AW139 lost tail taxying DOH [Archive]

airwave45

26th Aug 2009, 23:26

:ugh:

Helispanner, I _did not_ say the "tail fell off"
I said

AW139 lost tail taxying DOH

(which, having seen the picture isn't a million miles out)

Where is the weak point in the boom structure? why did this one fail there? (presumably under fairly light loading) bearing in mind the earlier tail strike.

Given the usual culture in the gulf, there probably will be two souls looking for 139 slots in the near future (for no good reason)

Epiphany, I did eat what remained of my hat, I took your current position into account first and barbecued it . . .(nowt but dust:ok:) cof cof splutter.

Cheers for the info gents, LMSLB . . .

The companies I work for have suspended flights on 139's for the forseeable (so I'm told) Shell / Conoco pending shareholder co review.
I believe GH grounded the 139 fleet for, ooh, minutes and minutes (well, 'till the next flight was due out anyway)

Drayman

27th Aug 2009, 10:31

EASA AD No.2008-0157 was issued following reports by operators of debonding of fuselage tail boom panels and requires repetitive inspection of these panels of certain part numbers of tail boom.This is covered by Augusta Bollettino Technico 139-134.To quote´This condition, if not detected and corrected, could lead to the structural collapse of the tail boom installation, resulting in loss of control of the helicopter`.I do not know if this helicopter had a tail boom that came within this AD but looking at the photo this failure comes in area 2 of the inspection sequence.Not withstanding the fact that the tailboom had a repair following the previous tail strike. I am eagerly awaiting the report into the cause of this incident that so easily could have been a tragedy.

iuk1963

27th Aug 2009, 10:38

quote frome MALABO: "A few have crashed..."

prothotype N.1 was lost few years ago and last year the one in Abu Dhabi (apparently no machine fault but the result of the investigation is still missing).

2much

27th Aug 2009, 12:55

These are not rumors the tailboom tore off just aft of the longerons. I have pictures but don't know how to paste them. There has been many tail boom delaminations with the 139 tailboom in the past and an BT was issued to carry out this inspection just about exactly where the boom tore from. All operators are anxiously awaiting the investigation from Agusta who are on site now.

[email protected]

27th Aug 2009, 13:31

Now where is 'lost at sea' when you need him to tell you what a wonderful aircraft this is?

Westlands have always been good at shielding operators in one nation from info regarding failures in another and sometimes operators within the same nation! Why should AW be any different?

A structural failure following a factory authorised repair!!!!!!!!WTF!!

SASless

27th Aug 2009, 14:06

To think it was repaired to save the expense of replacing it!:uhoh:

TheOldTimer

27th Aug 2009, 14:14

Crab, the 139 is a good machine. This event I am sure is not as simple as it may seem. Composite materials react in different ways to metal structures (Sea King) and I am sure that any advisory from Agusta regarding the previous damage and the authorised repair was good at the time. The investigation and report hopefully available in the wide world will benefit all and give valuable information on what is a new technology within the rotary world and in the future benefit the whole industry.
So in my humble view, uninformed comments like WTF, are not constructive and certainly do not add to what has previously been an informative thread.

TheOldTimer

27th Aug 2009, 14:19

SaSless Was it, maybe you could explain your comment on that????

hobiecat

27th Aug 2009, 17:04

I'm curious with your thoughts why ADA in Abu Dhabi has 5-6 "NEW" tail booms in for repair for debonding, some with less than 20 hours on them out of the factory? Granted this tailboom of Gulfs was repaired after a tail strike by Agusta. Something just doesn't seem right. I bet the directors and enginnering at Agusta are :ugh: :ugh:right about now.

KLALA

27th Aug 2009, 17:58

Gulf Has 3 Hel With Debonding Tail Boom All Of Them In The Same Location And Have Been Repaired By Augosta,coz They Under Wranty,and For This One Is Twisted Not Fail Off Like You Said Gayes.the Tail Boom Has Beem Twisted Then Fail Off On The Right Side Of The Helicopter,i Think From My Opnion The Pilot Look The Nose Wheel Insted To Unlook ,whill Checking The Chk List Befor Taxing

TukTuk BoomBoom

27th Aug 2009, 18:42

.......Yes, and let that be the last time someone posts a reply using text-speak....

Mike C

27th Aug 2009, 18:57

KLALA,

I think you should be very careful of your opinions.
I have inspected the concrete exactly where it happened (tail rotor slash marks)very very carefully, I know exactly where he taxied from, he had done about 90 degrees of turning to the left(high tr pitch setting) in a relatively tight turn, maybe 20 foot radius and IF the nose wheel had been locked then there is absolutely no scuff marks at all, and if the nose wheel had been locked I know from experience when I was new on type that he would not have been able to turn that tight. I know you work here and if you had taken the trouble to check, you would have come to the same conclusion.The concrete is quite white here and these marks show up very well.

Are we to assume that if we leave the nosewheel lock in, in a left hand turn the tail "falls off", I hope not, BUT be careful on that left pedal guys, you never know!!!

I know in a left hand turn there could be up to 200HP(thrust and moment arm,you work it out) pushing against a locked nosewheel(full pedal). Much higher stresses than in actual flight I would say. The torn aluminum alloy skin's looked awful thin to my untrained eye, and must admit to a little apprehension now.

I can only assume the Italians have done the maths.

The Captain concerned from my experience of a number of years is a very conscientious and honest guy, I have seen lots of offshore/shuttle shortcuts on checklists done a few myself I am afraid, but the pretaxy/take checks from my observation here are always done, and feel positive that he and his Captain copilot would do the same, unlike the previous incident with this machine on a helideck.

Regards

Mick.

IntheTin

27th Aug 2009, 18:58

KLALA, from reading your post, and at the same time trying to understand your obvious limited English, am I understanding that you think the pilots were at fault here?
I can tell you that both Captains are very experienced. They had turned 90 degrees to the left from the pad they were on when this happened. If the nose wheel had been locked then they couldn't have done this.

My Opnion The Pilot Look The Nose Wheel Insted To Unlook ,whill Checking The Chk List Befor Taxing

[email protected]

27th Aug 2009, 19:27

Theoldtimer - strangely enough, just because I fly the Sea King doesn't mean I am ignorant of composite materials - I went on a very interesting university course a few years ago all about composite materials in aerospace structures including visits to industry to see the production and autoclave process in action. I did learn that whilst composite materials are generally lighter and stronger (in compression) than aluminium, any repair process, unless perfectly carried out in ideal conditions, is susceptible to cracking, debonding and delamination and that impact damage to such structures is usually fairly terminal since composites tend to crack and shatter instead of dent and deform like aluminium.

So, since AW are clearly well aware of such issues, one would have thought that the highest level of scrutiny would have been applied to a repair carried out on a vital structure like the tail boom which had suffered a previous tail strike. This clearly did not happen or the tail structure would not have collapsed in the way the pictures show - hence my comment 'WTF' whilst not being constructive is, I think, appropriate since the failure really should not have happened.

hobiecat

27th Aug 2009, 19:29

Klala,

You better get your torque meter checked. Your on the worng track bubby.:ok::=

Ned-Air2Air

27th Aug 2009, 20:18

Crab - Interesting in your post about the aluminium bending and composites dont. Us Kiwis are very familiar with the way composites let go when they do, and I refer to the mast of the Team New Zealand boat in the Americas Cup on Auckland harbour.

It let go with a hell of a snap like a bomb going off and that was the end of it, whereas in the old days the metal masts would start to signs of failure before they went.

Ned

widgeon

27th Aug 2009, 20:41

My first week at Westlands , many decades ago , I witnessed the destructive test of a prototype Scout composite tail boom. One minute it was complete and then after a loud crack there were lots of graphite splinters flying around.
Of course this was the bad old days when all we could get was unidirectional High Modulus Fibre prepreg . The newer woven fibre is a little more forgiving. I suspect the mast was mostly Uni directional .

heliski22

27th Aug 2009, 20:47

SASless

AW in Italy sent an e-mail to all operators regarding this incident yesterday afternoon, probably around 24-hours after the incident. They say they will provide updates.

22

SASless

27th Aug 2009, 21:04

Bad news travels fast.......wonder how many emails they got asking questions before they sent out the one they did? I would not want to be on the desk in Vergiate or Gallarate where ever it is these days!

On second thought.....if Ivanna L. was still around.....indeed I would!:E

tottigol

28th Aug 2009, 00:18

"Gulf Has 3 Hel With Debonding Tail Boom All Of Them In The Same Location And Have Been Repaired By Augosta,coz They Under Wranty,and For This One Is Twisted Not Fail Off Like You Said Gayes.the Tail Boom Has Beem Twisted Then Fail Off On The Right Side Of The Helicopter,i Think From My Opnion The Pilot Look The Nose Wheel Insted To Unlook ,whill Checking The Chk List Befor Taxing"

BADA-BADA-BOOM!:ugh:

KLALA

28th Aug 2009, 04:34

Mick
I think you are a wise man and I do respect your experience , I did not say both pilots are less of experience but I meant that may be by mistake he locked the nose wheel instead of un locked ,the rezone I said that ,the tail boom will required a lot of force FOR twisted ,this force will come if the helicopter is moving in the taxing and suddenly it stopped in the direction of rotation .but I have news for you from 1 HR I received an e-mail from one of my friends he send me a copy from AD ISSUED BY EASA IN 13 AUGUST 2008 THE AD NO 2008-0157 AND IT SAY THAT SOME OF THE AB/139 HELICOPTERS REPORTED FINDING DEBONDING OF FUSLGE TAIL BOOM PANELS THIS CONDITION IF NOT DETECTED ANDCORRECCTED ,COULD LEAD TO THE STRUCTURAL COLLAPSE OF THE TAIL BOOM ANSTALLATION ,RESULTING IN LOSS OF CONTROL OF THE HELICOPTER ,BY THE WAY AM NOT WORKING IN THIS CO,BUT I HAVE FRINEDS THER ,THANKS FOR THE INFORMATION YOU GIVE ME ABOUT THE CASE

spinwing

28th Aug 2009, 05:41

Mmm ....

Well as this drama continues ... it would seem that some of us flying these hotrods have had certain information witheld from us with regard the seriousness of this de-bonding issue ....

I have heard of 'booms going from NEW to delaminated in less than 11 hours time in service ..... This is NOT good .... Agusta had better sort this out sooner rather than later.

Not a happy camper! :mad:

Um... lifting...

28th Aug 2009, 05:51

Agusta Statement

Somma Lombardo, Italy
August 26th, 2009

AgustaWestland Statement

To: All Customers and Operators of AW139 Helicopter Model

All AgustaWestland Authorized Service Centres for the AW139
Helicopter Model

Subject: AW139 Tailboom Damage Occurrence

Dear Customer,

The purpose of this Letter is to inform You that on August 25th, 2009 an AW139
aircraft had an occurrence resulting in major damage to the aircraft tail boom while ground
taxiing.
There were no injuries to the crew or passengers.
AgustaWestland has immediately dispatched a technical team to inspect the aircraft and
assist the event investigation.
No similar occurrence has taken place with any other AW139.
AgustaWestland has been notified by the Customer that the rest of its fleet is continuing
operations.
AgustaWestland wish to remind You the importance of strictly applying the inspection
procedures for the tail boom area prescribed in the applicable AW139 Maintenance
publication.
AgustaWestland will keep all AW139 Customers and Operators informed of any follow‐on
information that may emerge.

Sincerely Yours,

Marco Sala
AgustaWestland
Vice President
Customer Support & Services ‐ Italy

Ned-Air2Air

28th Aug 2009, 06:15

I love how they call it an "occurrence", is it still called this if it had let go once the aircraft was airborne :mad:

blakmax

28th Aug 2009, 06:18

Gulf Has 3 Hel With Debonding Tail Boom All Of Them In The Same Location And Have Been Repaired By Augosta

How were the disbonds repaired? By injection of fresh adhesive?

[email protected]

28th Aug 2009, 06:22

The problem with composite materials is that you are flying something that is essentially held together with glue (very hightech glue though) and if the adhesion is anything other than perfect then debonding and delamination can occur.

I was told that another problem is that traditional Non Destructive Testing Techniques don't work on composites so finding voids in the structure or identifying cracks is much more difficult. Any gurus know what techniques are being used nowadays? The Westlands test on the BERP blades on the Lynx was a 'tap-test' and the engineers had to listen to the sound the blade made when tapped with a 2p piece for hollow areas.

I'm sure AW are far more swept up than that now - anyone have the inspection procedures for the tail boom to hand??

TheOldTimer

28th Aug 2009, 06:50

Crab, as usual you are absoluty correct. You got in first re your last mail however will still give my input in the hope it will add something useful. I think I may have attended a similar series of lectures on composite structures, construction, stress paths and conditions leading to the degrading of structural integrity. One issue was the detection of de-lamination and the integrity of the bonding agent. As you say put simply metal structures tend to relieve stress loads by flexing, composite remain ridged, hence absorb these loads within the structure. The issue re continuing structural strength with no obvious signs of de-bonding of the skins as in this case it seems, is tricky to detect. Tapping the skins with a coin just doesn’t do the job. I do understand that ultra sound systems can give a better indication of the condition of the honeycomb central core and its bonding to outer skins.
In short, without visible external indications it’s almost impossible to gauge internal structural integrity.
Repair of this type of structure is another matter and extremely complex and generally outside the capacity of the operator. I would be very surprised if the repair indicated in this thread was concerned with the structure, rather more lightly to have been non- structural parts, and with the guidance and approval off A.
That all being said lessons need to be learnt and I am sure they will be.
I remember the discussions on this type of construction years ago when they first arrived and the reluctance of us non believers to believe that they could replace the metal structure

TukTuk BoomBoom

28th Aug 2009, 09:05

So....Aw139s have tailbooms fall off, S76s have substandard windscreens, 332s have gbxs blow up and the S92 has a run dry problem...
Bring back the 412!

[email protected]

28th Aug 2009, 09:46

Oldtimer - thanks for the input, what you said here In short, without visible external indications it’s almost impossible to gauge internal structural integrity. is why it seems strange to try and repair such a complex structure as a tail boom rather than replace it. I know the answer will be cost but losing an aircraft (which is what would have happened in this case had it got airborne and then suffered the failure) is so much more expensive.

A very good friend of mine is something of a composites guru but his area of expertise is maritime based as he repairs all the lifeboats for the RNLI. If a lifeboat is crunched he often has to replace much larger sections of the vessel than were obviously damaged, just be sure that any unseen damage is removed as well.

TheOldTimer

28th Aug 2009, 10:03

Crab, agree, but I have no idea what the inspection criteria would be re the 139 tail boom, however whatever they may have been they did not detect a severity that required replacement it seems. As you know impact damage on these structures has to be severe to cause visible deformation of the structure, it is more lightly that any visible indications were chipping of the external paint surface. Possibly the use of an Ultra Sonic detection system called FLAUS that has I understand shown success in detecting hidden de-bonding and internal voids may have help, I don’t know if this technology is in use or approved for use.

chopjock

28th Aug 2009, 11:52

Are we all agreed this incident would be far less likely to happen in the air?

Hedge36

28th Aug 2009, 12:44

Well, we hope.

birrddog

28th Aug 2009, 19:40

chopjock, do you mean in forward flight or in the air full stop (e.g. incl. hover)?

sycamore

28th Aug 2009, 19:55

I would have thought this `occurrence` should be classified as an accident,as the aircraft was taxiing,rotors running,etc, period of operation of the aircraft..

Ned-Air2Air

28th Aug 2009, 20:42

Could those who fly the AW139 answer this question.

Obviously there was substantial stress on the tailboom when they were doing their specific taxy procedure and thats when the boom gave way.

Now lets say they had become airborne without the separation happening, what other stages of the flight would have had that same large amount of stress on the boom that it could have given way.

Not being an aerodynamicist I am unsure of what parts of a flight regime have the highest stress loadings.

Ned

chopjock

28th Aug 2009, 21:33

I would have thought that when airborne, even at MTOW the tail boom would be easier to pull around than when on the ground. Therefore had it been air taxi ing, this accident may not have happened.

A bit like pushing a boat when afloat is much easier than pushing a boat aground.(or on wheels).

SASless

28th Aug 2009, 21:39

Chopjock,

If this "incident" had occurred in the air it would be an "accdient" for sure.

It is an interesting question you pose about the forces....but I would imagine hovering cross wind in a very strong wind at max weight would also produce quite a bit of force. Whether making tight turns on the ground....or forgetting the nose wheel lock and trying to make a turn would exceed the airborne forces is for the design engineers to answer to.

No matter.....tail booms are not supposed to fail this way for any reason shy of a massive overloading such as a tail boom strike or crash. (IMHO anyway!)

Rigga

28th Aug 2009, 22:39

It's bin a few years but...

