Bell 525 fatal accident July 2016
Extract from one report, "It proceeded to the south, where it flew a pattern at altitudes between 2,000 and 3,000 feet. The last data point is at 1,975 feet at a speed of 199 knots at 11:47 a.m. A witness on the ground reported hearing a loud twin boom sound before seeing the helicopter make an immediate descent. Most of the helicopter debris is highly compact, though some is hundreds of feet away from the primary crash site. Dallas television station KXAS reports a section of the tail boom was located 1,500 feet southeast of the primary crash site."
A witness on the ground reported hearing a loud twin boom sound before seeing the helicopter make an immediate descent
This morning the PR people at Bell announced they were suspending all promotion of the Bell 525 at Farnborough as a matter of 'respect'.
It was also announced that the CEO Mitch Snyder was no longer travelling to the UK for Farnborough for the same reason. Of course there were a number of other important Bell Helicopter programmes that he was there to promote besides the 525.
While perfectly understandable in the circumstances this is the second European media/PR event that Snyder has cancelled within a month. For the future health of the company there must surely be a limit to the number of times the CEO of any company can allow unfortunate circumstances to diminish his marketing role.
It was also announced that the CEO Mitch Snyder was no longer travelling to the UK for Farnborough for the same reason. Of course there were a number of other important Bell Helicopter programmes that he was there to promote besides the 525.
While perfectly understandable in the circumstances this is the second European media/PR event that Snyder has cancelled within a month. For the future health of the company there must surely be a limit to the number of times the CEO of any company can allow unfortunate circumstances to diminish his marketing role.
Considering the importance of the 525 to Bell and how much they have invested in it, his time is probably better spent overseeing their investigation into the accident and protecting the order book by reassuring customers.
What leads you to that guess, Sp 206?
Is that how a Rotor system would react to RBS.....Rotor system dipping to the rear or would the aircraft nose up and roll? How extreme would it be before the Pilot lowered the Collective and moved the Cyclic Aft upon realizing the aircraft had entered RBS?
Would there be any warning signs of the onset of RBS....what would they be?
Would there be any warning signs of the onset of RBS....what would they be?
ADS-B does not report vertical or horizontal airspeed. Instead, ADS-B reports horizontal and vertical velocity relative to the Earth. This velocity is useful for air traffic control functions and ADS-B applications. Airspeed can be provided by other aircraft sensors.
When you look at RBS analytically it is generally with respect to the airspeed over the rotor disc. As you go faster eventually the whole retreating side has negative airspeed. RBS comes before that point as the blades need a higher AOA to preserve symmetry of lift. Extra blades are no guarantee against RBS.
Seems much too early to speculate. I am sure the data will tell its tale. My heart goes out to Bell, my former competitors yet we are brothers all.
Seems much too early to speculate. I am sure the data will tell its tale. My heart goes out to Bell, my former competitors yet we are brothers all.
Last edited by IFMU; 10th Jul 2016 at 23:40. Reason: Last paragraph added
Rotor Stall Behavior and Loads
SAS, the modern multi-bladed rotors that I have experience with at Sikorsky do not exhibit the severe roll left tendency that I and perhaps others of the S-55/H-19 generation were schooled in.
I can only speak to the S-65,S-70, S-76 and S-92 series. If you drive these rotors into stall, the N/rev vibrations go up dramatically accompanied by non-linear ( with load factor ) increases in both the rotating ( push rods e.g. ) and stationary ( MR servo e.g. ) vibratory loads. A flight test team exploring stall limitations will of course be utilizing telemetry ( TM ) and the relevant loads will be a priority data stream for real time monitoring. Typically, the test team will work to a do-not-exceed limit ( DNE ) for these loads. The DNE is established based upon accelerated fatigue load testing of said components in the test lab, and are based on a very short component life, based in turn usually on a three sigma reduction from the component failure data. It is common for test teams to cycle count fatigue damage on these parts and they are replaced, discarded as required.
As you can see, this approach and taking into account that the published report is that the subject 525 was using TM, and against the personal experience just covered, makes it seem unlikely the cause was as simple as an entry into a rotor stall condition. But I hasten to add that I don't know anything about the 525 rotor, thus I'm out on a limb even going that far. It would be fair to add that during the 39 years at SA, there were a few rotor stability issues that cropped up, each with a different cause and a different solution, and I'd guess that other OEM's are no stranger to those occurrences.
