Number of blades in final stage of LEAP turbine

[amsm01]

Oddly specific question I know but would anyone be able to tell me (or provide a reference for) the number of turbine blades on the final stage of the LEAP-1A, which has 7 low px turbine stages; and the LEAP-1B, which has 5 of them - so I’m after the 7th and 5th stage respectively. I understand they’ve different amounts, presumably due to differing thrust ratings and engine sizes but I’ve drawn a blank trying to find dimensions of these things or the number of blades. Any bright ideas?

[aerolearner]

I can help with these, but I understand it is not much...
The information is there, but you need a large screen, good eyesight and a lot of patience!

LEAP-1A
https://downloads.regulations.gov/FA...tachment_3.pdf

LEAP-1B
https://downloads.regulations.gov/FA...tachment_3.pdf

[amsm01]

I have all three and thank you aerolearner! This is exactly the sort of diagram that I was trying to find - I’ll get to work on this tomorrow and my thanks again!

[aerolearner]

An update about LEAP-1A. Both sources say that each of the 7 LPT stages has 147 blades.

https://www.scribd.com/document/705213920/21665565
https://www.scribd.com/document/4363...-Inspection-V0

[amsm01]

How curious, this makes the diagrams even more useful. Must thank you again, this is really such useful reference material - no way I’d have ever dug this up myself so I very much owe you a drink aerolearner!

[amsm01]

On a related subject prompted by the part lifetime docs posted by aerolearner in the initial reply: they contain cycle lifetimes for various elements within the turbine - spool, disk etc - but not the blades themselves. Therefore is it correct to understand these are not life-limited parts as such? I found this a bit surprising given they’ll be exposed to the same high centrifugal forces and heat or is this mitigated by the fact they’re inspected and x-rayed whereas the larger components listed in the doc can’t be?

[tdracer]

Originally Posted by amsm01

On a related subject prompted by the part lifetime docs posted by aerolearner in the initial reply: they contain cycle lifetimes for various elements within the turbine - spool, disk etc - but not the blades themselves. Therefore is it correct to understand these are not life-limited parts as such? I found this a bit surprising given they’ll be exposed to the same high centrifugal forces and heat or is this mitigated by the fact they’re inspected and x-rayed whereas the larger components listed in the doc can’t be?

Breaking a turbine blade is not a big deal - high vibes and a probable engine shutdown, but unlikely to threaten continued safe flight and landing. Bursting a turbine disk is an uncontained engine failure and potentially catastrophic (think Qantas 32) - hence the discs are life-limited parts.
The blades are removed and refurbished on a regular basis (although LP turbine blades are not that highly stressed compared to the HP blades) - but are not considered 'life limited' parts. Of course, turbine blades do get scrapped (and they do track hours and cycles to determine the refurbishment intervales), but as long as they pass inspection during refurbishment they can be re-used indefinitely.

[amsm01]

Thank you tdracer, really appreciate the insight and given your explanation, it certainly makes sense why they aren’t categorised as LLPs - you learn something new every day! And in reply to aerolearner, having counted the blades in the 1B I came up with 109 - and 146 in the 1A so it’s probably not too far off the correct number given the double sources saying 147 in the 1A. Thanks both again

[megan]

Originally posted bt tdracer
Breaking a turbine blade is not a big deal - high vibes and a probable engine shutdown, but unlikely to threaten continued safe flight and landing

May be with airline equipment td but certainly not the case with all turbines.

Analysis of two failures I've had with Turbomeca products. In both cases all the blades on the disc had been trashed. Operator had four similar failures that I'm aware of. Following lifted from the safety authority report.

Occurrence report 200100584

On 7 February 2001, a Sikorsky S76C helicopter belonging to the same operator, with two crew and ten passengers on-board, was in a hover with the flight crew completing before take-off checklist items. The pilot reported that while trimming the engines, a "pop" was heard. He then noted that the left engine turbine gas temperature (measured at point T4 within the engine) was in excess of 1000 degrees C. The helicopter was then landed uneventfully. The flight crew reported that the only cockpit indication of imminent failure was the almost simultaneous illumination of the left engine chip (magnetic particle) detector advisory.

Examination of the helicopter revealed minor shrapnel damage to the left engine exhaust extension and engine cowling. There was no reported engine fire. The left engine was removed and sent to the engine manufacturer for disassembly and examination. The manufacturer's final report noted a separation of turbine blade number six of the GG second stage disc. The blade was separated above the 'fir tree' attachment point but below the blade platform, and had punctured the second stage nozzle guide vane turbine ring. One adjacent blade (number seven) in the direction of turbine wheel rotation was also noted as cracked.

