Thanks for the input.
What is the approximate full stroke of the actuator relative to the washer thickness? Would the missing washer offset on null and full stroke positions be subtle/adjustable at some other point or clearly way out? In any case it is hard to fathom leaving the factory that way or with missing lock means.

dont make the assumption, or imply that i did, that it fell out into the grass. more than likely it stayed within the cowlings, at least until the crash anyways

Here is a question.

Can someone explain the process / thinking around the AD? In so much that the commentary for the reason of the AD says:-

What I don't understand is the aircraft crashes almost 2 weeks ago. If the AD is truly a precautionary element what prompts the thinking of this 2 weeks later? Surely either its a precaution thinking out loud type AD, in which case why not flag earlier? Or they are concerned about something they have seen more recently, in which case why not just say?? It just seems bizarre.

Yeah that would be fair but what evidence? The video footage was available almost instantly and so now evidence surely means physical evidence from the accident aircraft or of prior events that have been caught prior to an accident with other in service aircraft. It can't be the former if the narrative around the AD is faithful and if its the latter does that take a week to get out with the commitment of intelligent minds?

It would be incredible if this isn't off the back of something seen or suspected and so why it can't be said more plainly - or indeed co-ordinated via the AAIB goodness knows... Hey Ho.

True enough, the reason I was wondering about the washer is that it is the only thing that might not be obvious if misassembled since a washer is often seen under a nut. Of course that does suggest that if it was suspect it would have been mentioned in the AD.

Most likely is that they suspect a fault in the feedback link but may not know exactly what at this stage.
Hard to fathom how pin and lockwire could be missed on assembly so might be a failure.
Here is one scenario: If the lockwire was incorrectly installed could it eventually wear into the locking pin, possibly on the head side?
.

It is a slow news week....

I think TR blades tend towards zero pitch when uncontrolled. Gazelle definitely did in manual. Lynx as well, albeit with a powerful spring bias unit? Sea King? Is it generic or type specific?

CAA PAPER 2003/1
Helicopter Tail Rotor Failures is well worth a read. Here are some extracts.

There are two major types of TRF:
a) A TR drive failure (TRDF) is a failure within the TR drive system with consequent
(usually total) loss of TR thrust. Example causes are internal fatigue or external
impact resulting in a broken drive shaft.
b) A TR control failure (TRCF) is a failure within the TR control system such that
normal pilot control of TR thrust has been partially or totally lost. Example causes
are internal wear or external impact resulting in a severed control cable. The
resultant TR applied pitch, or power, could be free to fluctuate, or may be fixed
anywhere between high pitch (HP) or low pitch (LP) setting, including that of the
current trim pitch (TP).
Both of these TRFs are time critical emergencies. The pilot has to identify and
diagnose the TRF type and react with the correct control strategy within a few
seconds (or less), to prevent the aircraft departing into an uncontrollable flight state.
Even if the pilot recovers from the initial transients, yaw (pedal) control will have been
lost and the ability to manoeuvre safely and carry out a safe landing will have been
significantly degraded. The TR and its drive and control systems are clearly flight
critical components and should be designed so that their probability of failure is
‘extremely remote’. The airworthiness design requirements for UK military and civil
aircraft define ‘extremely remote’ as being less than 10-6 [1] and between 10-7 and
10-9 [2,3] per flight hour respectively.