The only sure way of finding disbonding or voids within and GRP/CFRP structures is by using Ultrasound Techniques (a La - The Midwife's Pics of your loverly babies in mummy's tummy)

The 2p test (it used to be 10p in my day!) is okay for finding voids or disbonding near to the surface but is not reliable for deeper structures. TAP tests on Blades are looking for disbonding at the surface or the skin joint to the Spar - but not in the Spar itself.

X-Rays are also not particularly accurate and are open to "interpretation" issues.

Voids are prone to 'speading' under stress so all small knocks should be registered ASAP and assessed 'as soon as practicable' (I love that phrase!) and mapped for future periodic monitoring.

Also, to repair CFRP is a really specialist job - you cant use a normal drill, for instance, as it won't cut the Carbon Fibres without dragging them through the resin, and any repair scheme would need to be drawn up (though not for this particular Tail Boom!) by more specialists, and conducted by yet more specialists. (I remember a RAF Harrier GR7 once waited 18 months for an intake "Bird-Strike" Plate repair scheme that took two weeks to complete)

Although it is a very useful material - I personally don't like CFRP for structural uses!

212man

29th Aug 2009, 01:51

I distinctly recall Nick Lappos telling us that the stresses on the tailboom are far greater in flight than while ground taxying. So, it's probaby wrong to even discuss whether it would have failed in flight - it definitely would have done!

Outwest

29th Aug 2009, 02:32

So, it's probaby wrong to even discuss whether it would have failed in flight - it definitely would have done!

Could not agree more!!

I have never used full pedal while ground taxing, however have "hit the stops" many times in flight, with many different types of a/c.......cross winds, hot and high, etc.

This would have been a catastrophe if the a/c would have gotten airborne........

2leftskids

29th Aug 2009, 05:43

You would use more pedal hot and high because the tail rotor is less effective in these conditions. So yes you would use more pedal but not necessarily apply more force to the boom.

In a strong croswind a larger percentage of tail rotor thrust would be used to counter the effect of the wind to the vertical fin and boom itself so would not be transferred as an increased tortional force to the boom.

blakmax

29th Aug 2009, 06:13

Guys, the problem with NDI of bonding problems is not just limited to composites. Adhesive bonds depend on chemical reactions at the interface between materials and in some cases those chemical bonds are susceptible to degradation in service, usually by hydration at the interface. This occurs on composite and metal bond surfaces. NDI can ONLY detect defects which cause an air gap, and hence interupt transmission of sound waves. NDI can NOT detect interfaces which are degrading but have not yet separated. In other words, NDI can tell you if you have a disbond, but it can not tell you that you are about to have a disbond.
I did ask what the repair method was (and go no response) because if it involves injecting fresh adhesive, that will only fill the air gap so that NDI will pass the repair. It will NOT and can never restore bond strength because the surfaces are not chemically active and hence no chemical bonds can be formed.
My next question would be "what was the production method used to prepare the composite bond surface?" because if it relied solely on sacrificial peel plies (removed to produce a "clean" surface) then there is a potential for further problems. The surface must not only be clean, it must be chemically active, and removal of peel plies does not produce a chemically active surface. Worse yet, there are some peel plies which transfer release agent onto the bonding surface, hence causing contamination.
The real problem is that the FARs for certification of aircraft structures require testing for strength and fatigue, neither of which are the cause of bond degradation in service. It is actually possible to fully certify a structure which has a high potential for in service failure through bond degradation. And NDI will not find the disbond until it has occurred. The trick is to find the disbond before it reaches a critical size.
Now, can anyone tell me A: what is the repair method? and B: do they use peel plies during construction?
This is not an attack on this particular manufacturer. There are many well intentioned constructors out there who make the same mistakes.

rotorspeed

29th Aug 2009, 10:34

Surely the tail boom structural failure force on any helicopter is well in excess of any load that could possibly be applied through max pedal, whether on the ground on in the air?

spinwing

29th Aug 2009, 11:58

Mmmm ...

blakmax .....

We who have been flying these machines in the middle east have for some time been concerned by statements presumed to have come from Agusta saying .... "these de-bonding issues are only prevalent on machines operating in the middle east" .... would the adhesive breakdowns you mention be more problematic with the heat issues we suffer ie a/c parked in high ambient (+47C) temps, differential expansion/contraction of the structure ... also the area at issue is prone to exhaust heating action from rotor downwash.

All of which we now find very worrying .... :uhoh:

JohnDixson

29th Aug 2009, 12:25

FAR 29.351 describes the design yawing loads. I'd assume the JAR's are similar.

Thanks,
John Dixson

blakmax

29th Aug 2009, 13:47

Hi spinwing

You say the manufacturer says these de-bonding issues are only prevalent on machines operating in the middle eastRegretably these issues have been reported elsewhere. Reference NTSB A08-25_29 dated 09 Jun 2008. Interfacial degradation is exacerbated by hot and wet environments. Hence, a number of OEMs have incorrectly blamed tropical environments for similar failures. I am also aware of manufacturers blaming wash fluids, bird crap, whatever. The real reason is a failure of resistance to environmental degradation at the interface and this relates directly back to the failure of the FARs (or other regulations such as JARs, DEF STAN 00 970, MIL STD 1530 etc.) to require demonstration of the durability of the interfacial chemical bonds in the operating environment.

The current FARs require that bonds use "processes known to produce a sound structure". But what constitutes a sound structure? If it passes NDI is it a sound structure? Not necessarily because NDI can not interrogate bad interfaces. If it passes bond strength tests, is it a sound structure? Not necessarily because if the test is conducted before the interface is degraded, then there will be a false-positive result. If it passes fatigue testing is it a sound structure? Not necessarily because if the test is conducted before the interface is degraded, then there will be a false-positive result. None of these requirements address the time-dependent degradation of the chemical bonds at the interface. Hence, the FARs and other reg's which require strength and fatigue tests and the use of processes that produce a sound structure do not prevent a common (dare I say the most dominant) failure mechanism in adhesive bonded structures.

I stress again that it is NOT the manufacturers who are liable, it is the regulators who permit certification of structures which are susceptible to a known deficiency. Eventually some smart lawyer will stop sueing the manufacturers and realise that it really is the regulations themselves which permit structures to be certified such that they fail to prevent a known, common failure mechanism. I published a paper on this in 1996 and I am aware that the regulators have seen this paper ( see my web site at adhesionassociates.com ).

I have been trying since 1996 to have this deficiency rectified. If it takes a smart lawyer to defend a manufacturer or another even smarter lawyer to seek damages for defcient regulations, I would be available for discussions. My objective is not to make money from other's demise, but to prevent further needless loss of life. In the case of this specific aircraft, the on-ground failure avoided loss of life. It does not mean that the structure which in all probability met certification requirments was actually airworthy. I am well aware of other cases where "rotorheads" have not been so fortunate. The problem is not just limited to rotary wing aircraft. There are examples from fixed wing structures as well.

And can I stress again, if ANYONE has ANY testing evidence that injection repairs for disbonds actually restore ANY strength, I'd be very interested. The difference between removing the NDI signature and actually restoring strength is critical. Similarly, if an adhesive disbond exhibits interfacial failure (metal or composite) this is a processing and certification issue, not a loads issue.

On a related issue, I have data (admittedly old data) which shows that injection repair for delaminations in composites actually REDUCES fatigue life compared to doing absoultely nothing. I actually forced one OEM to admit that injection repairs for delaminations could not restore ultimate strength and at best slowed down delamination growth. Any comments?

Regards

Blakmax

212man

29th Aug 2009, 16:40

It sounds reminiscent of when the Mosquitoe was deployed to the far East 65 years ago!

Aser

29th Aug 2009, 17:28

OMG! :ugh:

Some more detailed pictures, sent by a friend:
http://i66.photobucket.com/albums/h263/aser_martinez/A7GHC/DSC04051.jpg
http://i66.photobucket.com/albums/h263/aser_martinez/A7GHC/DSC04052.jpg
http://i66.photobucket.com/albums/h263/aser_martinez/A7GHC/DSC04054.jpg
http://i66.photobucket.com/albums/h263/aser_martinez/A7GHC/DSC04055.jpg
http://i66.photobucket.com/albums/h263/aser_martinez/A7GHC/DSC04057.jpg

I'm glad everyone is safe

Regards
Aser

TheOldTimer

29th Aug 2009, 17:45

Thank you blakmax, for the interesting input and the background technical information. I had always felt that with the new emerging composite structures that a important part of the structure, the glue and the efficiency of its bonding would have been well covered, maybe I need to expand my reading list.
The matter of the structure being affected climatically and the approach taken after events that could damage or reduce its load carrying capability is a matter I feel needing some attention. I feel myself that these structures need a regular form of inspection using a system that will be able to assess any degrade of its integrity. Certainly after an event that may have cause a weakness that could remain hidden for a number of years a more involved inspection should be required for the life of the structure.
The fact that after repair defects could spread under repeated loading and not be visible or seemingly detectable in the early stages using present methods is interesting. I am sure that JAA offer some advisory on this whole subject ?
I have just seen the snaps from Aser, I apologies to Crab, indeed the statement WTF seems to be well justified.

Thanks for your indulgence.

noooby

29th Aug 2009, 18:05

Crab, while I agree with your comments on the use and damage characteristics of compositie structures, I disagree with people calling the 139 tailboom composite. It is actually constructed from Aluminium, just like the fuselage. It is called composite by many because of the fact that it uses adhesive for the structural joining of parts instead of rivets, in which case we should be calling most, if not all, metal rotor blades composite as well.
The only true composite part of the airframe is from the aft edge of the cockpit doors forward, which is kevlar. The rest of the helicopter structure is Aluminium honeycomb construction. The tailboom construction is quite similar to the forward section of the 412 tailboom, and you should see how they can debond if you don't have Alpine Ejectors fitted!!
I dare say, like the AS365, earlier S-76, and some other types, a heat shield of some sort is needed until a permanent solution is found. I do see that there is a newer tailboom out with different honeycomb material between the metal skins. Will that solve the issues that some customers are having? Only time will tell, but I'm sure they are working on it.

The Old Timer, who said the defects weren't visible? Some people are saying nothing was visible, others are saying that there was something visible before this happened. Hopefully we will all soon know something definite.

spinwing

29th Aug 2009, 18:55

Mmmm ....

Nooby .....

An opportunity I had to (personally) visually inspect 2 areas of de-bonding on the 139 airframe both of which had the surface skin removed to reveal the "honeycomb" below ... showed a material which was not aluminium honeycomb but which appeared to some sort of "Kevlar" type honeycomb product (or similar) if I am not mistaken (and it has been known for that to happen, on more than one occasion) I would respectfully suggest that this construction is then truly a "composite construction" by definition ie one composed of dissimilar materials! Those areas were ... The Top deck area between the engines normally occupied by the exhaust ducts , and the area of de-bonding on the RHS of the tail-boom exactly where this "incident" damage occurred.

I must acknowledge however to not having been privy to an Engineers course on this Airframe (just a "drivers' conversion!) much to my personal frustration! .... as usual I stand to be corrected as necessary!

Cheers

noooby

29th Aug 2009, 20:13

spinwing, you are correct, the honeycomb material used is not aluminium. Nomex I think on the new tailbooms? Not sure about the older tailbooms (I'd have to have a look at the BT).
Composite construction for the tailboom yes, but people seem to be implying that the boom skin is made of composite material, which is not correct.
Poor gluing seems to be one of the issues, specially with the top deck area. The other main issue seems to be dissipation of, or protection from, excessive heat. Improvements in the manufacturing processes are definitely needed!

alnic

29th Aug 2009, 21:08

http://www.helicoptersmagazine.com/images/stories/New%20products/r22_delambladeweb.jpg

An example of a delaminated rotor blade for those who haven't seen one.

9Aplus

29th Aug 2009, 21:50

This one above belongs to R44 and Frank already deal with this subject.....
BTW have some info that Robinson is going to perform real life test of
new protective covers with two units flying from UK to East up to middle
of nowhere island....

Ned-Air2Air

29th Aug 2009, 23:58

On the AH1Z I remember they had problems with the exhaust heat causing MAJOR issues to the tailboom and their solution was to create new exhausts directing the heat upwards and away from the boom.

Could the heat from the exhausts be a contributing factor to this problem and could the same solution be applied.

Just wondering.

Ned

blakmax

29th Aug 2009, 23:59

Noooby is correct. There is a lot of confusion about differentiating between composite construction (adhesive bonded structures) and composite materials (fibres in a plastic matrix). I have never understood why the FAA lumps these two completely separate technologies into one. See for example AC-20-107A, and the entire basis for training in composite repairs. Just because a technician can successfully fabricate a piece of fibre-glass does not mean that he can prepare a metallic surface for sticking that fibre-glass to it.

Adhesive bonding and composite fabrication are two completely different technologies. The design methodology is different, the failure modes are different, testing methods are different. The only commonality is that they start off as a sticky material which hardens to form the final product. The FAA lumps them together for the purposes of certification, yet if adhesive bonded structures were designed on actual strength (rather than an average shear stress basis which is known to be invalid) the certification methodology could be dramatically streamlined, saving heaps in certification costs.

What is totally missing in the FAA system at present is a requirement to certify that the bonding processes (for metals or composites) actually do produce durable bonds which will last for the life of the aircraft and will not result in failures like the one shown in the adjacent photograph.

The way to tell if the AW139 failure is due to bond durability is to inspect the failure surface using the Mark I eye-ball. If the adhesive has fractured leaving a complete coating on both surfaces, then is is a cohesion failure and the failure is due to a design issue or an overload. However, if the adhesive separates at the interface, then this is a bond durability issue which is 99.99% due to the process used to prepare the metal for bonding at the time of manufacture. And just to be clear, humidity and temperature actually only accelerate failure of bonds with poor durability. An absence of humidity and temperature simply delay the same outcome, which will be interfacial failure ata a later time than if humidity and temperature were involved.

noooby

30th Aug 2009, 01:31

blakmax, haven't seen a tailboom delam on a 139, so don't know about the failure type there. Have seen and repaired the upper deck delams, and the glue all stayed with the skin on the ones I saw. Basically none left on the honeycomb.
Smells like poor quality control at manufacture to me. I'm sure heat has something to do with it, but if the bond joint was 100% to begin with, I don't think we would be seeing these problems.

blakmax

30th Aug 2009, 02:54

Hi nooby and spinwing. What you are describing is termed adhesion fillet bond failure. This was a phenomenum which led to in-flight failure of a significant number of rudders on one type of military aircraft. It is quite significant in that the flatwise tensile strength of the sandwich structure (the load necessary to pull the skin vertically off the core) degrades to less than 10% of original strength. see DAVIS, M.J., Chester, R.J., Perl, D.R., Pomerleau, E., Vallerand, M., Honeycomb Bond and Core Durability Issues; Experiences within CREDP Nations, Aging Aircraft Conference, Williamsberg, VA, Aug 31-Sep 02 1998. I can mail you a copy if you PM me your address.

The real problem is that it is very difficult to detect until the problem is significant. If there is still contact between the core and the adhesive, it is possible to transmit sound waves (tap test or ultrasonic) and a false positive results. It is usually only when the skin experiences out of plane loads (aerodynamic, buckling or internal pressure due to heating) that the skin separates from the core and the problem can be found.

The problem typically occurs where water has entered a sandwich panel (often through those stupid injection repairs which never work anyway) and the interface hydrates. If the problem occurs early in the TIS, it is probably because the core was contaminated or not properly dried at the time of production. Nomex core absorbs a large amount of water just sitting on a bench. (I have measured 5% by weight.) If the production environment is not humidity controlled and if the core is not moisture evacuated prior to bonding, then the water will inhibit any chemical reactions at the interface and the bond will be weak.

Regards

Blakmax

noooby

30th Aug 2009, 05:14

Aha, Nomex is the water sponge. That explains sucking so much water out of the top deck during the repair!
Agusta BT139-159 is an optional bulletin that details the replacement of some (looks like most, if not all!) tailboom panels. The original panels have nomex honeycomb, the new ones have metallic honeycomb. Anyone with a 00234 tailboom, you have Nomex honeycomb. If you have a 00235 tailboom, the metallic honeycomb is inside. Supposed to have superior adhesion at high temps, and won't be as affected by moisture ingress.
Better go check your tailboom P/N's!

blakmax

30th Aug 2009, 06:45

Hold that thought nooby!

Another problem with honeycomb core is that you can not do an adequate surface preparation process on bare core. To get a durable bond to aluminium, you need a chemically active surface which is treated to develop a resistance to hydration. You can not abrade or chemically treat bare aluminium core material, so it is usually just solvent degreased (if anything is done at all!). Hence, the bond is still susceptible to hydration.

The real solution to hydration is for the manufacturer to use phosphoric anodised core which has an epoxy primer applied as part of the core manufacturing process. ALCORE and Hexcell offer these products. In that case the interface between the aluminium and the primer is appropriately managed, and when the part manufacturer bonds to the core, he performs an epoxy-to-epoxy bond on which is far easier to generate hydration resistance. The studies outlined in the reference in the previous message found that PAA and primed core was considerably less susceptible to adhesion fillet bond failure between the adhesive and core.