Metallurgical examination by the manufacturer attributed the blade failure to a low-cycle fatigue cracking mechanism. The manufacturer concluded that abnormal loading was the major contributing factor in the failure, given the reported absence of anomalous material features or evidence of high-temperature operation. Dimensional inspections failed to reveal any sign of non-conformity that could have led to the development of the abnormal loads. However, the manufacturer stated that turbine blade platform/GG disc interferences were also a potential factor that could have aggravated the fatigue failure of the blade.

At the time of the occurrence, Arriel 1S1 engine, serial number 15522, had accumulated 4,737.4 hours and 4,471 cycles since new. It had accumulated 1,740.0 hours TSO and 1,615 cycles since overhaul. Following overhaul, the engine was installed on March 11, 1999. Module three did not have turbine blade plasma coating modification TU204 incorporated.


Occurrence report 199602839

On 9 September 1996, a Sikorsky S76C helicopter belonging to the same operator, experienced an in-flight engine failure of the right engine while taking off from an oil platform. A loud noise was heard before the engine failure. (My note, there was no noise prior to the failure, the failure occurred with a massive explosion, akin to a grenade exploding - have experience with those as well) The right engine was shut down and the crew completed an uneventful single engine return to the Longford base. There was no reported associated engine fire. The right engine was removed and sent to the manufacturer for disassembly examination.

At the time of the occurrence, Arriel 1S1 engine serial number 15513, had accumulated 2,282.0 hours and 1,949 cycles since new. The manufacturer provided the operator with a final report noting the rupture (separation) of one GG turbine blade with subsequent rear bearing damage and GG seizure. Their report stated that the separation was suspected to be the result of blade rubbing with the second stage nozzle guide vanes with no signs of fatigue or abnormal over temperature operation. Module three had turbine blade plasma coating modification TU204 incorporated.

[Someone Somewhere]

I think those cases broadly line up with tdr's statement: they were contained engine failures that didn't impact the aircraft beyond the unexpected loss of power.

Not like a turbine or fan disc failing and tearing through the airframe.

That's not to say that engine failure can't impact safe flight, especially on a helicopter, but it's still much less likely than a disk failure.

To throw a very rough guess out there, are we at about 200k hours per IFSD but maybe 20 million hours per uncontained rotor failure? Couldn't find figures for the latter newer than 1980s on a casual search.

[megan]

The first failure I had the shrapnel was contained, in the second a lot of shrapnel exited tangentially, which could have had implications for the tail rotor drive as the shaft ran between the two engines.

One S-76 with Allison engines had a engine failure that severed the tail rotor drive shaft, severed the battery connection, and as result of the other engine being taken out as well, total electrical failure. We flew with a mod we called the BBQ plates, weighed some hundreds of pounds to contain the shrapnel until the engines could be modded.

[tdracer]

Originally Posted by megan

The first failure I had the shrapnel was contained, in the second a lot of shrapnel exited tangentially, which could have had implications for the tail rotor drive as the shaft ran between the two engines.

One S-76 with Allison engines had a engine failure that severed the tail rotor drive shaft, severed the battery connection, and as result of the other engine being taken out as well, total electrical failure. We flew with a mod we called the BBQ plates, weighed some hundreds of pounds to contain the shrapnel until the engines could be modded.

Perhaps the standards are different for turboshaft engines used on helicopters, but it is a cert requirement for large commercial aircraft turbofan engines that the turbine case retains debris from failed turbine blades (same with the compressor case).
In addition, rotor burst analysis do both 'high energy' and 'low energy' debris and what shielding is needed (and can be taken credit for) for the 'low energy' debris - and shed compressor and turbine blades are considered 'low energy' debris (fan blades are treated differently for obvious reason).
OTOH, burst discs are considered to have infinite energy and no shielding credit can be taken - a burst disc is going to go wherever it's going to go and take out anything in its way. The rotor burst analysis is commonly known as a "1 in 20", since it's assumed that 5% of uncontained disc failures will be catastrophic.

[wrench1]

Originally Posted by tdracer

Perhaps the standards are different for turboshaft engines used on helicopters,

No difference in standards between the two, but more a point of when those engines were certified. There was a revision to the Part 33 turbine engine containment rules in the 1980s which were after the certification date of most turboshaft engines of that time. In the case of the S-76 C30 engine mentioned above, when the AD was released as a result of that incident, those engine models were upgraded to meet the newer containment rules with the installation of an improved internal containment ring. The “BBQ plates” or protective shields, were the initial temporary fix of the issue. One positive outcome from using those external plates/shields was that as the engines were upgraded with the permanent rings, any mechanic who wanted got a free set of custom Inconel rivet bucking bars.