Recovery from the failure transient
For TRDFs, and TRCFs where the post-failure pitch angle of the TR blades is different
from the pre-failure trim position, the immediate effect is a yaw response. That is, (for
anticlockwise main rotors), nose to starboard following a TRDF or LP TRCF, and nose
to port following a HP TRCF. The level of initial yaw acceleration will depend on the
nature of the failure, and the level of yaw rate and attitude build-up will depend on the
forward speed. In hover, an unchecked TRDF will result in the yawing moment from
the main rotor torque reaction spinning the fuselage at rates in excess of 100° sec-1,
perhaps even as high as 150-200° sec-1. Typically, the higher the forward speed, the
lower the yaw rate and attitude excursion as any natural directional stability of the
aircraft will tend to reduce the severity of the motion. However, this is only true up to
some value of sideslip, beyond which it is possible that directional stability can
reverse, resulting in increased yaw rate and attitude excursions. Evidence from the
Lynx TRF AFS trial [5] suggests that the ability of the pilot to successfully manage a
forward flight failure is strongly related to the extent of the initial yaw/sideslip
transient. If this exceeds 90°, then the pilot is unlikely to be able to recover, as the
flight control problem is exacerbated by disorientation; if the yaw rate reduces to zero
below about 30° yaw angle, then the pilot has a much greater chance of recovering
from the failure. Accompanying the yaw excursions will be pitch and roll motion,
which can further increase the risk of disorientation. An additional effect of any roll
attitude transient is an increase in the main rotor disc angle of incidence, leading to
an increased risk of the rotor over-speeding as the pilot reduces main rotor collective
to contain the effects of the failure. The extent of the attitude excursions depends on
the aerodynamic design characteristics of the fuselage and vertical stabiliser, the
resulting directional stability, the type of attitude stabilisation present in the flight
control system and the pilot’s control actions

Chronus, #3, loss of tail rotor blade ,followed milliseconds later by loss of other blade(s),and probably the gearbox.. The C of G of the a/c will move to somewhere beyond the nose of the a/c ....

How can CG be fwd of nose?

As a former helicopter pilot I experienced three total engine failures in single engine helicopters and three lucky autorotations that saved the day for my pax, me and the company. Most of my helicopter time though is logged in twin engine helicopters, one of which also gave me an engine failure, but that was in level flight and therefore nothing of interest..

With Occam's razor in mind I got to my very own simplified conclusion that this accident started with a tail rotor failure that turned out catastrophic due to no visual recovery in the blurry video. The helicopter’s counter rotation and the engine sound made me think the helicopter descended with high power and high collective all the way to the ground.. The shape of the tail rotor blades on the downed helicopter made me also think that the helicopter hit the ground with the TR not spinning at all due to a TR drive failure… Why wasn’t the pilots able to lower the collective?

Slap,

That was some British understatement being used.

When you lose the Tail Rotor Blades, Gearbox, and associated items from the very rear of the helicopter....the CG shift is forward...and very significant....causing a strong pitching movement that might exceed the ability of the flight controls to compensate for and thus cause the loss of the aircraft.

It could be that during the preliminary investigation something was found that could or could't be the main cause so Leonardo as a precaution released the SB.

The investigation is probably far from over and as far as I know no one would release informations until all data has been analysed.

My guess would be because it was a time critical event.
Factors are : Height above ground coupled with the likelihood of occurence at critical moment of transit to forward flight, size of helicopter, twin engines, high rotation speed following TRFand location of site.
Any chance of succesful recovery would have entailed almost instant recognition of TRF, immediate power reduction and simultaneous pitch down for immediate descent and autorotation, which would involve a high ROD and require favourable terrain below. Might be survivable at low height but not so at the sort of height involved in this instance.
I would readily admit I know little about rotary wing, but I understand that they are far less forgiving than their fixed wing sisters. One thing in common must be that with the vertical stab gone on a fixed wing it also spells curtain time, but as in the cases of the JAL123 747 which managed slightly better than AA 587 they seem to keep going for a little longer before the inevitable. With rotary wing the equivalent of no TR means in no time at all the machine assumes the characteristics of a dandelion strung to a brick.
I really don`t think even the most accomplished pilot could have pulled it off. It would have been a far greater miracle than the Hudson one if they had, would be my humble view.

Chronus: Is this because of loss of yaw stability or the shift in CG?

Jag,you will lose your stability in yaw,coupled with the C OF G change; you might be able to crowd pax into the back to restore/mitigate the CG PROBLEM.The a/c if disturbed in yaw or roll will probably start Dutch -rolling which will require a lot of handling skills,especially during any configuration changes,and approach/landing....or may not.. A B52 lost it`s fin/rudder and landed safely-see u-tube..