So, the question is "Did Augusta use PAA and primed core?" If not, they simply delay the onset of the same problem.

Regards

Blakmax

PS don't you ever sleep?

[email protected]

30th Aug 2009, 10:13

Westlands have past experience with debonding items - the TR driveshafts on the Lynx had a titanium plug glued into the steel driveshaft tube and because there were no rivets or bolts, there were no witness marks to show the bonding was letting go.

Sadly this caused the loss of two AAC pilots when the bonding in a TR driveshaft let go at 200' in the hover on an airtest.

It then turned out there were 3 differently modified TR shafts in the fleet and the type that failed was being phased out.

blakmax

30th Aug 2009, 10:32

Sadly Crabby, this is what happens when people believe that you only need a clean surface to get a good bond, and short term testing as stipulated by the FARs lets such bonds pass certification.
The organisation I worked for actually listened to reason and after we changed our processes based on a reliable validation test, we had only three failures in bonded repairs since 1992, out of a total exceeding 4000 repairs. In every case, we could determine that technician error (or laziness) was the source of the failures.
As long as we have certification rules based on horns-and-hooves adhesive bonding technology we will see bond failures which may unfortunately result in loss of aircraft and lives where the bond is flight critical.
Even worse, the level of comprehension of forensic data by crash investigators is often very low. I don't mean any disrespect to our wonderful crash investigators, but this really is a specialised area and it probably only constitutes a few percent of their experience, and their knowledge is often hand-me-down based on bad historic information. I have seen examples of this low level of knowledge where one investigator found fatigue markings in a bond failure and concluded that the bond failed by fatigue. Pity that the striations were at the middle of the bond and the rest of the disbond was at the interface. The two features could never be related. I could list several other examples where incorrect conclusions were drawn, but I don't wish to extend the message length too much.

Regards

blakmax

500e

30th Aug 2009, 15:07

Somma Lombardo, Italy
August 26th, 2009
AgustaWestland Statement
To: All Customers and Operators of AW139 Helicopter Model
All AgustaWestland Authorized Service Centres for the AW139
Helicopter Model
Subject: AW139 Tailboom Damage Occurrence
Dear Customer,
The purpose of this Letter is to inform You that on August 25th, 2009 an AW139
aircraft had an occurrence resulting in major damage to the aircraft tail boom while ground taxiing.
Strange name for Failure, (my bold highlight) what about the rest of the tails that are showing serious problems, "occurrences or failures".
Do these machines have HUMS?.
What is the inspection regime to detect failure in-service.
If as some say there is no outward sign of de lamination, how is a pilot to know\inspect for a serviceable craft.

chopjock

30th Aug 2009, 19:19

If as some say there is no outward sign of de lamination, how is a pilot to know\inspect for a serviceable craft.

I suppose one way to test the tail boom for integrity would be to ground taxi at MAUW with the nose weel locked and apply full right pedal.:}

mickjoebill

30th Aug 2009, 23:39

Slightly off topic,
Does the skin of a tailboom on light helis such as AS350 Bell 206 or r44 form a significant part of its structural integrity?

Mickjoebill

chuks

31st Aug 2009, 06:11

In the picture the tail boom is displaced to the right, which argues that it was giving the nose a push to the left, or is there something here I am overlooking? (I know very little about helicopters.) I would expect to see the boom off to the left for a right yaw input there.

Is the tailboom so weak that it will break before the nosewheel tires skid, assuming that the steering is locked with full yaw input?

Um... lifting...

31st Aug 2009, 07:24

chuks-
The tailrotor on the 139 is a tractor (not your red, Donetsk, Ukraine manufactured tractor... but a tractor nevertheless). Many Bells (206, 407) are pushers. The Huey folk can tell all manner of tales on tailrotor placement and direction of rotation and the like... all rather confusing.
On the 139, the main blades turn counter (anti, as our British friends would have it) clockwise when viewed from above, as do most American, German, and Italian single-rotor helicopters. This would mean that an increasing torque input (due to increased pitch pull on the main blades, which one tends to need to do to set the 139 in motion on the wheels) would require an increase in left pedal (also known in some circles as antitorque), pulling the tailboom to the right.
The clever observer, viewing a sound 139, will also notice that the tail rotor is canted to the left to provide a bit of lift to the tail to counteract the rather aft CofG in the hover. The long-nose versions are a bit more forward CofG by all reports, but I haven't flown one of those just yet.
The 139 has a nosewheel that automatically centers and locks upon liftoff and has to be selected to unlock upon landing. Because of the aft CofG, the nosewheel is the first thing to lift off the ground in the hover, but by the same token, hasn't got an enormous load pinning it to the ground, either. I'm not particularly fond of this locking nosewheel feature. However, I don't pretend to know entirely what happened in this incident, though I watch with interest.
The boom is probably less likely to fail in the other direction because that is when loads from the tailrotor are lower.
French and Russian single-rotor machines tend to have directions of rotations that are reversed.

212man

31st Aug 2009, 08:02

Is anyone seriously under the impression that the nosewheel lock had anything to do with this?

heliski22

31st Aug 2009, 09:00

Yes, I did wonder about that 212man

heliski22

31st Aug 2009, 09:11

Noooby - did you mean P/N ..00234 or ..00134

Just been having a closer look at BT139-159. It does seem like an awful lot of bits and man hours for an Optional BT?

What is a little worrying, in light of this Middle East incident, however, is the question of just why would we want to "improve" (that's the word in the BT) our tail boom if there's nothing wrong with the one we've got, eh?

Answers on a postcard, etc., etc.....................

TheOldTimer

31st Aug 2009, 13:15

After a little time and I expect all operators of the 139 looking at the tail boom with large magnifying glasses and tapping like mad looking for any sign of a problem, I was wondering what the results are and the location of any failings??? A friend tells me the latest high tech tool is a little hammer rather than the coin?
:ok:

blakmax

31st Aug 2009, 13:19

OK, I'm not an Augusta expert (I'm not even a helicopter expert.) but what is that great phalic symbol attached to the separated residual boom and pointing at the sky in the recent pictures. It looks like some kind of strake which may have been bonded to the skin of the boom and failed as part of the event. I make that judgement on the basis that I can not see any fastener holes and also that anything that length would have bent if even a few fasteners were effective. A loss of such a length of stiffening structure would have made a very significant difference to the ability of the structure to survive any loads (wheel locked or not, and even if the pilot was wearing spotted underwear on a Monday afternoon two days after the full moon). :}

If it really was bonded and if the owners of this or other aircraft with similar failures can examine the surface to see if it has failed interfacially (areas with no adhesive on one surface) then there may be grounds for warranty action.

If this structure was not bonded, tell me and I will get back in my box. :sad:

It really is time that operators stopped accepting glib "you did xyz wrong" responses from manufacturers when expert assessment can demonstrate that the operator was not responsible for a failure. IF THE ADHESIVE HAS FAILED TO STICK, IT IS A MANUFACTURER'S PROCESSING PROBLEM. IF THE MANUFACTURER COMPLIED WITH THE REGULATIONS THEN IT IS A REGULATORY PROBLEM.

As I have already stated in this stream and others, curent FARs do not require assessment of bond durability other than static and fatigue testing, neither of which will prevent interfacial failure which is time dependent. Just do a search for "blakmax" postings.

This applies for other manufacturers, not just Augusta, as well as for non-rotary structures. If you overload the structure because the pilot forgot the brake or sneezed at the wrong time then the failure will separate the adhesive with residual adhesive on both surfaces at any location. That probably is a design or operational issue which could be debated. But interfacial failures are directly attributable to the manufacturer's process selection or the regulator's failure to exclude an important and common adhesive bond failure mode. There are very few things an operator can do which will result in interfacial failure of an adhesive bond unless the interface was already degraded due to the processes used by the manufacturer which probably met the regulator's requirements.

Perhaps action against the manufacturer to recover lost income due to reduced aircraft availability as well as cost of part replacement under warranty will actually address bond failure issues. Hopefully it will not be legal action on behalf of dependents after a loss of life incident. And before anyone asks, no, I am not a lawyer. I am an expert in composite and adhesive bonding technologies who wants to see best practice adopted by regulators and manufacturers so that we do not see failure of composite and bonded structures.

It is actually possible to design adhesive bonded joints such that they NEVER fail under any loads. That must be combined with validation of bonding processes that assure interfacial failure will never occur. So, if it is possible to eliminate bond failure by design methods, combined with validated processes which prevent interfacial failure then why do we continue to see bond failures? Because the current regulatory requirements for bond validation and common design practices do not preclude either failure mechanism.

After all, if temperature has an effect on durability of your bonded structures, then global warming is not your friend.

Regards

Blakmax

heliski22

31st Aug 2009, 14:00

blakmax

The "phallic symbol" you're looking at is a strake that is an aerodynamic add-on. It is fixed with fasteners (rivets) rather than bonded in any way. I can't remember if there's a sealant of any kind added to the join or if it's just a filler that is painted over. If you look at the second picture down in the group of photos on page 5, you'll see the holes where it was torn from the boom itself during the failure.

sycamore

31st Aug 2009, 14:11

212 man,I don`t think that it is necessarily suggested that the failure on this aircraft was as a nosewheel locked case; however,it would be interesting to know from `those who do` if the nosewheel is linked to the yaw pedals on the ground for ground manoeuvring or not at all. The nosewheel assembly looks to be directly attached to the oleo with very little `trail`,ie no trailing link, therefore all castoring has to come from the tail,or/in addition to any `inside` wheel braking. So all the yaw loads on the ground ,which will also include a rotary twisting torque,from the tail-rotor above the tail-boom will be felt in the `weakest` part of the boom.
Blakmax, the strake along the fuselage is for aerodynamic purposes to reduce yaw pedal transient reversals during sideways/crosswind hovering, by causing a loss/breakaway of the airflow around `deep`/heavily curved tailbooms due to the `Coanda` effect( teaspoon in running tapwater ),reducing sideloads in the low-speed areas.Developed in late `70`s by WHL as the Seakings got heavier,and pedal reversals at high torques moving on and off ships decks were becoming limiting. Early `breadboard` models were a piece of angle-iron,bolted on,cut to length as tests progressed,later production mods were pieces of angle-alloy bolted on....Works though !!

ShyTorque

31st Aug 2009, 14:41

however,it would be interesting to know from `those who do` if the nosewheel is linked to the yaw pedals on the ground for ground manoeuvring or not at all.

:confused:

In thirty years of flying, mainly rotary wing, I don't know of any helicopter where the nosewheel is linked to the yaw pedals, unlike some fixed wing aircraft. The locking pin is designed to lock the nosewheel fore and aft i.e. straight ahead.

I can't think of a situation where having yaw pedals linked to the nosewheel would be advantageous (or to the tailwheel, where fitted). I can think of some where it would possibly make the aircraft uncontrollable, for example a slow rolling takeoff in a crosswind.

Larger aircraft, such as the CH-53 and the Chinook do have steerable wheels. For example, the CH-53 nosewheel, and one of the two Chinook "tailwheels" are steerable (the other side castoring). However, this isn't done via the yaw pedals but my other mechanical means, the pilot's control being a left/right control knob on the console.

chopjock

31st Aug 2009, 15:01

Ok I've just seen the pics on p5. Goodness me It appears to me than heat from the exhaust was blown down with rotor downdraft on to the boom, combined with little or no airspeed (no cooling air) may well have softened the epoxy laminate enough to weaken the structure. Has anyone stood behind an exhaust on a Hughes 500 when the engine is idling? It's jolly well hot!
Are there any pics of this ship before the accident? if so look for heat / soot exhaust stains as a tell tale sign!

Epoxy (depending on type) tends to weaken above a temp of 150 ~ 180 deg C. I'm sure a turbine exhaust at 2m distance will be hotter than that. :eek:

I assume It's alright to guess what went wrong because no one was hurt.

Um... lifting...

31st Aug 2009, 15:06

The nose wheel steering system gives the pilot control of the aircraft
during ground movement. After take-off, centering mechanisms will
automatically align the wheels in a fore and aft direction for retraction
in the NLG bay.
For high speed rolling, the nose wheels need to be held directionally
aligned. The nose wheel center-lock is controlled by a switch on the
Landing Gear Control Panel (LGCP). Pressing the switch once
engages the center-lock, locking the nose wheels in a directionally
aligned fore and aft position. Pressing the switch a second time will
disengage the center-lock, leaving the nose wheels free for ground
maneuvering.

While bearing the weight of the aircraft, the nosegear (so I have been told by the engineers) can caster 360 degrees. The nosegear has a small amount of trail. It is not directly steerable, but is not difficult to get to turn from a dead stop with moderate tail rotor pedal input, with or without differential braking, unless it's locked.
There is a centering cam mechanism with automatic locking pin that engages when weight is removed from the weight on wheels switch. What the above quote from the RFM doesn't exactly spell out for you is that in practice you can only disengage the lock on the ground.
The locking pin is very strong. If not unlocked before towing (the external warning flag is not readily apparent), the locking pin is not the first point of failure in the gear assembly. The nose gear is designed to keep the entire aircraft aligned with the runway during a 60-knot running landing (it also tends to shimmy at moderately high taxi speeds).
At 6400kg, 1400kg of the aircraft's weight is on the twin nose gear, and 5000kg is divided between the two single main gear. Nose gear tire pressure is 137 psi. Main tire pressure is almost double that, as is the contact area of the main tires. The aircraft is supposed to be taxied on paved surfaces with the nosewheel unlocked no faster than 20 knots and no faster than 10 knots on grass surfaces. With the nosewheel locked, those speeds double. For emergencies, you can add 20 knots to each of those for maximum touchdown speeds of 60 and 40.

noooby

31st Aug 2009, 15:55

Blakmax and heliski, the tailboom strake is secured with screws to the tailboom, not rivets. The screws locate into inserts bonded into the honeycomb. Material is not very thick at the edges, so would probably tear sideways and leave the screws behind in this type of incident.
heliski, sorry if I was getting my tailboom numbers mixed up. the 00134 tailboom assy, has the 00234 tailboom as part of that assy. The 00235 tailboom is part of a different P/N tailboom assy. It can get confusing!
Um...lifting is correct. The nose gear can rotate through 360 degrees on the ground. Tunr any sharper than the towbar limit lines on the aircraft, and you may pivot the aircraft on one tyre, not good for the tyre and sure to leave rubber behind!

heliski22

31st Aug 2009, 16:08

Ok, Noooby, so we have the 234 - this does not give me a nice warm feeling just now, despite nearly 400 hours on the bird and the recent completion of BT139-134. :(

By the way, does anybody know if this aircraft was still operating at 6400kg MTOW or had it had the 6800kg mods? Just a thought?

9Aplus

31st Aug 2009, 18:22

Epoxy (depending on type) tends to weaken above a temp of 150 ~ 180 deg C
Epoxy weakening begins at 82-90 deg C on 150 all is down to plastic features....

spinwing

31st Aug 2009, 19:27

Mmmmm ....

... By the way, does anybody know if this aircraft was still operating at 6400kg MTOW or had it had the 6800kg mods?

Just a guess ... but from the pics above the aircraft still seems to be sitting quite low to the tarmac ... I think therefore that it was still in 6400Kg config.

I have noticed that after the 6800Kg mods they tend to sit higher up on the main gear ... prob due to the extra prx in the u/c struts and especially in this case with the boom nicely folded up against the airframe ....

amicus

31st Aug 2009, 19:45

To Rigga, blakmax and all,
All thanks for your input and I am a composites engineer with a few decades in composite experience. I have found most QA inspection techniques in all composites and adhesives to be sadly lacking, e. g. you can check that there is a bond, but that tells you nothing re strength of the bond.
My experience led me to invoking proof loading to limit load, (if you are initially worried re incurring damage, then proof load five times to limit load on early proof tests to prove no damage), as most reliable way as it solves the issues re both poor Q.A. or poor in-process control and is a lot cheaper than a bunch of questionable Q.A. techniques. If temperature levels are issue do proof test at max certified operating temperature via heat blankets.
I would follow same mandatory proof load procedure after any tail boom repair also, of course. Either it works or it doesn't with no doubts or questions, that is a benefit of proof loading and it catches lots of in-process faults that Q.A, doesn't catch. And given the tail boom and its problems, I would invoke 100% proof loading requirement and have a cheap permanent rig built for fast testing, no instrumentation needed, just a load vs. deflection accept./reject curve to measure.
Anyhow that it what i would insist on if I were at Augusta as clearly their Q.A. isn't catching defects and their in-process controls are questionable too.

JOURNIOO4

31st Aug 2009, 21:47

I think that the designer think that to avoid torsion and bending coupling in the fin root the tail gearbox has been located on the top of the fin and the transmition shaft is inside rather than outside the structure, in this way the shape continuity between the fin aerodynamic profile and the tail cone cannot prevent stress concentrations in the fin root section which result fatigue problems this is what I think happen in AW 139

blakmax

1st Sep 2009, 02:35

Thanks everyone for explaining what that phalic symbol was and how it was attached.