Henra had it right: let the pro's at Bell who know their rotor and new machine do what they alone are equipped to do.
I can only speak to the S-65,S-70, S-76 and S-92 series. If you drive these rotors into stall, the N/rev vibrations go up dramatically accompanied by non-linear ( with load factor ) increases in both the rotating ( push rods e.g. ) and stationary ( MR servo e.g. ) vibratory loads. A flight test team exploring stall limitations will of course be utilizing telemetry ( TM ) and the relevant loads will be a priority data stream for real time monitoring. Typically, the test team will work to a do-not-exceed limit ( DNE ) for these loads. The DNE is established based upon accelerated fatigue load testing of said components in the test lab, and are based on a very short component life, based in turn usually on a three sigma reduction from the component failure data. It is common for test teams to cycle count fatigue damage on these parts and they are replaced, discarded as required.
As you can see, this approach and taking into account that the published report is that the subject 525 was using TM, and against the personal experience just covered, makes it seem unlikely the cause was as simple as an entry into a rotor stall condition. But I hasten to add that I don't know anything about the 525 rotor, thus I'm out on a limb even going that far. It would be fair to add that during the 39 years at SA, there were a few rotor stability issues that cropped up, each with a different cause and a different solution, and I'd guess that other OEM's are no stranger to those occurrences.
Henra had it right: let the pro's at Bell who know their rotor and new machine do what they alone are equipped to do.
Last edited by JohnDixson; 11th Jul 2016 at 00:12. Reason: Word left out/grammar correction
Brother Dixson as usual makes a very difficult explanation sound childishly simple!
The only helicopter I know for sure I had RBS happen was in a CH-47A due to a slightly over gross external load......and it occurred in the 46-48 knot range. Being Tandem Rotor....it was a very unusual result.
We had managed to trundle off with just over 16,000 pounds of 105 MM Howitzer Ammo due to a great amount of indiscretion. We realized our mistake when the FE counted the Projectiles.....and we tweaked to the fact we had grabbed a CH-54 Sky Crane Sling Load by mistake.
The only helicopter I know for sure I had RBS happen was in a CH-47A due to a slightly over gross external load......and it occurred in the 46-48 knot range. Being Tandem Rotor....it was a very unusual result.
We had managed to trundle off with just over 16,000 pounds of 105 MM Howitzer Ammo due to a great amount of indiscretion. We realized our mistake when the FE counted the Projectiles.....and we tweaked to the fact we had grabbed a CH-54 Sky Crane Sling Load by mistake.
SAS, just out of flight school, having wangled an assignment to the Test Board at Ft Rucker, I got to be instructed in the CH-47 by Boeing Test Pilot Jim Campbell*. Asking him re stall ( I had had two very realistic H-19 events post flight school ) in the Chinook, Jim said he had flown some of that work, and due to the lift split between the fwd and aft rotor, the aft operating with higher thrust, thus higher angle of attack, the aft rotor would stall first. He said it was therefore a sort of self-correcting situation: the aft rotor would stall first, the ship would go nose up, decelerating the aircraft and installing the rear rotor. That sound right?
* Heard later that unfortunately, Jim left Boeing for Air America and was KIA in SE Asia.
Your " CH-54 Pretend " flight might provoke some of the uninitiated to carp about failure to fill out a 365F Weight and Balance Form etc etc , and I thought you might comment re the very real issues of flying cargo in an active combat situation.
* Heard later that unfortunately, Jim left Boeing for Air America and was KIA in SE Asia.
Your " CH-54 Pretend " flight might provoke some of the uninitiated to carp about failure to fill out a 365F Weight and Balance Form etc etc , and I thought you might comment re the very real issues of flying cargo in an active combat situation.
Last edited by JohnDixson; 11th Jul 2016 at 03:07. Reason: Added thought
No sure about that explanation. Typically you would say that for the same lift a rotor with more blades can be smaller in diameter, and can turn faster. As the retreating blade stalls when the airspeed of the a/c approaches the speed of the rotor blade relative to the airframe, that point it reached at higher airspeeds in a rotor that spins faster.
