugal forces on fan blades and disc from rotation, with 'varying' stress (ie stress which alternates) such as might arise in non rotating parts like casings etc. A steady stress which is below the safe stress for the material is quite different to an alternating stress which will cause failure if the cycles reach the fatigue limit for the material. I think we may need to ask people to state their engineering qualifications before entering this discussion. pm me and I'll tell you mine if you are interested.

AD = 2 hour inspection when hub assembly failed at less than 25% of life limit cycles per previously linked article?
http://www.iba.aero/wp-content/uploa...e-May-2017.pdf


Similar fan disc failure some of us may remember.
Lessons Learned

Olympia 463

Nope, not confused.

There is such a thing as a "Goodman diagram" (combined steady stress and alternating stress) which applies equally to rotating parts as well as stationary parts. I simply put them in perspective as to how rare it was to combine them in a fan disk failure unless something was seriously amiss.

As a propulsion technician of over 40 years maintaining and overhauling gas turbine engines, I am used to boroscoping various parts of the engine. In my experience, some defects/cracks propagate quite slowly and can be monitored safely... i.e. combustion chamber liners and HPNGV vanes. Other defects, particularly on rotating components fail within a very short time (a few hours) of the defect becoming visible to the naked eye.

I would have thought that a crack on an LP fan disk would go from "visible" to "failure" very quickly. Perhaps someone better qualified in that field could comment on that? I'm guessing that the mandated inspection of the engines must be NDT rather than boroscope or visual to be of any value?

It looks like the AD is available here: http://rgl.faa.gov/Regulatory_and_Gu..._Emergency.pdf

The "AD Requirements" section says: "This AD requires a one-time visual inspection of the GP7200 series engine fan hub, with the compliance time based on the number of accumulated flight cycles, and removal of the fan hub if damage or defects are found that are outside of serviceable limits."

The "AD Requirements" section says: "This AD requires a one-time visual inspection of the GP7200 series engine fan hub, with the compliance time based on the number of accumulated flight cycles, and removal of the fan hub if damage or defects are found that are outside of serviceable limits."[/quote]

Yes. However, the visual inspection is to be performed iaw a proceedure that will detail the exact method used.
Also worth noting:
"Interim Action
We consider this AD interim action. An investigation to determine the cause of the failure is
on-going and we may consider additional rulemaking if final action is identified". And:
"FAA’s Determination
We are issuing this AD because we evaluated all the relevant information and determined the
unsafe condition described previously is likely to exist or develop in other products of the same type
design".

My interpretation of this is that, the FAA think there may be further failures.:uhoh:

Is the GP7200 engine (as fitted to the AF A380, if there are variations) typically transported with fan attached, or with the fan detached from the propulsion core and re-attached at the destination?

What size of aircraft is required to transport the fan and the fairing pieces, as well as the core? Are AF going to have to wait a while for an oversize cargo aircraft to be available to take the replacement (ballast :) ) engine to Canada?

(A quick look at some sales documents implies that the engine with fan will not fit through the doors of a 747-8F but will fit into an An-124.)

Perhaps AirBus will use their Beluga, can not find anything published re the range of that aircraft though.
The Airbus A300-600ST (Super Transporter) or Beluga, is a version of the standard A300-600 wide-body airliner modified to carry aircraft parts and oversized cargo. It received the official name of Super Transporter early on; however, the name Beluga, a whale it resembles,[2][3] gained popularity and has since been officially adopted

Does the a380 have any provision for 5 engine ferry flights (as the 747 does)?

No. Nor do the majority of 747s, come to that.

Normal transportation without Fan case , same as GE 90-115

My fracture mechanics training came primarily from Dr. DW Hoeppner who had recently moved from Lockheed where he had been involved in fan disk failures on L1011s equipped with RB211 engines. He started off our first class by showing several airplane and helicopter components that had failed in crashes involving multiple loss of lives. Gets your attention and builds a focus on material selection. At that time, 1974, I remember it being a given that materials were going to fail. The focus was to determine life expectancy to that failure point, material selection matching the requirements, and inspections with the ability to find a crack prior to failure. In the ensuing 40+ years materials may have been developed that will have infinite life. Age and retirement have me out of the loop - note my name reference to an airline of the past. However this failure takes me back to Dr. Hoeppner's lectures and avoiding rush to judgement. One of our lectures was a few days after the DC-10 crash near Senlis, France. He came to class carrying a DC-10 cargo door latch, jumped on the desk and yelled at us how some engineer's mistaken material selection resulted in the death of 346 people. Illustration of rush to judgement. The latch was the failure mode but it was due to poor system design, not material failure.

My rush to judgement on this failure is there is a material inclusion not properly detected initiating a premature fatigue crack that precipitated the front falling off. Having learned from Dr. Hoeppner that this strategic part will have a fault tolerant material, the crack could grow quite large prior to failure. This would result in an AD requiring inspections at intervals where it is felt any crack will be discovered prior to becoming critical size that will result in the front falling off.

SLF coming out of lurk mode to share some info about the ranges of a Beluga and an An-124.

Purely FWIW:

NOT from the Airbus website, which left the range blank:
According to airliners.net:
Beluiga -- A300-600ST -- range with a 30 tonne (66,150lb) payload 4000km (2160nm)

According to Wikipedia:
An-124-100 range:
10 tons of cargo= 14,100 km (7,613 nmi)
20 tons of cargo= 13,250 km (7,154 nmi)

Аn-124-100М-150 range:
40 tons of cargo= 11,900 km (6,425 nmi)

55 An-124 have been made.

According to a PDF on the "mtu.de" website, a GP7000 series engine weighs 13,400 lbs without the housing.

Looks as though they would need to use an AN-124 to ferry the ballast engine to Goose Bay.

Has the luggage been retrieved and sent to the pax?

I'm not saying that your judgement is wrong, but considering the fracture toughness of a titanium fan disk you need a "huge" detectable defect to have progressed to start-stop cyclic failure this early in life.

Other possibilities are still open and we as readers have no new facts to go on.

I'm waiting for a hint on the RPM in the last sec trace and better pics of the attachments.

How will AF rescue the aircraft from Wired

https://www.wired.com/story/how-air-...red-engine/amp