For Amicus, proof testing will certainly help improve detection of production defects and is actually mandated in the FARs for bonded and composite structures where NDI is unreliable. However, even proof testing will not prevent in-service degradation of bond interfaces. Similarly, any QA strength tests on coupons will not prevent degradation of bond interfaces. Strength and proof tests are only a snap-shot of bond strength at the time of testing. Interfacial degradation is time dependent, with moisture and temperature accelerating that degradation. If you test before the interface has had time to hydrate, then you get a false positive result.

The secret to success is to use tests that actually interrogate the resistance of the interface to hydration BEFORE construction. You simply can not test for this after construction. Demonstrate by wedge testing that the proces you are using can actually develop a hydration resistant interfacial bond, and then duplicate that process in production and you will have a structure which will not fail interfacially in later service. The test is ASTM D3762, but ignore the stated acceptance criteria in that ASTM. The correct critieria are in DOT/FAA/AR – TN06/07, Apr 2007 which can be obtained by email from the FAA Tech Center Library. US nationals can download the document, but because it is a .gov website, we foreigners can not access it for homeland security reasons. Too hard? send me a PM and I'll send a copy to you.

Regards

Blakmax

Skin King

1st Sep 2009, 04:25

Having been directly invovled now with at least 15 AW 139 Tailboom disbonds I would like now to share my observations. Please remember I'm not a composite engineer but simply the poor guy with the tap hammer who has to make the call as to whether or not the Tailboom Assy is serviceable.
Firstly, all but one of the delaminations I have seen started on the curve section of the R/H side of the Tailboom structure aft of where the top longeron finishes. The longerons are a hat section sheetmetal part that are rivetted into a recess on the internal side of the bonded panel and are approximately 28 inches in length. There are 3 longerons per side. Cut outs of the delaminated areas show that where the longeron finishes there is a 1 inch internal doubler strip between the outside skin and the Nomex core that I assume is suppose to transfer loads from all 3 longerons to the bonded panel they are rivetted to. As I said before I am not a composite engineer but this 1 inch internal doubler strip appears a little insufficent to carry the loads between the longerons and the bonded panel.
I can liken the 139 Tailboom build as to having a VW swinging off a pool que which is attached to a tree stump by masking tape.
Cut out samples have shown that in some cases that there is a lack of adhesive to the outside skin (aluminiun skin is still visible in parts) while in other cases failure of the core to adhere to the outside skin i.e there is adhesive on the outside skin and not the Nomex core.
No evidence of obvious water contamination to the core was visible in any samples taken.
Teletemps where placed in various locations on the Tailboom Assy and the average skin temps while on the ground where around 80 degrees C. Yes it's hot here. Temps whilst in flight where unconclusive so I would not like to speculate.
Ok having said all, it is my personal belief that there is a combination of adverse factors at work here the worst being the sudden transition of loads through the longerons to the Tailboom bonded structure. Couple this with poor bonding processes and constant heat soaks then disbonds are inevitable.
My advise to all engineers on 139's is to visually inspect the area Aft of the R/H upper longeron for skin ripples or waviness on your daily inspections and after each flight. But I guess the majority of us know that already.
I do hope Agusta comes up with a fix soon as I'm really getting bored with changing Booms. :ok:

blakmax

1st Sep 2009, 07:00

Thanks for the details skin king. However, let me assure you that there is no way structural loads of any sort can cause interfacial failures i.e. where there is an absence of adhesive on one adherend. Of importance, was there an impression of the core in the adhesive sample taken from over the disbond where the adhesive was absent from the core? If there is no impression, then that was not in contact at manufacture. If there is an impression, then it is probably inadequate drying of the core before bonding.

The disbond from the aluminium is more of a problem. If the surface of the adhesive was glossy, then there is probably another processing issue such as inadequate pressure application during bond or poor fit-up of the parts. If the surface of the adhesive replicates any features of the aluminium it was probably in contact but has disbonded due to hydration of the interface. The regulators need to ask Augusta what the process was, and not accept the usual "proprietary" crap. Lives are at stake and the regulator should step up and take control.

I certainly hope that the booms you are fitting are being replaced uner warranty because there are very few things an operator can do which will cause interfacial failure. To address the temperature theory, heat applied to adhesives reduces the strength of the bond but also increases plastic behaviour. Such failures should exhibit elongation marking on the bond, not interfacial failures.

Skin King

1st Sep 2009, 07:55

Thanks Blakmax for the reassurance about loads and disbonds. As I mention before the samples taken showed two types of failures.

1) There is adhesive and core impression on the outside skin. No adhesive on core cells.

2) There is adhesive on the core cells and only random areas or no adhesive on the outside skin. There is just a slight core impression in the areas of random adhesive.

It appeared to me that there maybe should have been an additional layer of unsupported film adhesive applied during layup particularly in the curved areas.

Anyway I'll leave it to the composite experts at Agusta to sort out their process woes.

And YES the replacement booms are under warranty even though some only last about 2 days before they disbonded as well.

The 139 is a good A/C and as with all new A/C has some growing pains. Luckily for us history dictates these pains get sorted out fairly quickly and I'm sure we will see a new improved Tailboom Assy sooner than later.:ok:

blakmax

1st Sep 2009, 09:08

Thanks skin king, your description basically confirms what I said. Maybe Agusta could do with a consultant expert to help sort out the problem. I'm not cheap, but neither is replacing tail booms on a regular basis. I'd be glad to help if they PM me.

Regards

Blakmax

Skin King

1st Sep 2009, 10:02

As mentioned in a previous post Agusta Optional BT 139-159 which replaces the Nomex core lateral panels of the Tailboom with aluminium core is available but before everyone gets too excited it does require specialised tooling and an approved jig of which there is only ONE..... Yep!! Italy.

It's good to see that Agusta are well aware of the problems :Dand lets hope that along with a change in core material came a change in the manufacturing process of these panels. Time will tell but as we are approaching our cooler months we may have to wait until next year to find out.

shak'n

1st Sep 2009, 11:09

In my limited experience with AW139, we too had debonding i.e. bubbling of the tail boom skin only on the left hand side directly under the #1 engine (LHS) exhaust stack. This also coincided being directly under the tail boom coanda strake.

The organization I was with had this bubbling/debonding occur on at least 2 out of the 3 aircraft it owned and all with less than 300hrs. One with only around 150hrs when the debonding was observed. The role of the aircraft was SAR/EMS requiring a lot of normal cruise flight regimes but also periodical slow speed searches, hoisting etc requiring extended hovering.

We observed one bubble appeared quite large immediately after shutdown but seemed to reduce in size after cooling down. At the base I was at, teletemp decals were fitted only on the left side (unfortunately) and the tail boom monitored. During normal flights, T/Os, cruise and Ldgs, no portion of telitemps discoloured. We then had to do a slow speed search (<20KIAS) for approx 1 hour duration which included a little downwind maneuvering . On shutdown from this sortie the first couple of segments had discoloured. We then did periods of hoisting and hotter segments became discoloured. I do not remember what the hottest temp we finally recorded but it was within the hottest one or two. Unfortunatley we did not put a teletemp on the RHS of the boom in the a similar position in order to get a comparison of temps on either side of the boom to see if both sides were seeing similar temps.

It is clear though that the bubbling/debonding is linked to the high temps being recorded there on the LHS. When looked at in conjunction with the Coanda strake, the airflow or vortex which it sets up beneath it particularly in the hover, must trap the exhaust gas from #1 engine immediately below it. It in effect becomes a Weber-like BBQ recirculating and increasing the localized heat against the skin. See the link below which shows the visualised airflow around these strakes in the hover admirably:

http://www.nwlink.com/~blrweb/pdfs/H_65%20flow%20analysis.pdf

(Hope this link works out)

I believe the laminar flow on the RHS (while in the hover and throughout the flight regime) prevents localised heat build up and the ensuing softening of epoxy and bubble formation. As I said, regrettably we did not fit teletemps on the RHS of the boom to compare the heat build up of each side against the flight envelopes. I am now no longer in a position to access a AW139 but recommend to anyone who does, try it and see what the teletemps on each side of the boom do with different flight regimes - in particular hovering, slow speed maneuvering, or extended downwind operation (on-ground, taxiing, hovering, slow speed).

A very easy quick fix: If the above is true, the simplest quick fix would be to amputate the first 2 feet of Coanda Strake to allow laminar flow of the rotor downwash to remove localised heat build up against the skin on the LHS of the tail boom i.e remove the hot exhaust gas of #1 engine with the laminar flow in the hover or at slow speed maneuvering . The remaining length of strake should still provide adequate Tail rotor thrust offset in the hover as originally designed to do but prevent localised "weber" effect of trapped, oven-like temps beneath the strake immediately under the exhaust stack of #1 engine.......

I'm not suggesting this as a permanent fix as I think a better tail boom design/product should be designed by those so-called design engineers who get paid all the big bucks and have limited responsibility for their end product need to do their job properly as we as aircrew are expected to do every day.

But, food for thought as a short term solution don't you think???

blakmax

1st Sep 2009, 14:13

The suggestion that this problem is driven by heat has some initial plausibility, but realistically the only way that adhesion fillet bond failure can be driven by heat is by the water content of the cells causing steam pressurisation during the operational heat cycle, combined with weak bonds between the adhesive and the core, and as I already suggested this is probably due to inadequate drying of the nomex prior to bonding and a lack of environment control in the bonding facility at the time of bonding.

Service heat can never cause interfacial failure between the skin and adhesive.

This is not operator caused. The PM reports I have is that these defects occur in incredibly short TIS after fitment (a few hours???) Now I don't care if the operator had a dozen naked maidens dancing on the boom, that would not cause interfacial defects. They are processing problems and that is a manufacturer issue which should continue to drive their warranty claims through the roof.

In defence of Agusta, I have seen exactly the same deficiencies in bonded structures from other rotary manufacturers as well as fixed wing manufacturers. THIS IS NOT ISOLATED TO ONE MANUFACTURER. It reflects a deficiency in the regulations in respect to adhesive bonded structures, combined with deficient engineering training in adhesive bonding. This is not a reflection on any manufacturer including Agusta, because to loosely quote Dick Chaney, you may know what you do know but you can not know what you don't know.

The route to success is to ask those who actually do know more than you to help you to sort out your adhesive bonding problems. The alternative is to take decisions with substant risks to company profits as well as human life. There are experts who are out there who can help and it is no discredit to a company to recognise that their knowledge base is limited to the requirements (and deficiencies) of current regulations. The choice is do you produce a product which barely complies with the requirements of the regulations or do you produce a product which meets the requirements of the regulations and has a justifiable reputation for excellence and never requires warranty replacement? While I clearly state my capabilities in this area, I acknowledge that there may be other experts out there. I simply urge manufacturers such as Agusta, EC, RHC and others to seek help to establish standards which actually produce quality products with sound structural capabilities and not just to pretend they have divinely acquired knowledge which is often demonstrably not backed up by service performance.

Remember this: No matter what design methodology is used, the shear strength of an ineffectively manufactured bond will eventually reach zero, despite the best intentions of the manufacturer. What varies is the time for the bond to achieve zero strength, or the intervention of a load which exceeds the degraded bond strength and results in premature failure. Unfortunately refusal to acknowledge gaps in technological capabilites have a high potential to cost lives, and those costs exceed warranty claims any day and subsequent legal costs may exceed the cost of specialist support by a considerable margin.

Regards

blakmax

JimL

1st Sep 2009, 15:58

This thread is a tribute to Blakmax and the other practitioners who have volunteered knowledge and advice.

Well done - keep the information flowing.

Jim

212man

1st Sep 2009, 16:17

Hear hear!

Blakmax, do you have any friends Who are metallurgists, gas turbine experts or AFCS specialists? If so, I think we have pprune sewn up!

Thanks alot for your contribution - it's been a joy to read.

SASless

1st Sep 2009, 17:19

I have a rule.....I must learn something new every day!

This week I have more than caught up with my sleep having read Blakmax's posts.....after reading each one of the posts I accomplished my daily goal and thus could fall back into the hammock with a clear conscience! ;)

Definitely enjoy his posts!:D:D:D

griffothefog

1st Sep 2009, 17:59

SAS,

What if he's a spook from Eurocopter? Me paranoid??? :E

heliski22

1st Sep 2009, 18:17

And hear, hear again!

I've been checking this thread several times a day just to see what new gems of wisdom have been brought forward. Supplemented by the observations of others with some hands on experience, blakmax's posts have been fascinating!

In the Unversity of Life, every day's a schoolday!!

Griffo, just cos you're paranoid doesn't mean they're not out to get you!!!

blakmax, if I read your last post correctly, is it probable that the longer the aircraft is in service without defects showing up, the less likely it may be that there will be problems at all?

22

Hilife

1st Sep 2009, 18:26

I’m not a lover of anything AW, but there are a lot of 139’s out there flying in all manner of operations and this is the first I’ve heard of a tailboom failure and supposedly on a low time machine.

Appreciate manufacturing flaws are possible, but post #24 suggests that this machine had history of a tail strike and after an inspection the machine was released to service.

I’m not privy to the circumstances and severity of the tail strike – if indeed this ever happened, but if true, it could be the root cause of this latest incident and not down to poor build, manufacturing processes, excessive localised heat or bonding design flaws.

Just a thought.

Shell Management

1st Sep 2009, 19:19

Hilife

A very good point - and there was indeed a previous incident. This fact has not been lost on the oil companies. That is why there have been no groundings of our AW139 fleets.

We hope the full picture will emerge over the next few days.

blakmax

1st Sep 2009, 21:50

Firstly, thanks guys for the encouaging comments, and no, I am not a spook for any company. I am a private consultant having recently retired after 37 years involvement in aircraft science and maintenance on military aircraft. I am happy to share my specialist knowledge on adhesive bond forensics in the hope of improving safety.

Now to deal with the tail strike theory. If the adhesive bonds were damaged by impact or overload of any description, then the failure would be by cohesion, with the adhesive fractured. There would be residual adhesive on both surfaces. The descriptions from people who have taken sections from disbonds clearly indicate interfacial failure, and a well bonded structure should never fail interfacially even under impact conditions.

And Yes, 212, I do have a friend who is a metallurgist and he is a damned good failure analyst.

Still waiting to hear from Agusta....

Thridle Op Des

2nd Sep 2009, 04:52

I'm wondering if the tailbooms are subcontracted to Alenia Aeronautica in Naples. On the FW side they are rapidly progressing towards pressurised structures in bonded composite, albeit with slightly different structural approaches between Boeing and AB so the science of adhesives is becoming much more interesting.

airwave45

2nd Sep 2009, 08:00

Given the known softening of epoxy at temperature,
The static temps on the booms during the day being 70-80 Deg C (ambient temps are mid 40's add in direct sunlight in the sandpit, 70 Deg C is easy)

The guys fly the 139's in an unusual profile offshore in the Qatari gulf (where A7 GHC failed), the cruise and approach are same as everywhere else, the lift to get back in the cruise is unusual.

The guys will lift off the deck, wait (I assume for egt, but I'm just SLF) then 9 times out of 10 will lift backwards up about 100' (Top of the legs of the jack ups are about that high, so it's easy for SLF to guage height)

the heli's come up at about 75 to 85 deg off the deck (90 being straight up) I'm used to helicopters exiting the decks in forward flight and notice the exit procedure as it's pretty close to the bow leg and hacks me off as I know neither driver is looking backwards to guage clearance from the leg.

This exit will surely have the boom / fus passing through the Exhaust gas, heating an already hot structure.

I assume this is done as the 139 is such a tail heavy beastie, but don't know.

If the helicopter is flown backwards and up through it's own exhaust gas 5 to 10 times a day, this might not help the already hot and stressed boom bonding product.

a little bird told me that 2 other booms have been found out of spec on the ramp at GH, that would make 3 on that ramp with issues.

Shell / Conoco _have_ grounded 139's here. (we are back to 412's) others are still flying 139's

spinwing

2nd Sep 2009, 09:00

Mmmmm ...

Airwave ..... you are making comments about pilot technique which are really not necessary to this thread .... if you are concerned about how they arrive or depart your platforms you should take one of the pilots aside and ask him to explain to you why they are using their to your eyes unusual procedures.... they most likely will be quite happy to explain. A lot is to do with decks placed out of wind or with NO WIND with high ambient OATs and High humidity often all at the same time ..... makes for different techniques than used in the North Sea.

With regard heat being blown down on the booms during takeoff .... I don't think that is a problem ... helicopter by design have a bloody big fan that tends to blow hot air across the booms all the time ... whether on the ground on takeoff or in flight .... were stuck with that!

Cheers ;)

Um... lifting...