Slap,in my case of t/r and gearbox departure ,the change of CG was about 3" forward of the Fwd limit,`past my nose, anyway`!! The stick has reached the back-stop ,and the a/c has changed direction by about 50-60*;entering autorotation and chopped the engine leads to a further pitch down ,but there is a `pimple` of a hilltop with cleared scrub that we are pointing at.
To flare requires another quick fwd stick and back again with lever,as it`s also uphill,but at zero groundspeed.Landed upright,but burst a main tyre on a treestump.. ...There is a pic on `Rotorheads around the World-views from cockpit(not video) ,page 15....

That is not always the case SAS, I agree on the forward shift, however actual CG position at time of accident is a big part of the equation.
You can PM me and we can talk the details.

Why look for a TR failure ?
 

Statistically there are many more accidents due to unanticipated yaw -- so poorly named as LTE -- than due to TR failures. Looking at the initial yaw, there is no abrupt acceleration. If it is a TR failure, I guess you have to look for a very progressive failure...
AAIB investigations will say it.

.................................

You will not have "CG move in front of the nose" even you lose the TR while keeping the H stab. Thats BS unless helicopter had been loaded nose heavy anyways. Most of the weight are lining up with the MGB and the moment of inertial will hard to overcome. FM says it all.

MWR ... If you are in any way correct, this implies a really poor original design... a prime goal in mechanical components design just as important as stessing and robustness is clarity of assembly whilst minimising potential for mis-assembly...
e.g. all bolts in a locality doing much the same job being of the same length, usually arranged by say stepping a casting thickness .. EVEN IF IT COSTS WEIGHT..

Nailed it!

He has not really "Nailed" it. the "Unless" statement is still hogwash.

To have the CG in front of the nose of the helicopter, you would need a MRB sticking over nose, the portion beyond the nose, weighing more than the entire helicopter behind it.

To keep this sensible and simple, The Helicopter CG should be directly under the Main Rotor Mast centroid both Longitudinally and Laterally. That's the optimum ideal position giving 100% MR flight control in all directions.

However, practically speaking, there has to be some loading of the helicopter which would cause the CG to move around the rotor centroid. It cannot move that much. Even in a large helicopter like an EC225 it can only move 0.5m longitudinally.

There are limits to how far the CG can move away from the rotor centroid before effective flight control is compromised.

If the TRGBX and its blades depart the airframe, the CG will make a significant move forward. It is highly likely in ALL helicopters that this will result in a significant compromise of the forward CG limit and effective aft cyclic to counter the forward movement of the CG will reach the aft stop before the pitch moment forward can be cancelled. The results.....well not pretty.

The Datum, as someone has already posted, is simply a point in space to facilitate all calculations to be in the positive range. Usually it is quite a way out in front of the helicopter nose to facilitate the fitment of long PITOT probes during certification. For simplicity, it remains there. Therefore, the limits for longitudinal CG are expressed in positive numbers and to keep the theme at the EC225 4.4m to 4.9m. This actually means, behind the datum. Not in front of the rotor centroid.

I know 99% of Rotor heads know this but to clear up the incessant garbage postings by the odd individual who seems confused....enjoy!

The only occasion I know off where the main door departed into and then removed the tail rotor assembly resulted to a massive pitch down which the pilot tried to correct with aft cyclic. The caused the main rotor, in sympathy with the pitch up of the tail boom to slice off the boom just aft of the mounting point.

The ensuing pandemonium resulted in the aircraft descending almost vertical and killing everybody on board.

Bristow lost a 412 in Nigeria and the evidence showed a Cabin Door had come into contact with the Tail Rotor.

The flight was at night in IMC weather.

The aircraft and occupants were not recovered as I recall.....just some odd bits and pieces.

A ROV was lost on the first dive and the search was called off afterwards.

Bristow lost a 212 in the Sumatran jungle when the T/R departed in the cruise with half of the 90 box. The T/R was not found within a few hundred yards of the crash site. It later found by some Indonesians who were paid a lot of rupiahs, the distance from the locus could not be determined. The rotor was examined by the AAIB and it was determined that it had been hit by an oil cooler duct. The aircraft had caught fire and the port side was burnt out but the starboard panel was still attached. It was assumed that a faulty catch on the panel had failed and the panel had flown over the tail boom and hit the rotor.