2nd Sep 2009, 09:08

The pause is a momentary check of the power index (measure of first power limit in the 139) to determine the limit to which to pull during the takeoff.
The backward takeoff permits a safe landing back to the deck if at any point during the procedure an engine fails. Standard in the 139 for certain PC1 operations, though I can't see why it would be relevant to this thread.

chopjock

2nd Sep 2009, 10:04

I think Airwave has very valid points and are relevant to this thread. Surely anything the pilot can do to minimize the risk of exhaust heat to the tail boom is good? (Even though he shouldn't have to).

For example, in the cruise at 155kts there is virtually no downwash on the tailboom and max cooling air flow.
The more time spent in the hover and climbing backwards will obviously add risk to the tailboom getting hotter for longer. :)

Imagine this, a pilot carries out a twin engine type lift from a platform, backing away and climbing until safe to fly away on one engine. That's all very well and good so long as you still have a tail boom left. :eek:

Heli-phile

2nd Sep 2009, 10:27

I don't think I have ever seen a more disturbing failure on an aircraft structure.
My thoughts go to the "lucky" crew on this imminent flight, ...and the previous flights!!

:ooh:

ShyTorque

2nd Sep 2009, 10:44

The quoted "unusual", rearwards climbing, departure is likely to be the normal Class A takeoff profile for a raised helideck, as published in the Agusta Flight Manual. The crew are required to fly the aircraft in the published manner to satisfy the performance criteria for public transport operations.

Shell Management

2nd Sep 2009, 11:11

Yes, ShyTorque, very likely the FM has been followed. What is unusual is that other people who don't know what this is are judging it!

airwave45
Gulf Helicopters withdrew its other AW 139 from service pending further understanding of the cause of this incident. Shell are not currently placing any restrictions on the use of small number of AW 139 in Shell service.

blakmax

2nd Sep 2009, 11:29

Ok, irrespective of the trajectory of the aircraft at take off and/or the presence of dancing girls on the boom, there is really a high thermal input to bonded sandwich panels in service. I am aware of some measurements on composite structures sitting on the tarmac in northern Australia where the temperature exceeded 80C (180F for those in the US) just through solar heating. Exhaust gasses may or may not add to this depending on the previously mentioned big cooling fan just above.

Irrespective of the temperature, what is the mechanism which causes the failure? In reality it is the presence of the afore mentioned moisture absorbed by nomex (about 5%) and epoxy (about 2%). When heated, this leads to pressure in the cells both due to the expansion of air but also the release of water vapour which adds to the pressure. It is this pressure which results in the failure of the bond to the core. Even worse, if this panel is heated to above 100C (212F) in service or for hot-bond repair, then the pressure greatly increases.

Now there is another aspect which must be considered. If the nomex core has not been adequately dried before bonding, then the moisture evolving during the adhesive cure cycle will inhibit chemical reactions between the core and the adhesive, thus weakening the bond which forms. In service, the pressure developed in the cells as described above, when combined with the weak fillet bonds due to poor chemical reaction results in excessive pressure which disbonds the skin at the fillet bonds and results in the skin distortion described elsewhere in this thread as the skin disbonds.

So, the upshot of all of this is that the heat cycle in service facilitates the failure; it does not actually cause it.

It is still a processing problem.

A further aspect to this. A number of people have suggested that there is a softening temperature for adhesives at around 80C (180F). That is true for most 120C (250F) curing adhesives, but for 177C (350F) curing adhesives, that temeprature is closer to 140C (about 285F). That temperature is the glass transition temperature Tg above which the adhesive transforms from a glassy rigid material to a soft rubbery material. However, exceeding Tg does not cause interfacial failures, so if the failures in this case are interfacial adhesion failures, then Tg is not a factor.

Regards

blakmax

PS OK SASless, you can go and have another snoo. (Snoo is the singular of snooze.) Now I think it is time for one more boo.

IntheTin

2nd Sep 2009, 12:08

blakmax,

With all this discussion about moisture, would washing the aircraft down after the last flight of the day have some bearing on this problem?

Brian Abraham

2nd Sep 2009, 12:16

blakmax, thanks for your erudite and informed contributions. Reading your posts one could come to the conclusion that little is known in certain segments of the industry about bonding (Your quote "It is still a processing problem"). I would have thought that the industry would have extensive hands on experience and be mature with little left to learn, with the development of Redux bonding occurring back in 1941. First use on a all metal aircraft was the DH Dove (1945) I think, and used extensively on the DH Comet (1949). The importance of surface treatment of aluminium was well appreciated during the early Redux development period. Am I missing something?

blakmax

2nd Sep 2009, 13:29

Ok, I'll try to respond as rapidly as the messages come in, but soon I must snoo!

Firstly, it is not the wash fluid which are the problem. It is the persistence of humidity, combined with temperature. Remember that higher humidity simply accelerates the problem for a deficient interface. It is a moisture diffusion problem, not a moisture flooding problem. A durable interface is resistant to any hydration mechanism, be it wash fluids or humidity.

Secondly, the major difference between Redux bonding and current processes is that the Redux adhesives were phenolic based and they had a strong acid-base reaction which produced durable bonds. My understanding is that current epoxy adhesives do not exhibit as strong an acid-base reaction and hence they rely on the chemical activity generated by processing of the surfaces. They are therefore more susceptible to hydration, so treatment must address hydration reistance.

Why don't we change back to phenolic adhesives? Because they are condensation polymerisers which produce water as a by-product of the reaction. That water leads to micro-voiding as that water turns to steam during the cure cycle, which results in micro-voiding which reduces the bond strength. Don't be fooled by thinking that vacuum draws such volatiles out; in reality vacuum reduces the pressure and therefore makes the bubbles larger. It is the adhesive which is drawn out, not the volatiles. In contrast, epoxies are addition polymerisers, and they do not produce water as a by-product and hence they are less susceptible to micro-voiding. But the interfacial reactions are more dependent on the chemical activity of the surface and the resistance to hydration.

Now to my favourite subject: Quality. Almost every manufacturer currently uses Quality Assurance (QA) testing to verify structural integrity. Indeed, there are regulatory requirements that are directed towards QA. As already stated, strength tests only produce a snap-shot of current strength. So if the test is conducted before the interface has had time and humidity to hydrate, a false positive result occurs. If combined with NDI (which can not interrogate the interface) then one has a warm, fuzzy feeling that the product is structurally sound. In some cases, such products have an absolute certainty that they will fail in service. Again, I stress that such products can actually demonstrate compliance with the regulations (FAR, JAR, DEF STAN 00 970 etc.)

The problem with QA is that it is nothing more than a "leave pass". If you can meet the requirements of selected tests (usually strength tests based on coupon samples cured with the part) and if it passes NDI, then it is an approved part and you get the weekend off. IT DOES NOT MATTERWHAT YOU DO TO GET THAT LEAVE PASS.

In contrast, there is a far better way of managing quality, and that is Quality Management (QM). In this process, every step of the process is firstly validated using tests that guarantee bond integrity and importantly guarantee bond hydration resistance. Process specifications must match exactly those validated processeses. Then the easy bit. Just make sure it is correctly implemented by certification of compliance with validated process specifications by competent technicians using approved materials that have been correctly handled and exposed in only a controlled environment and using equipment that is appropriate to the task..

Recognise that you can inspect a product and test coupons all you like. You will NEVER change the quality of the product. Manage the quality and it will pass every test you throw at it. And it will actually be a quality product which should not fail in the manners described in this thread.

How can I be so confident? These principles reduced the repeat repair rate for bonded repairs from 43% in 1992 to three bond failures since then, and every one of those could be tracked to technician error or laziness.

We do not have an engineering problem; we have an education problem.

Regards

blakmax

PS If you contact me by PM, blakmax has no "c" in it. We convicts down under can't spell.

PPS SASless. You can take a weekend off after this! Put your feet up, old chap! :-)} (The additional } is because I have a beard. Never did understand why grown men scratch the hair off their face so they look like women!)

Um... lifting...

2nd Sep 2009, 14:44

Wrote a thoroughly boring paper for a supply chain management course on essentially this very subject. The operation in question was a small composites shop manufacturing generally small aircraft panels.
Their longer-term goal is to get into becoming a key supplier for major aircraft OEMs.
We initially found their waste to be rather high. We then found that their rejected parts rate was high. Working backward through the process, we discovered that their clean room processes weren't particularly well controlled, nor were their quality functions managed. We then found that their sourcing for such items as prepreg and honeycomb was large, infrequent lots (this is bad because these items have a shelf life, even when properly stored) resulting in either waste or usage of expired materials.
There were other issues, but in short, they would be unable to become a key supplier for a ISO 9001 customer (which most major aerospace manufacturers are... though having read this thread I begin to wonder if Agusta is) without changing basically every aspect of their supply chain, from procurement through quality.
As I read blakmax's posts, I get the sense that there is a lot of variability in the quality of production in these parts, to include voiding, moisture, improper curing, and whatever other problems can arise in the process. I would think any of these would make the final product unstable with regard to widely varying environmental conditions.
So, yes, blakmax... I think quality may be the thing to be attacked. I've learned a great deal reading your posts. Thanks for that.

500e

2nd Sep 2009, 15:52

I have always thought that ISO9001 was a recipe for disaster, have read lots of manuals, from war and peace to concise & to the point ones.
The problem is if it was wrong in the first place it is then a struggle to rectify, as long as you do it by the manual it must be OK, yes I know you have updates, revisions but the No. of companies that consistently get it wrong suggests there must be a better way.
Some consultants write a large part of them with no in depth understanding of the business they are writing for.
blakmax
Interesting to hear your thoughts of vacuum causing larger voids, we have used to remove excess resins & entrapped air rather than vapours, but I can follow the argument regarding humidity, something else to factor into the equation.
How would you guard against volatiles entrapment other than slow cure and super critical mixes ?
PS
I only scrape small portion of skin is this in keeping with the rugged down under

DOUBLE BOGEY

2nd Sep 2009, 17:03

Hi sorry I have been away, did the tail "Fall Off" or has the TRDS thrashed about following failure and "Severed" the tailboom. Have not had time to trawl all the way through the posts.

Cannot accept that a pilot with heavy feet could make it fall OFF, surely the design overload is there to cater for such an eventuality, over and over again.

If its about glue and sticky stuff not being sticky enough!!! then thats really shocking!!

God was smiling on the crew that day for sure.

DB

heliski22

2nd Sep 2009, 20:12

DB

Subject to formal confirmation, I think we can be sure it failed and "fell off" as you say, mercifully just before take-off. As to whys and wherefores, it really would be worth your while trawling through the posts ans siphon out what blakmax in particular has to say about composites and their capacity to fail based on less then perfect production techniques. Supplemented by the observations of some people who've worked on this part of the aircraft, it all makes for very interesting reading.

22

blakmax

2nd Sep 2009, 23:43

Hi 500e. Just to clarify about applying vacuum to remove air etc. That is a waise move and should be continued, but with some modification. It is important to remove trapped air.

The problem is that epoxies absorb humidity in the cured and especially the uncured state, and many small voids in composites and adhesive bonds are actually caused by release of that moisture as heat is applied. As the resin or adhesive becomes fluid, the bubble will expand because of the low pressure outside and the increasing pressure in the bubble. That forces the adhesive out. Very high vacuum and you end up with an aero-bar (a bubbly confectionery sold down under).
The trick is to make the voids smaller before the adhesive gels. Try this: Apply full vacuum during heat up so that any trapped air can be drawn out. Now once the adhesive has started to flow, reduce the vacuum to about 10 inches Hg (about 5 psi) for the rest of the cure cycle. There is enough pressure to hold the shape, but the voids reduce significantly because of the increase in pressure outside the joint.

Regards

blakmax

airwave45

2nd Sep 2009, 23:46

Spinwing, Um..Lifting,
Thanks for the info, and the kind delivery.
You guys are seated up front while the turbines are running, I have to walk through the blast off the exhausts to ballast the aircraft before you get to settle back to cruise mode, and on specifically the Puma, it's ferkin murderous walking through the exhaust gas.
Assuming the turbines on the 139 and the puma are vaugely similar and the boom on the 139 is hovering (pun intentional) at around 80 Deg C on the helideck, backing upwards is going to put that slightly toasty exhaust gas right back back down on the boom about where it failed.

Now, it only takes 10 to 15 seconds to back up to 100' ato, you try putting yer face in the exhaust blast for 10 seconds and tell me it won't make a signifcant difference to the temperature of yer face (or the boom).

Say all you like about procedures for Cat ab, xy of anything at all, I write procedures for the stuff we do on the rigs, we guestimate at $41,500 / hr to run a rig offshore (inc all the support services) so, it's not small beer.

Just because it's written in a "procedure" does not make it right. these are procedures written by engineers unknown, not the holy grail. Any procedure is subject to review and input (not to be ignored or circumvented willy nilly)

By inference the recomendation from AW is to back the aircraft up from the helideck, putting hot exhaust gas over the boom (unless you can tell me it goes elsewhere ?) then transiton the heli into forward flight (I'm assuming there will be some difference in tailboom loads between backing and forward flight . . ?) with a now distinctly toasty composite, glued boom section.

The only comment from a pax pov on the GH crews is that they have been universally smooth, controlled and impressive.
If commenting on flight profiles is a faux pax on my part, well fly flight profiles that don't raise eyebrows elsewhere . . . :ugh:

I fly gliders, in Scotland, in wave and have the utmost respect for rotary wing pilots.(some very senior rotary pilots also fly fixed wing and have taught me much over the years) None of this is meant as a "dig" at anyone, My arse is on the line as much as yours. I know that, even if some (pompous twats) dismiss it.

If Shell in the Hauge have not ceased 139 ops, subsiduaries elsewhere have.

Europe has _THE_ best composite manufacturing in the world. (notwithstanding input from elsewhere)
It is where F1 composites are manufactured, The best gliders in the world are manufactured, racing bicycle frames are manufactured, in short, Europeans are pretty good at sticking light things together.
Americans, are very good at building heavy and throwing horsepower at them afterwards. (which actually seems to be the way to go in many applications)

I go to play in glued structures and trust them implicitly, however, they are painted white (to keep the heat down) and don't carry kerosene burning turbines. They work very well from 20 Deg C to -30 Deg C.

I believe that the flown profiles are likely to significantly increase the temps on the booms.

From what I know of the properties of the adhesives, this is not a good thing.

I most certainly _AM NOT_ having a "go" at the guys who take me to work, who I believe to be doing a very good job with the equipment they have.

(now if "someone" would look in the direction the bus was going, I'd be even happier . . . .specifically on T.O.) Those rig legs are pretty solid.

Brian Abraham

3rd Sep 2009, 02:40

blakmax, with your experience it would be interesting to hear your thoughts on where we are going with composites, as in the Boeing 787. Perhaps you could start a thread in Tech Log, or where ever the Mods might suggest.

griffothefog

3rd Sep 2009, 06:47

airwave45,

Now you've really got me thinking..... I have never heard of any offshore helis using a back-up take off profile on jack up rigs unless it is away from the deck into clear air over the sea. The 139 does have a very tail down attitude of around 7-10 degs, which would be exaggerated from a pax in the back point of view, and with the normal "towering take off" profile from offshore installations, could feel like a back up profile is being used? If they truly are using the profile you suggest, I would be F*****G gobsmacked :eek:
As for the tailboom theories/bonding issues... I will leave that to some extremely experienced pruners on this thread, but my two cents worth is that its been operating in the gulf of Mexico for many, many years without any issues with the tail or exhaust gas problems, and lets remember, its not made of chocolate!! :ok:

I think its Mr Plum in the repair room with a glue stick :E

212man

3rd Sep 2009, 06:54

Griffo,
I was surpsrised too, but on checking the RFM I see that there is indeed a back up procedure for elevated helidecks! I would have thought it was intended for hospital roof tops and the like - not Jack Up Rigs :uhoh:

By the way, I don't believe this procedure would have any adverse effect on the tailboom heating relative to the overall useage - there are too many variables and too brief an exposure. Heating from exhausts can be an issue in general though, and I know a previous type I flew had a modification to the tail rotor drive cover for this reason.

DOUBLE BOGEY

3rd Sep 2009, 09:05

CAT A (Group A) (Class 1) Elevated Heliport Techniques have traditionally ALWAYS employed a reverse flight profile from the helipad up to and between 80 and 200 feet dependant on the type or any obstacles that may be in front of you.

Offshore Helicopter Operations are conducted (generally) under the alleviations afforded by JAR-OPS 3.517 - Take-off and Landing with Exposure (ie. the inference being that should the engine fail during the "Exposed" period, a safe forced landing may not be assured).

This situation exists partly due to the vaguaries of the offshore environment, turbulence around helidecks and so forth, and partly because it is not really possible to determine the performance of the helicopter prior to arrival at the deck, due to the obstacle environment etc.

The "Exposure" period is also "Theoretical" because if it was not; the only time you would realise that you may or may not have exceeded the exposure period, is when it actually happens, and guess what, you are already exposed.

The implied risks associated with operating with an exposure period during take-off and landing are mitigated by the requirements of the approval, in that to be so approved you must operate the helicopter under an engine reliability programme, have a full HUMs kit and protocols fitted for trend monitoring, AND MOST SIGFNIFICANTLY operate the helicopter IAW the procedures acceptable to the authority.

These proicedures were first published under JAR-TGL 14 and follow in principle the older HAPs modelling studies for offshore operations. These procedures DO NOT incorporate a reverse flight take-off procedure. They are based on a vertical take-off to a TDP of approx 25 feet and then a rotation away from the deck using the momentum of the intitial vertical climb to hopefully clear the tail should an engine fail. There is no performance solution in play and safety is assured by the close monitoring of the 2 engines in play as indicated earlier to provide an assurance within defined acceptable probabilities that neither engine will fail .

I do not know what kind of approval the AW139 has for offshore, and although the Operator may be JAR-OPS, the Procedures they are required to operate must be acceptable to their responsible Authority which may differ from those currently applied in the UK-NS sector.

From a performance point of view, ignoring the turbulence environment for a moment, a helicopter conducting the CAT A approved Elevated Helipad Profile, at the correct mass for WAT, has the capability in the event of OEI to either, Re-land safely back on the helipad, or, continue the take-off safely, meeting all required safety margins, depending on the point in the profile at which the OEI occurred.

For a helicopter operating to TGL-14 and the HAPs model, this absolute level of safety in the event of OEI is not provided, but may be there depending on the ambient operating conditions and obstacle environment.

THEREFORE - if some offshore operators are electing to (or required to) use a CAT A technique offshore, from a performance point of view, provided the ambient conditions and obstacle environment is suitable, the helicopter is being operated to a greater level of safety than one using the JAR-OPS 3.517 Class 2 With Exposure alleviations.

Hope this helps.

blakmax

3rd Sep 2009, 10:20

Brian Abrahams, Started a new thread under Tech Log

Regards

Blakmax

spinwing

3rd Sep 2009, 10:33

Mmmmm ....

Ok ... Heli Ops (Offshore) in the UAE at the moment do NOT have to operate to PC1 requirements.

I do not know if they have to in the State of Qatar .... I very much doubt it.

As 'Griffo' and I both know the decks on most 'jack ups' used in the Arabian Gulf are generally a bit on the small size and therefore I really doubt if a "back up' type take off procedure is used ... it'd be a short route to either disaster or un-employment to do so! (and I might add I recently lost a collegue to what has been assumed to be a 'backing up accident' so I'm a bit sensitised to them).

As "DB" indicated ... the 'normal' take-off is 'usually' to a hover then to lift to 25-40' above the deck if nothing untoward has occurred then 'rotation' happens and using the a/c momentum to clear the deck if it were to go wrong from that point.

NOW can we keep this thread on track with the regard the 139 bonding issues ... especially as I see there is another thread has been started to take care of the rig takeoff 'red herring' ...

Cheers

DOUBLE BOGEY

3rd Sep 2009, 10:47

Spinwing, apologies for the thread diversion, got carried away for a bit there!!!

DB

blakmax

3rd Sep 2009, 11:07

Spinwing

I am not too sure if you want to continue my discussions on adhesive bonding issues here, or should I also start a new thread eleswhere? I did start a new thread on 787 composite issues in response to Brian Abraham's request because that is really far from the AW139 issue.

I'll take guidance here. Do I start a new thread to talk about generic bonding issues, do I continue posting here or do I simply make love elsewhere (F**&&* off)? The subject is of relevance to the current thread but the subject may be more happily located elsewhere if requested.

Regards

Blakmax

Senior Pilot

3rd Sep 2009, 11:16

blakmax,

I think that most of us agree that your dissertations in this thread are both informative, and very relevant to the AW139 topic under discussion. Unless you specifically want to create a new thread on Rotorheads, I would rather see your posts staying here :ok:

spinwing

3rd Sep 2009, 11:20

Mmm ...

blakmax ....

I consider your posts to be extremely relevant to this thread ... and also very educational.

Please, please do continue ..... I was more concerned with the possibility of a "thread hijack' with regard rig take offs ... which I really thought bore no relation to the 139 de-bonding issues.

Your input IS very valuable!

Cheers :ok:

blakmax

3rd Sep 2009, 12:49

Guys (and girls) and for the down under crowd, Blokes and Shielas (strewth, Bruce that is laying it on a bit thick!)

The rig focussed group really do have a point. If the thermal shock loading is high, the moisture evolved will produce sufficient cell pressure to force the skin off weakly bonded sandwich structure. Recognise that this is the initiating mechanism, not the root cause of the problem, which still remains a processing issue resulting in weak core to adhesive bonds which fail when subjected to excessive cell pressure.

It is still a bonding issue, Agusta.

Regards

Blakmax

sox6

3rd Sep 2009, 14:12

Unless a report has been issued or you have examined the damage yourself are the highlighted words not a bit presumptuous?:

Recognise that this is the initiating mechanism, not the root cause of the problem, which still remains a processing issue resulting in weak core to adhesive bonds which fail when subjected to excessive cell pressure.

It is still a bonding issue, Agusta.

Or are you in fact the messenger for someone with an ax to grind?

[email protected]

3rd Sep 2009, 15:02

sox6 - blakmax has done an excellent job here in highlighting shortcomings in people's knowledge about the construction of the aircraft they fly and operate.

Unless you can come up with a really good reason why a tailboom that had been repaired and declared airworthy should suddenly fail in such a spectacular fashion I suggest you let blakmax educate you as he doing to the rest of us.

We have all taken composite technology for granted and assumed since it replaced/enhanced aluminium construction that it must be as reliable and failure proof - after all, all the manufacturers are using it so they surely wouldn't compromise safety for a few more kilos useable AUM would they?

Blakmax has clearly detailed how, unless the composite is put together properly in the first place, it is doomed to failure at some point and environmental factors are simply the trigger and accelerant.

DOUBLE BOGEY

3rd Sep 2009, 15:17

CRAB

Thanks for putting us all in our place, however, there is no information released that supports any of the hypothesis that has been put forward so shall we keep all the other SWAGs open as well.... like did the bump on the helideck weaken it.

If it is a heat/debonding type issues then it is clear that none of us have learnt anything at all since the first SMS began.

A Greek SMS Reported some while back that a de-bonding issue caused by excessive heat had caused a catastrophic mainplane failure and a fatality as a result. The Pilot I think was called Icarus!!!

DB

sox6

3rd Sep 2009, 15:25

I am always open to being educated. I simply challenge the level of confidence of some statements being made. The rest us may have a low level of knowledge ourselves but automatically assuming someone with more must have all the answers is just being naive. Lets not confuse knowledge about composites with knowledge of this failure.

If a psychology professor was posting here about a recent accident and discussing the pilot's errors with that level of confidence would that be automatically accepted just because of a few academic qualifications?

Brian Abraham

3rd Sep 2009, 15:57

I am always open to being educated. I simply challenge the level of confidence of some statements being made. The rest us may have a low level of knowledge ourselves but automatically assuming someone with more must have all the answers is just being naive.
sox6, if you want to do yourself a favour, go back through the thread and you'll find blakmax lays his credentials on the table. Impressive they are, an expert in composites. He is not alluding to what caused the failure, but giving an education in the problems associated with the use of composites.

sox6

3rd Sep 2009, 16:06

Brian If you do yourself a favour and look on this very page you will find I have specifically challenged some apparently very explicit statements above that do more than simply allude to this failure.

JimL

3rd Sep 2009, 16:30

It is not clear whether we should be posting on this thread or the other one - started to discuss the take-off procedure.

AW have developed a take-off procedure for the AW139 which is in conformance with the PC2e philosophy; further, they have had it certificated as a CAT A helideck procedure. It is analogous to the North Sea HAPS procedures but, when flown into a clear sector (the 180 degree obstacle clear sector), it provides engine failure accountability up to its new MCTOM of 6.8t.

In fact it is good enough to provide deck-edge clearance (and of course drop down) even when taken off from the TD/PM of a helideck of any size in the North Sea and elsewhere (212man, remember the discussion we had some time ago). With a take-off mass of 6.4t, the drop-down is minimal up to about ISA + 20 with zero wind.

As the greatest danger to the aircraft in a helideck departure/arrival is collision with obstacles, heliport procedures should not be used. There is less danger with a PC2-with-exposure procedure.

Having said all of that, we are all in the dark as to the procedures used, and reasons for using them, unless one of them is prepared to join this discussion.

Jim

PS - FSTD TGL # 14 (new)–- Guidance on the qualification of Electrical Motion Systems for FSTDs

TheOldTimer

3rd Sep 2009, 17:56

Dear me what a shame, a very educational and fascinating series of posts and a thread that was in my humble view, one of the best I have seen recently, seems to be going in a direction that really adds little to the original development of the discussion.

Capt JB

3rd Sep 2009, 17:56

Ok:O, interested to know what the :mad: happened there. I have A7-GHC as registration, from web records.
If anyone has the serial number and operator of this bird it will be very much appreciated to keep up with my database

Sandy Toad

3rd Sep 2009, 18:18

BT 139-193 and 139-194 Issued .Both Mandatory Precautionary Inspections for signs of de-bonding. First one for Helicopters that have experienced "Tail Scrapes, Tail Strikes or Tail Bumps". Latter for rest of fleet.

[email protected]

3rd Sep 2009, 19:05

DB, I would be very surprised if the tail strike hadn't weakened the structure - surely that is why it had an AW approved repair carried out. The point we keep coming back to is that the industry may not be quite as clever as it thinks when manufacturing or repairing composite structures.

Sox6 - if you read blakmax's post that prompted your reply in context, he is referring to the fact the environmental issues such as OAT and exhaust gases can precipitate failure but only if the bond is not perfectly manufactured - I read this as a generic statement regarding compositis and not one specifically dealing with the tail boom collapse.

He has been critical of the aerospace industry as a whole as well as the regulators, he just happened to mention AW in that post.

500e

3rd Sep 2009, 19:29

DB
".Like did the bump on the heli deck weaken it."
Now there is a scary thought, I understood it was passed as serviceable after repairs by AG.
sox6
blakmax is questioning the methods\ quality of build and giving his suggestion for failures of this type, he would appear to be able to back up the talk with results in the field, he has laid out types of failure with suggestions as to the reasons and a rough guide as how to examine same, the thing I find worrying is there appears to be no way to check for incipient flaws?.
So the only way to prevent failure is in build quality which is what he appears to be saying, as Crab said,
"Unless you can come up with a really good reason why a tail boom that had been repaired and declared airworthy should suddenly fail in such a spectacular fashion I suggest you let blakmax educate you as he doing to the rest of us".

Thanks for the information you have imparted both here & in PMs direct BM . We have been doing in nearly correctly, without understanding the reason! will reduce vac to around 5lb rather than 8\9 we were using. This is on a boat hull not aircraft, relax everyone:ok:.
I have no knowledge of bm other than he gave me some Free pointers regarding bonding, which as this is his living I found most helpful.

blakmax

3rd Sep 2009, 22:30

To quote a famous Australian "Crikey!" Things were quiet last night and when I awoke this morning all hell has broken loose!

For Sox6. You do have a point, and we will eventually have to wait until the report is out before we can confirm my theories. However, my discussions were not restricted to just this incident. If you go back through the thread you will see that several posts relate to bonding issues in the tail boom on a number of other aircraft. Indeed there are examples of failures after very short TIS. The desriptions of the disbonds in all of the other cases are consistent with interfacial failure between the adhesive and the skin or the adhesive and the core. These types of failures are indicative of processing problems. There are very few actions which an operator or maintainer can do that will result in interfacial failures.

The next issue is; Will the report actually be undertaken by people who can recognise bond defects? I mean no reflection on investigators, who really do a good job, but adhesive disbonding probably constitutes a very small proportion of the investigations so there is only a limited opportunity to develop the knowledge base necessary to differentiate between the different types of failures and to identify the probable cause. I am unaware of any formal training available in adhesive bond forensics. Therefore much of the knowledge base is derived from hand-me-down knowledge which is only as good as the original data.

I have also seen a number of glaring errors in formal reports on accidents involving another helicopter type. In one case, the investigator found what he beleived was fatigue markings in an adhesive bond. Unfortunately these features were in the middle of the adhesive layer ahead of the failure surface and the actual failure was at the interface. There is no way any fatigue in the middle of the layer can be related to a failure which occurs at the interface.

In another case, the investigator showed an example of a good bond from an area away from the failure site. In fact that "good" bond exhibited about 50% microvoiding which was not even commented on.

From my side, this subject has been my primary responsibility on a day to day basis for over 37 years. I have had the luxury of having the time to develop expertise in the field, the opportunity to access a wide range of bond failures and to identify processing issues which cause specific failures. These skills enabled me to write an engineering standard on composite and adhesive bonded repairs, write two handbooks on repair design and repair bonding processes, develop three training courses and participate in the development of two other courses. Since the implmentation of these measures, the bond failure rate in repairs fell from 43% in 1992 to only three cases since then and in each case the cause could be easily attributed to poor practices.

The reason I have been getting involved with this thread is to increase people's understanding of composite and adhesive bonding issues.

You will also note that I have offered my services to Agusta on investigating any bonding issues for this specific case as well as those reported by others in this thread. I acknowledge that until something more definite is available, this still remains a WAG.

Regards

blakmax

spinwing

3rd Sep 2009, 22:38

Mmmmmm ....

..... The reason I have been getting involved with this thread is to increase people's understanding of composite and adhesive bonding issues.
.....

And if I might say ... your doing a bloody good job of that too!

Thank you :D :D

blakmax

4th Sep 2009, 00:04

Thanks for the comments spinwing. I did forget to address the issue of axe grinding. Up until this incident was reported on PPRuNe, I had never seen an AW139 and I have never had any contact with Agusta. I have only had one small contract with another rotary craft manufacturer and I have no ongoing relationship with that company. I have also stressed in the past that these observations are relevant to a wide range of manufacturers of rotary and fixed wing structures. Hence I have no axe to grind with any specific company and I have no interest in supporting one company against another.

On a different tack, this particular failure presents a unique opportunity to investigate the integrity of the bonds in the failure area. Usually, the samples have been subjected to crash related forces and environments which often mask the evidence or at least complicate the assessment. This example has not experienced fire, has not impacted the ground at high speed, it has not been imersed in water and all of the parts should have been recovered. It may be possible to determine the direction of disbond growth during the failure, so it may be possible to determine the locus of failure and the probable cause without the conclusions being clouded by unrelated events.

Regards

blakmax

TheOldTimer

4th Sep 2009, 07:16

Blakmax.
I would like to add to Spin, crab and others comments regarding your valuable, very informative and nicely placed humour that has given us all a valuable insight into the black art of composite s and the problems and considerations that designers and manufactures need to address. Since my first exposure to the inclusion of these structures in aircraft design, initially as non load carrying and graduating to become critical structures, I admit to having been concerned. This in part I felt to be an old farts reluctance to accept change. A responsibility of the designer and manufacture, policed by legislation to regulate and ensure that the best and safe practices were followed. Your articles on the scientific side of this has been informative in the extreme.
To finish, education and understanding is important as we all know, openness in the release of information to the industry also critical. This was not generally the case in my time.
Thanks blakmax and please continue to educate us all, let’s not return to the dark side of the past.
Thanks all for indulging an old timer well past his sell by date.
:ok:

blakmax

4th Sep 2009, 09:57

Thanks OT. I really appreciated all of the support both on and off line. Adhesive bonding is a truly amazing technology when the basic fundamentals are understood and correctly implemented. However there are problems with the certification basis and these deficiencies are combined with management that follows what worked last time (but didn't really work if they understood the basics). Iwas astounded to find that over 75% of aircraft designers base their designs on an average shar stress approach which was shown to be incorrect as far back as 1936. The only way that these designs work is that they have a high level of conservatism built in to the design allowable stress levels, and they undertake extensive testing. The stupid thing is that if they udnerstood the fundamentals of adhesive joint design developed in the late 1960's not only would their designs be just as conservative, they could save millions in certification costs.

It is actually possible to design a bond between aluminium up to about 0.15 inches thick such that the aluminium breaks, not the adhesive. If this rule was applied, then every one of the thousands of tests normally undertaken would result in only breaking the metal. So why do all of those tests? Design the joint, undertake a few tests to demonstrate that the bond is stronger than the metal and certification is a lot cheaper.

Regards

blakmax

widgeon

4th Sep 2009, 10:03

It is funny that despite all the advances in technology one of the best tests for a debond is still the tap test with a quarter ( or probably 10P in the UK ).