For; Blade Slap,totti,M3- perhaps I should have put the phrase about the CG` beyond the nose`` in italics/commas,or a couple of smilies.It was meant to imply that the loss of a tail-rotor and gearbox will have a very significant FWD C o G change,irrespective of where it was originally,and if it was FWD AT THE TIME you will get a significant NOSE DOWN pitch.
In my case ,I had 2 engineers at the front of the cabin,looking at the rear of the engine and reduction g.box checking for oil leaks,as we had just done an engine out,change a component /pipe on the back of the gearbox,, engine back in,and RTB.,so my CoG was well Fwd anyway.
I might suggest if you are pilots or gingerbeers that you go ,find the appropriate tech manual which shows all the weights for tailrotor ,g/box,etc and work out the resultant change of C o G ,if they should depart,......
SAS,DB, thanks....

Not all aircraft are affected the same, it seems that the 212/412, perhaps the H-1 series and civilian counterparts suffer the most.
I am aware of more than one 407 TRGB and a portion of tailboom departures back in '98/2000 where a successful autorotation was made at least once to the water (PHI GoM 1998 or '99) with a full load of passengers, the pilot reported no excessive pitch down moment.
As far as consulting the appropriate tech manuals, I have access to a couple of types (which I cannot disclose) and we ran calculations confirming that the loss of the TRGB and the TR blades does not move the CG forward to a catastrophic unrecoverable value.
HTH

Totti, OK,I`ll` add a couple of other bits as your calculator is warmed up;
IF the helo now enters a full autorotation,how much aft stick movement is produced at a variety of airspeeds,to keep a level attitude;
now do it all at a fwd C o G...
now consider the fact that the aircraft has yawed(depending on airspeed and fin size)which may also have a `pitch` effect..
now consider where the stick/disc relationship is to tail- boom.....and your stick margins vs control stops

Answers on a postcard,,,...or sheets of A4.....

Not rocket science to suggest that the longer the tailboom, the further from the C of G is the TRGB and the more marked effect its loss will have on longitudinal C of G.

I don't need to calculate anything,read my post again.
Crab, you are absolutely correct.

Totally agree, I sincerely hope that my conjecture on misplaced washer is way off. I mentioned it only because at least as presented in the AD it is hard to see how both the split pin and lockwire could be missing without being noticed and was pondering what else could be mi-assembled.

GeordiMike also responded that lack of washer would likely affect the centering of the control system enough that it would at least be noticed if not totally unserviceable.
(My thoughts based on his feedback)

Again most likely they suspect something in that mechanism but do not know (or cannot say with certainty) exactly what so the goal is to have it looked at for anything suspicious.

did you magnify the photo?
If you haven't, please do. I see the arrangement in the photo as indicated in the drawing.

Quote:
I saw that the photo matches the drawing as expected since the photo is of a correct assembly. My question was about possibility of an incorrect assembly with the washer on the wrong side, under the nut, and if that was even possible without causing an obvious problem. If it is possible then the hardened rod shoulder could stress the 'hinge bracket element (90)' and cause a failure,especially if the shoulder was significantly smaller than the outside diameter of the washer.

Without detailed drawings and system knowledge whether misassembly is even possible is an open question though feed back from GeordiMike suggests it probably would be obvious due to offsets in the servo loop.
Also if washer misplacement (or missing) is possible seems that it would have been an explicit check in the AD which does not mention the washer.
Then again something did go wrong, as others have said will be clear in final report.

thats why I asked, I figured that as well, if the washer was suspect for any reason it would be part of the AD.

Incidentally for simple random information totally unrelated to this case, I was disassembling some AS350 control rods today, and while cutting the lockwire, the nut started turning. the lockwasher design didn't stop the nut as it was just a place for the lockwire to terminate.
The "double locking" in this case did its job although the proper torque was not on the nut.
However, nothing further would have happened because the lockwire, torgue and locktabs, on the other end would have all had to fail too before bad things happened.