500e

4th Sep 2009, 11:17

The tap test only finds relatively large dis-bonds not a myriad of small ones, read post 78 of this thread, how do AG suggest the pilot checks for problem ? or does it require an engineer to examine before every flight.
The post regarding bubbles would suggest there was a considerable void to start with, presumably this boom was in service & subject to testing on on going basis.

twinpac1958

4th Sep 2009, 12:20

For those that do not know, Fuselages AND tailbooms are manufactured in Poland and trucked to Vergiate (It) for assembly.

spinwing

4th Sep 2009, 12:27

Mmmmm .....

....For those that do not know, Fuselages AND tailbooms are manufactured in Poland ......

Vaguely remember mentioning this back on post #11 ...... but it was sooooo long ago now :hmm:

Cheers .... :E

twinpac1958

4th Sep 2009, 12:29

Anyone who has tried to contact Agusta for help should know that nobody would have picked up the call anyway.....

Anytime I've tried to call anyone at Agusta I have never been successful and sad part is that there isn't even voice mail....

iuk1963

4th Sep 2009, 13:34

"For those that do not know, Fuselages AND tailbooms are manufactured in Poland and trucked to Vergiate (It) for assembly."

ok, so what...

9Aplus

4th Sep 2009, 13:50

....For those that do not know, Fuselages AND tailbooms are manufactured in Poland ......Was in PZL Swidnik 14 months ago, only main part of 139 airframe was produced
there.
Concerning tail booms, Bell 412 booms are in production there for long time now.

Anyone who has tried to contact Agusta for help should know that nobody would have picked up the call anyway.....That may be true for Philly but not for Vergiate, Cachina Costa or Somma Lombardo...:*

Therefore your claim is bit unfair.... even to me, who stand besides and wonder
what is going on with that tail booms.:uhoh:

Was on 139 test ride on 2008. for approx 45 min, for me that was one of the best experience in helicopter world. :ok:

sulli26

4th Sep 2009, 15:36

http://www.caa.co.uk/docs/33/20090904EASAAD20090198E.pdf

Algy

4th Sep 2009, 16:07

Anyone able to shed light on the common factor in the seven aircraft identified in today's EASA AD as requiring inspection within 5hr rather than the 25hr for the rest of the fleet? Serial numbers: 31006, 31020, 31022, 31042, 31136, 31157 and 31248.

Shell Management

4th Sep 2009, 16:36

They all have had a backend bump and no, they are not all from the sandbox.

Algy

4th Sep 2009, 16:42

Indeed not: one in sandbox, two in GOM, one Norway, one Scotialand, one Tokyo Police and one Pakistan Army!

spinwing

5th Sep 2009, 00:10

Mmmm ....

And of course .. c/n 31028 although not mentioned ...... no longer exists :(

TheOldTimer

5th Sep 2009, 06:49

I don’t wish to repeat any findings already posted but had news that a operator had found de-bonding detected by the tap test and visual inspection on the compressively loaded side wall of the boom., at about the location seen as the failure point on the photographs back in this thread.
As blkmax has now given us all a understanding of the issues related to design and manufacture and the strength degrade that bonding failure can cause, would it now be the time to strengthen our understanding of repair techniques to these structures. Would repair be an option? In these cases,? I feel not, however, if you could enlighten us blakmax I certainly would be keen to have your views.
:ok::confused:

ARRAKIS

5th Sep 2009, 07:06

PZL Swidnik is manufacturing ONLY the fuselages for the AW139. Information about the tailboom production is :yuk:.
AW139 tailbooms (and some parts of AW101 fuselage) will probably be produced now, that AW will became the owner of the factory.

Arrakis

blakmax

5th Sep 2009, 10:41

Hi OldTimer. Thanks for the information that is is a compression loaded area because that adds to the suggestion that adhesion fillet bond failure may have played a part. For the skin to separate there needs to be an out of plane force and that would come from compression buckling of the weakly bonded skin. Once the disbond exceeded a critical size, the structure would be susceptible to collapse by overall buckling. (In deference to previous comments, this is only a suggested failure mechanism which is consistent with features observed in other bond failures in this region together with the stated loading mechanism. Examination of the part by a person competent in adhesive bond forensics will be required to confirm the cause of failure.)

I'd also express some concern that if the processing is to blame, and if that process has been used elsewhere, then there is a reasonable probability that the problem is not just limited to the disbond area; it just means that the structure has not seen loads that would result in the skin separating in other areas. The real solution is to eliminate the causes during manufacture rather than chasing damage which has a high potential to continue to occure in other locations.

Firstly let me state that I do not have design loads, and in any repair to critical structure, validation by FEM and/or testing is appropriate.

The first step is to verify the failure mechanism. Adhesion fillet bond failure is characterised by an absence of adhesive on the core. If adhesion fillet bond failure is confirmed, then an inspection of the core itself is essential, because fillet bond failure is often coupled with another core defect known as weak node bond failure, where the core cells separate at the sites where the cells walls were bonded together in the process of manufacturing the core. This inspection is best performed using a Mark 1 eyeball and a magnifying glass. Any sign of node bond failure necessitates complete removal of the core. Repair is the same as for aluminium core below.

Now for the manner of repair. If I were to repair adhesion fillet bond failure, the first thing to do is to remove the skin surrounding the disbond until the adhesive actually fractured, in contrast to the fillet bond separating from the core. For nomex core dry the core using a vacuum bag and heater blankets for at least six hours at a temperature of 85C +/- 5C. Bond an insert repair over the core and then bond a splice repair between the insert and surrounding skin.

For aluminium core, the story is less friendly. Because the core surface at the fillet bond zone has hydrated to enable adhesion failure to occur, it is prudent to remove all core in the area where the adhesion fillet bond failure has occurred. Fabricate a replacement core insert and bond that in place after preparing the skin at the bottom of the cut-out. Use a film adhesive to bond the core to the face sheet and a foaming adhesive to splice the core to the insert. Sand the core to the OML shape and bond a repair patch over the core and adjacent skin.

Now if the failure is between the skin and the adhesive, then it is a different matter. If the adhesive has a glossy appearance and has no visible reproduction of the surface of the metal, then it is probable that there was a pressurisation or fit up problem and the adhesive was never in contact with the skin. If the defect is limited to a specific area (verified by removing skin until the adhesive fractures from the core) then an insert repair may be possible.

If the skin shows signs that it has been in contact with the skin (the surface is not as glossy and there is replication of the surface of the metal) then there is a problem with the surface preparation method due either to contamination or the use of an ineffective process which does not prevent hydration. No repair is possible.

If I haven't bored everyone to death, I could discuss how to heat the structure because most aircraft SRMs contain heating methods that even Walt Disney would not be able to make people believe.

Regards

Blakmax

212man

5th Sep 2009, 11:54

Now I see why the S-92 uses tin and rivets!

The Sultan

5th Sep 2009, 15:40

212,

Too bad that the 92 guys did not spend a little more time on the transmission lube system.

I am sure Agusta will take care of the issue and not wait a year to see what happens next.

The Sultan

Hilife

5th Sep 2009, 16:09

Just like they did for the the A109E tail boom and the AW101 TR Hub cracking. Come to think of it, the 101's MGB reliability isn't anything to brag about either.:rolleyes:

pants on fire...

5th Sep 2009, 16:54

From the Sultan.

I am sure Agusta will take care of the issue and not wait a year to see what happens next.

You might well know all about that?

Didn't Bell play a significant part in the design of the problem parts? Maybe the Sultan actually designed it? :ouch:

9Aplus

5th Sep 2009, 19:20

Bell pays for 50% of 139 design R&D, most of the work done in Swidnik and
AW Italy

ARRAKIS

5th Sep 2009, 20:55

I was told that another problem is that traditional Non Destructive Testing Techniques don't work on composites so finding voids in the structure or identifying cracks is much more difficult. Any gurus know what techniques are being used nowadays?There are some new techniques for composites NDT. The first one that comes to my mind is active thermography. The tested element is heated/illuminated with a strong light source and observed with an infrared camera. Internal defects like delaminations, etc.. are becoming visible, because they are affecting the way the material heats up/cools down.

Arrakis
PS. What about other AW helicopters tailbooms?

blakmax

5th Sep 2009, 23:32

I was told that another problem is that traditional Non Destructive Testing Techniques don't work on composites so finding voids in the structure or identifying cracks is much more difficult. Any gurus know what techniques are being used nowadays?

Not true. As well as theremography mentioned by Arrakis, conventional ultrasonics are effective in finding delaminations and disbonds. Eddy currents are not effective, but then composites rarely exhibit the type of cracks that occur in metals unless they are associated with other damage which can be detected using ultrasonics. The old tap test is effective in finding delaminations close to the surface, but not so effective in thicker composite structure, and while it may be effective in finding disbonds and delaminations, it is important that other more accurate methods (ultrasonics or thermography) are used to determine the size of the defect prior to repair.

Remember that NDI is only capable of finding air gaps. It can not find weak bonds which are still in contact but which have a high potential for subsequent failure. I once asked an over-confident technician to tell me where a sample I had just made actually had good bonds. He easily found the deliberate void in the centre and then confidently told me the rest of the sample was well bonded. Pity it was joined together using double sided tape! NDI is effective at finding bad bonds. It can not give total assurance of good bonds.

Regards

Blakmax

[email protected]

6th Sep 2009, 06:53

So my statement is true then - eddy currents and visual inspections (traditional NDT) don't work but modern techniques (thermography and ultrasound) do:)

blakmax

6th Sep 2009, 07:53

Yes Crabby, these are a waste of time. Eddy currents require a conductive medium and while carbon composites do set up eddy currents, the nature of damage in composites (delamination, fibre fracture, intra-ply splitting, fibre pull-out) does not lend itself to detection by that method.

Visual is only of any value when the defect is excessively large so that the skin distorts. Composites are linear elastic to failure and therefore they tend to spring back into place. In sandwich structure visual may be of assistance only if the core is crushed sufficiently to prevent the skin springing back.

The most rediculous visual inspection prize goes to a certain brand of helicopter which issued an AD requiring visual inspection for disbonds before every flight. Now, a dark disbond on a black blade viewed against a bright sky? What chance?

Regards

Blakmax

ARRAKIS

6th Sep 2009, 08:25

So my statement is true then - eddy currents and visual inspections (traditional NDT) don't work but modern techniques (thermography and ultrasound) do.
Basically yes.
The advantage of thermography over ultrasound NDT is speed. I don't know exactly about ultrasound, but thermography can also detect weak bonds. No need for any voids. This is all about the heat flow through the bond and some smart data processing.

Arrakis

blakmax

6th Sep 2009, 09:35

Hi Arrakis. I have seen some pretty impressive stuff using thermography for larger area disbonds or delaminations, but I was unaware of the ability to find weak bonds. Do you have any published reference material for that? (Interested, not disbelieving.)

I could see the possibility for localised areas of weak bonds, but my understanding is that you need to have areas of good and bad bonds to enable a comparative assessment. In other words, thermography does not provide a numerical output based on a known standard, it relies on variations of temperature within the zone being inspected. If the entire region is weakly bonded, then such a comparative assessment may not be possible.

Regards

blakmax

blakmax

6th Sep 2009, 12:44

Sorry Old Timer, I forgot the usual repair method specified by OEMs for disbonds: INJECTION REPAIRS.

Injection repairs involve drilling holes in the skin of sandwich structure and injecting fresh adhesive to "re-bond" the area. Remembering that adhesive bonding involves chemical reactions on clean, chemically active surfaces which are treated to develop hydration resistance, these requirements are physically impossible to achieve by just injecting fresh adhesive. You can not guarantee that the surface is clean even if you flush the holes with solvent. You certainly have done absolutely nothing to activate the surface (solvent cleaning alone does not do that). And for metals, you have not treated the surface to prevent hydration. Let me be blunt: Injection repairs are nothing more than bovine excrement.

The only two things achieved by injection repairs is that the air gap is filled so that (as discussed elsewhere in this thread) NDI no longer finds the gap. The other thing achieved is that the technician can sign off the repair with a warm fuzzy feeling he has fixed the problem. Structurally, he has made not one ounce of difference. This situation is true for any disbond, composite or metal.

If you are looking for a short term fix, my solution (tongue placed firmly in cheek) is to paint the disbond area bright pink, because that is a better structural solution. Firstly, you have not perforated the skin and therefore have prevented moisture ingress which would cause corrosion or further disbonding. Secondly, you have identified the area so that you can cut down future inspection time. Thirdly, if you carefully map the area, you can monitor disbond growth rates. And fourthly, the aditional paint has added to the corrosion protection of ther area. Structurally, the result is the same. You have done nothing to restore the strength of the disbond.

Now I have recently had some comments that my statements are too definitive. Let me stretch the boundary there. If there is one single shread of evidence (apart from removing the NDI trace) from ANY manufacturer, researcher or regulator that injection repairs actually work, please provide the reference.

Does any one want to take a bet of ten pints of Guiness that I get no response?I have a serious number of examples where injection repairs have not rebonded anything. I also have a number of examples where injection repairs have led to in-flight failure of fairing panels approx. 10ft x 8 ft which impacted the fin, HSTAB and rudder. Had this aircraft been in a nose-down pitch with the HSTAB raised, the panel may have jammed the HSTAB with disasterous consequences. All because of an injection repair. In another case, an aircraft lost a rudder due to injection repairs performed at manufacture. Pictures available on request.

So why do we see injection repairs in every repair manual? Because these repairs are "grandfathered" to the next model and the regulators accept that process. In this case "grandfather" has dimentia!
Regards

blakmax

TheOldTimer

6th Sep 2009, 15:07

Thanks blakmax for the info re my question. Maybe the pink paint job would be better, certainly would add to the visual effect overall. As I understand a number of aircraft have now been found suffering from various levels of de-bonding, the next interesting chapter will be the discussion on the approved? Repair scheme from the big A, if and when it reaches this thread. I still think replacement is the safe option.
:ok:

TheOldTimer

6th Sep 2009, 15:25

Sorry blakmax forgot to mention, had a snap sent to me of the approved testing device yesterday, very pretty looking miniature hammer, has a part number and everything. Maybe this move towards high tech is the reason for finding more of those pesky voids that eluded the two pence piece. Retirement has its advantages, a little worrying as off on a holiday soon and flying Airbus, aren’t they mainly composite??? Sorry ignore that, thread drift.:cool:

spinwing

6th Sep 2009, 23:47

Mmmmm ....

Tap Hammers .....

Hey OT .... I gave mine away to an eager apprentice years ago ..... I now find my self back at the Lathe making myself (and a few chosen 139 mates) new tap hammers .... what a way to spend ones leave time????

Wonder if they have to be approved so the "Tap" can be calibrated??? :confused:

:E

212man

7th Sep 2009, 00:17

I hope the 787 designers are taking notes.....

TheOldTimer

7th Sep 2009, 03:33

Spin, maybe you have to make a tap test audio recording of a series of toppings and send to A, they then issue a chit. I wonder if the hammer heads are cracked, they could then indicate a false pesky void, hence the sudden increase in the discovery of them. 12 ish in one day seems to be the record, so a mate informs me. Wonder if that’s true? I also understand that sales of pink paint have increased and stocks are being shipped to Italy, so with little hammers and pink paint, problem solved.
OT
:ok::sad:

Skin King

7th Sep 2009, 06:34

Sorry for my absence over the last week but I've been busy tapping out 139 Tailbooms.
As mentioned before I am the poor technician at the coal face having to deal with composite problems on a daily basis now for nearly 30 years
(yes blakmax I have work with you a long time ago. Say hello to my mate Jimmy S for me.) What I have found over time is that we as technicians, particularly on helicopters, are restricted by repair schemes laid down in the OEM's manuals to the use of cold bonded repairs only.
I know Airbus and Boeing use, in the majority of cases, hot bonded repairs.
Why the difference between fixed wing and rotory wing composite repairs ? This is a question I ask myself everyday, I can only put it down to the fact that most helicopters are operated and maintained in less than favorable locations to carry out hot bonded repairs.
It may also have something to do with the cost associated with having approved hot bonded repair capabilities.
I will go on record as saying I detest cold bonded repairs with rivets as I have seen everyone fail, but I also understand that it would be near impossible in most cases to carry out a hot bonded repair whilst in the field.
What I believe Bell and Agusta should do is place a supplement in their SRM's that allows the customer the option of carrying out a hot bonded repair.
So OK, I'm stuck (no pun intended) as with all other technicians in the field with performing cold bonded repairs ( this is where you come in blakmax) how do we improve our processing of the repair area and parts and to ensure we get the best bond possible? I'm talking surface preparation and core cleaning here...TAKE IT AWAY BLAKMAX :ok:

One last thing regarding NDT of composite structures in the field I use the Eurocopter tap hammer which is simply a 2 1/2 inch long cylindrical steel roller bearing. Works a treat and has cost Agusta millions.:)

blakmax

7th Sep 2009, 09:53

G'day skin king. I am still trying to work out who you are so I have sent you a PM. Top right near log-out.

By the quality of your comments about hot bonding vs cold bonding and the integrity of SRM procedures, my persistent nagging over the years has paid off, at least in your understanding of how things could be improved.

You have highlighted one of the issues that stuns me. Why must the SRM only have the dumbest repair? Surely there has to be a grading of capabilities. I fail to understand why SRMs contain only the repair which is easiest to perform while at the same time providing the worst outcome for everyone, including organisations which are capable of higher level repairs. You would presonally know that the only way RAAF at Amberley ever achieved a virtual elimination of repeat repairs was to scrap the OEM processes and adopt scientifically validated procesess which had a verified capability to produce bond durability. This is why I wrote DEF (AUST) 9005. And you would agree, this is not rocket science! I have taught electricians to perform bonded repairs, so what is so difficult that we must continue to use ineffective processes because Blogslovia doesn't have an appropriate level of competency? Can't we have a grading level for authorised repair stations and assign repair approvals on the status of the repair station?

I suggest that many OEMs do not actually comprehend the impact of their own SRM processes on repair integrity and durability. Further, many hide behind a fascade of OEM infallibility which could not sustain knowledgeable scrutiny, as you know from the numerous cases I managed for RAAF. I managed to convince and then assisted a certain large manufacturer in St Louis to re-write their manuals and any organisation who adopted the amended publications gains the benefit of those changes. So it is possible to work with manufacturers to correct deficient bonding practices.

The frustration I have is that the only two ways I can see to force such changes are by 1. Operator pressure to reduce maintenance costs and 2. Legal action after some unfortunate people have died and their dependents seek legal redress. Logic is not effective against a company position of corporate infallibility.

Regards Blakmax

PS respond to the PM.

victor papa

7th Sep 2009, 15:43

Went through my OEM manuals today to look for a repair for a impact mark on Nomex honeycomb with aluminium skin with possible debonding in the surrounding area. Blakmax, you should be happy to here that drilling a hole and injecting something was not an option. In order to comply with the repair I had to get the local aviation approved AMO in who specialises in composite repairs and has stock and experience in working with all the listed(of which there are quite a few) materials to be used during the different stages of the repair.

Sorry, forgot to mention it is not even a structural part or load carrying part at all.

Skin King

7th Sep 2009, 23:51

To anyone at Agusta who is out there and willing to listen,

PLEASE STOP ISSUING REPAIRS FOR LARGE DISBONDS TO AW139 TAILBOOMS !!!!! :=

They are difficult to accomplish even in a facility with all the resources, use antiquated techniques and simply do not restore the panels to their original strength.
My advice to any operator who chooses to carry out these repairs is to monitor the area around the repair on your daily and after each flight.
Or better still just ask Agusta for a new Tailboom.
I don't want to appear all doom and gloom but in all my years invovled with aircraft I have never come across a more important safety issue in regards to a helicopter structure. It would be nice if Agusta could at least forward an information letter to operators detailing the actions they are taking to fix this immediate problem. Sure they have aluminium core material now in place of Nomex and they say they have improved QA standards, well give us the details...What does this new core material achieve? What are the new QA standards? Are they validating each panel manufactured? I believe we as operators and the pax we carry have a right to know. And if there was an obvious shortfall in QA standards in the manufacture of past assy's we should know that as well. I'm sure if a car maker had a serious safety issue with one of their products it would be recalled regardless of the cost to them.
Time to step up Agusta.
I know, it was more than 2 cents worth :O

blakmax

8th Sep 2009, 02:55

Skin King, You are saying what a number of others have said to me in PMs. The real (and only valid) solution is to get the manufacturing correct so that disbonds don't occur in the first place.

I am advised by PM from some members that the repairs to which you refer involve large area disbonds and the repairs involve the use of a room temperature curing paste adhesive with lots of blind fasteners added. There are multiple issues with this repair. Firstly, I'd like to know what surface preparation was used on the skin and doubler prior to bonding. (Please don't tell me it is just a solvent wipe!) Secondly, unless they do something to remove the moisture from the core, the adhesive will not bond to the core. Thirdly, the paste adhesive will most probably just run down the cells, leaving barely any adhesive to form fillets to the core. Fourthly, the fasteners will provide a moisture path into the surrounding core making the problem spread wider and faster. Elevated temperature curing adhesives would probably develop stronger bonds, but given that the core to adhesive interface is so weak, the skins would probably pop off the core in the surrounding areas.

A further concern is that these disbonds appear to be driven by a combination of weak interfaces and compression loads. The surrounding areas are probably just as weakly bonded and the only reason they have not failed is that they have not seen the compressive loads. Even if these repairs were successful, there is no guarantee that adjacent areas would fail in a similar manner.

I am told that one approved repair lasted less than 30 minutes. The real solution is to fix the manufacturing processes. I'd be happy to contract to help.

Adhesive bonded repairs can be very effective, but the processes for those repairs must also be just as valid as the processes for manufacture should be.

Does anyone have a contact in Agusta who I can deal with?

Regards

Blakmax

K1W1phil

9th Sep 2009, 02:05

Hi Blackmax,

with regard to your excellent letter to this thread, you are correct about many misguided constructors out there, in particular reference to sacrificial peel plies and the potential to produce contamination of the adhesive bond.
Also any form of repair by injection of adhesive is flawed as there is no way that the internal structure can be verified as ready for re bonding, anyone can inject and lead themselves to believe that they have corrected a problem but they are only temporarily masking a potentially serious issue. but as this thread is mainly speculative right now it would be silly to try and pre empt a thorough and detailed defect examination. save that the failure of this component is dramatic and the consequences of this event happening in flight would be tragic.
As per your first posting in this thread it would be interesting to see what repair scheme was applied to the boom following the instance of ground contact.
Rumor only, I hear this aircraft type is prone to suffering from exhaust efflux impingement on the boom and coupled with the high ambient temps experienced in the middle east could this be causing disbonding? maybe if it is, a simple mod as per the 212/412 to put a bend in the exhaust duct could help.

Phil

spinwing

9th Sep 2009, 02:41

Mmmm ...

Phil read through the whole thread again .... all has been covered with regard
heat & exhaust re-circulation etc etc ...

:ok:

PS also the exhaust ducts tend to push the hot gas out to the sides on the 139 rather than directly out the back as on the 212/412 ....

9Aplus

9th Sep 2009, 12:31

Does anyone have a contact in Agusta who I can deal with?

Regards

Blakmax

Contact DONE to Somma Lombardo

turboshaft

9th Sep 2009, 14:24

This episode puts an interesting light on the announcement made a couple of years ago by EC - surely the home of 'plastic' helicopters - that the new EC175 would use an aluminum structure (ostensibly for ease of 'on deck' repair) rather than a composite one.

ARRAKIS

9th Sep 2009, 17:35

A question came to my mind.
How would behave, a composite NH90 airframe in a nice, dusty place... if you have a few bullet or rpg-7 holes to patch.... :{

Arrakis

blakmax

9th Sep 2009, 23:51

Guys

Several of the last few postings have suggested that repair of composites is too difficult and metallic structures may be easier to maintain. In some ways that is correct, but what we have is more of an education problem than engineering or materials problems. There are some basic rules which become obvious if you understand the basic mechanisms involved in adhesive bonding and importantly the mechanisms involved in bond failure. There are a few production traps, which if addressed correctly should eliminate many of the problems we see in service.

With regard to field level repair, many see adhesive bonding as too hard. The main problem is contamination, including from humidity. We have addressed this by having a portable airconditioner and dehumidifier which we pump into a roughly constructed tent. The positive pressure from blowing air in keeps the contamination out, and the humidity and temperature are controlled. It does not take too long to cofigure such a system once you have spent the effort to get the basic equipment in place. We have used that method for repairs installed in Northern Australia during the wet season.

The real problem is that owners don't want to invest in the equipment and developing the skill levels necessary. They just want to do what they have been doing for years and not to step out of their comfort zones. I don't blame them but they effectively turn their backs on the most effective and efficient method for repairing structures. I have data which shows that the best you can ever achieve for a mechanically fastened repair in metallic structure (assuming that the dmaage is away from edges etc.) is a restoration of about 65% of original strength. For composite structure the efficiency is down to about 45%. With a properly designed and implemented bonded repair, the strength restoration in thin materials is 100%.

With regard to battle damage, most manufacturers use mechanical repairs, but even these have traps such as larger edge distances and fastener separation distances than for metals.

The message is to get the production processes correct in the first place, improve training and then acquire the necessary equipment. Don't dumb down the repair methods.

Regards

blakmax

blakmax

10th Sep 2009, 01:22

OK, people. I have had some photographs of skin sections cut from a typical disbond in the tail boom sent to me from a reliable source. I am assured they are from an AW139 and from the area which typically disbonds in the tail boom.

In the samples, there are regions just at the edge of the disbond where the core has been in firm contact with the adhesive and has formed what visually appears to be an acceptable bond. To enable core to transfer shear loads, it is essential that the adhesive forms fillets to the core. As the adhesive melts and flows, it wets the surface of the core and by capiliary action flows down the sides for a small distance, thus forming the fillets. These fillets are critical to the performance of bonded sandwich structure, and in the example examined the core had fractured during removal of the sample indicating that these particular cells were well bonded.

However, over the remainder of the surface of the samples I have seen in the photographs there is evidence that the contact between the core and the adhesive has not been sufficient to form the fillets. It appears from these photographs that the capiliary action has only resulted in adhesive contacting just the very end of the core, forming a small bead and not flowing down the sides to form the fillets. The disbonds being investigated when these samples were taken appear to have been due to separation of the beads of adhesive.

To explain the difference, imagine you are to weld one piece of material to another to form a T joint. In the first case, press the sheets together one perpendicular to the other and weld both sides, forming fillets. In the second case, hold the second sheet perpendicular to the other but leave a gap so they are about 5mm apart. Now weld just enough to fill that gap. Clearly the sample which formed the fillets will be much stronger.

There is also extensive micro-voiding in the adhesive layer itself due to the presence of small bubbles. While some micro-voiding is tolerable, the extent in the samples I have seen would not be consistent with what would typically be observed in well bonded structure.

One possible mechanism for disbonding is that the absence of the fillets has led to the loads being transferred by the small bead of adhesive at the end of the cells and in the presence of compression loads which place that bond in out of plane tension, combined with or acting independently of the internal pressure associated with heating, the outcome is that the small amount of adhesive simply can not sustain the loads.

Suffice it to say that the lack of fillet bonds would have meant that this area would not have been as strong as it could have been had the fillets been formed correctly and it would therefore be debatable if the structure could have sustained the loads the design required.

Now, could Agusta have found this with NDI? Probably not. In defence of Agusta, there would have been sifficient contact between the adhesive and the core (even through the small bead at the end of the cells) to enable sound transmission. Hence ultrasonic or tap inspection would have passed this defect. Is it possible to inspect for this defect in current aircraft? Probably not. NDI will only help AFTER the disbond has occurred.

Such a failure to form the fillets results from a number of production issues, and the presence of micro-voiding adds weight to one of the theories I have advanced. There are several other possible production issues that could result in the same defect. The only way to resolve this is to assess the probability of each potential cause occurring within the current production process.

I can not go much further without actually seeing the failed example and assessing AW's production methods and quality processes. Agusta may be able to sort this out themselves, but I am willing to assist if requested. I would much rather help to get the problem fixed than to help some lawyer get rich after some poor b*stard has died.

Sorry if some of the wording is a bit obtuse, but I have tried to take sox6's advice and not to be too prescriptive without actually seeing the articles and assessing the production process.

Regards

blakmax

9Aplus

10th Sep 2009, 06:52

:ok:for last one, hope they will be smart enough.....

outhouse

10th Sep 2009, 12:40

Having been away from PP for some time I have how an opportunity to catch up on my reading. I found this thread fascinating and the technical stuff by blakmax an educational gem.
Many thanks
outhouse.

Encyclo

10th Sep 2009, 17:27

Good analysis Blakmax,

The issue you bring up in the defect not being detectable through NDI/NDT is the reason OEM actually do DT (destructive testing) on a portion of the parts they build, specifically to identify production/quality issues as you describe.

I would hope AW being the professionnals they are (not being sarcastic:=) would have identified adhesive flow issues after the panels are cured in the autoclave.

amicus

11th Sep 2009, 19:58

Based upon all that I have read in this thread, I would insist upon proof loading all existing articles and new booms and all field repairs until AW can prove conclusively that their in-process works on a repeatable and controlled basis. Even then, I would continue to require 100% proof load their booms for several years before reducing the proof test requirement to an acceptable statistical basis.
We can debate all the nuances of hydration and fillets forever, but safety demands a short term and provable system and that is exactly what 100% proof loading provides. Clearly AW has not had and does not have control of in-process bonding procedures and NDI is not capable of detecting these incipient production defects. I have no issue whatsoever with correcting and improving In-process at AW, but we still have a bunch of highly questionable booms in service and inadequate repairs, so proof loading to limit load is the only short term sane and workable procedure from a safety of flight standpoint. Maybe Blakmax can get his yearned for consulting gig and ultimately fix things at AW, but that will take much time and meanwhile we have existing booms and a lousy cold bond SRM repair schemes from AW. This is a safety issue and needs addressing immediately.

slowrotor

12th Sep 2009, 04:05

Here is another method that anyone can use to inspect for disbonds:

Sometimes in the evening or morning, dew will condense on the aircraft surface and the substructure can be seen in the pattern of the dew. A spar or bulkhead disbond might be seen as an area without dew. The dew clearly shows the exact edge of a sound bond. Or at least a bond that has thermal contact.

I have observed this on my aircraft, a Grob G109 many times. Not sure if it will work on honeycomb, but it sure works for skin to spar bonds.

Just another way to visually inspect when conditions permit and the humidity is just right.

spinwing

12th Sep 2009, 05:49

Mmmmm ....

Unfortunately .... this method will not help those who operate in the "Middle East" ......

:hmm:

Eng AW139

12th Sep 2009, 05:55

Lots of morning dew over here..in the Gulf

spinwing

12th Sep 2009, 06:02

Mmmmm....

Yeah trouble is its usually all over the Heli-decks making them lethal slippery ....

D'ya reckon you'd be able to find a de-bond using that method?

Good luck with that! ;)

Heli-phile

12th Sep 2009, 08:10

Where are AW139's still operating and where have they been "withdrawn from service"??.

Not being involved with this particular type I am not up to speed with what inspections have/are being conducted resulting from the tail boom "re-deployment"

Blakmax and other contributors - great posts, keep up the pragmatic and informative posts

blakmax

12th Sep 2009, 09:17

Sorry to say folks, but in Post # 235 I requested information on the surface preparation method used for bonding the repair to fix the disbonds. I tongue in cheek made the comment "Don't tell me is just a solvent wipe". Sadly, I have been advised that the procedure stipulated was a hand abrade and solvent wipe. Now if you have followed my explanation for surface preparation to involve a step to provide resistance to hydration, then you will realise that a hand abrasion and solvent wipe on an aluminium surface will never produce any resistance to hydration whatsoever. I am certain that these repairs will disbond, probably in the shorter term. If you have such a repair, have a look BF to see if it is lifting.

Hand abrade and solvent degrease may produce adequate short term strength if tested directly after the specimen is made. But if you leave the specimen in a warm, humid environment for a few months, the specimen may not even stay together long enough to be tested.

Now I am not familiar with the EASA regulatory system, but FAR 29.605 with reference to processes states " The methods of fabrication used must produce consistently sound structures." For a repair to restore the certification basis for an aircraft, the repair processes must meet the same certification requirements as the original construction. Is there any evidence anywhere that scuff sand and solvent wipe actually produces a "consistently sound structure"? I can show hundreds of examples where it has not.

The worry is that this was a repair to fix a disbond and it has replaced that disbond with a repair which almost certainly will result in another disbond. The only reason these repairs do not fall off is that they installed fasteners to restrain the patch. The question is, has the repair design provided adequate strength to carry the load through only the fasteners? Next question: Even if the fasteners can carry the loads, is there enough stiffness to prevent crippling failure once the insert disbonds from the skin?

I agree with Amicus. This is a safety of flight issue, and in the short term it is relevant to the AW139 and the focus is on Agusta and something must be done.

However, there are just as many other platforms out there using the same repair methods and built with similar deficiencies in their construction methods. It really comes down to defining what is a "consistently sound structure". The only long term solution I believe is to amend the airworthiness regulations to mandate demonstration that production processes not only meet static strength and fatigue requirements, but also demonstrate long-term bond durability.

I think it is time that the regulators ackowledge that the current regulations do not adequately address a possible (common?) failure mechanism which can result in loss of aircraft. I am aware that the FAA is trying to manage this by amending an Advisory Circular AC-20-107, but what value is advice? Surely a possible (probable?) cause of structural failure requires mandatory regulation, not optional advice? How important do we feel this matter is? If even one regulator or manufacturer would organise a conference or meeting on the subject and cover our out of pocket costs, my company would be prepared to waive professional fees to present at that conference.

Regards

Blakmax