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I had an RBS experience in a Sea King, high and fast and we got a very rapid onset 1:1 lateral oscillation.
It is 5 blades, fully articulated.
RIP to the lads, and condolences to the families.
It is 5 blades, fully articulated.
RIP to the lads, and condolences to the families.
Viper 7
Viper, that sounds like pitch-lag instability, not stall.
Part 2: I looked in vain online for an old H-3 flight manual. There used to be a write-up on pitch-lag instability in the USN military manual.
Anyway, re H-3 pitch-lag: The H-3 rotor has some alpha-1 coupling in the geometry. In the H-3, as the blade lag angle increases ( which occurs with increased power ), the pitch angle decreases a bit. Hardly noticeable, but coming back into a hover on the 61, the pilot needs to nudge the collective up a bit more when he applies collective to come to the hover. Really a non-issue.
However, if a blade damper has a problem ( e.g. sticky relief valve ) then at higher speeds, and having nothing to do with stall*, the rotor can excite a pitch to lag angle instability. Pilots unfamiliar with the excitation may well call it a 1:1 or 1/rev, but in fact it is 2/3 per rev. Fix is to look at the dampers and check their timing.
* USN H-3 had a cruise guide indicator. The main rotor servo ( typically highest loads on the newer machines have highest loads on the aft longitudinal servo, but I honestly don't recall which servo on the H-3 had the highest loads ) had an LVDT ( linear,variable,differential transducer ) which measures the loads on that servo and feeds an indicator with range markings. Like the same system on the S-65 series, anything over 30% in indicative of increasing degrees of stall.
On the H-3, with a damper problem, one can see pitch-lag when the cruise guide will be barely indicating anything. It will be at higher power, thus faster speed, but its not a simple 1/rev. An easy way to evaluate the difference is that with a simple 1/rev, ( and assuming it is large enough to get one's attention )there is usually one blade out of track and quite visible from the cockpit. With pitch-lag, the main rotor feels and appears to wobble at a slightly slower frequency, and one can see that.
Rotor heads subsequent to the S-61 have a very flat alpha-1 geometry, that is, almost zero blade pitch change as the lag angle changes, and as a result have been absent this particular instability.
Part 2: I looked in vain online for an old H-3 flight manual. There used to be a write-up on pitch-lag instability in the USN military manual.
Anyway, re H-3 pitch-lag: The H-3 rotor has some alpha-1 coupling in the geometry. In the H-3, as the blade lag angle increases ( which occurs with increased power ), the pitch angle decreases a bit. Hardly noticeable, but coming back into a hover on the 61, the pilot needs to nudge the collective up a bit more when he applies collective to come to the hover. Really a non-issue.
However, if a blade damper has a problem ( e.g. sticky relief valve ) then at higher speeds, and having nothing to do with stall*, the rotor can excite a pitch to lag angle instability. Pilots unfamiliar with the excitation may well call it a 1:1 or 1/rev, but in fact it is 2/3 per rev. Fix is to look at the dampers and check their timing.
* USN H-3 had a cruise guide indicator. The main rotor servo ( typically highest loads on the newer machines have highest loads on the aft longitudinal servo, but I honestly don't recall which servo on the H-3 had the highest loads ) had an LVDT ( linear,variable,differential transducer ) which measures the loads on that servo and feeds an indicator with range markings. Like the same system on the S-65 series, anything over 30% in indicative of increasing degrees of stall.
On the H-3, with a damper problem, one can see pitch-lag when the cruise guide will be barely indicating anything. It will be at higher power, thus faster speed, but its not a simple 1/rev. An easy way to evaluate the difference is that with a simple 1/rev, ( and assuming it is large enough to get one's attention )there is usually one blade out of track and quite visible from the cockpit. With pitch-lag, the main rotor feels and appears to wobble at a slightly slower frequency, and one can see that.
Rotor heads subsequent to the S-61 have a very flat alpha-1 geometry, that is, almost zero blade pitch change as the lag angle changes, and as a result have been absent this particular instability.
Last edited by JohnDixson; 11th Jul 2016 at 13:56. Reason: Added information. Forgot one thing.
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The flight test of a new aircraft involves a great deal of instrumentation and real-time monitoring by excellent pilots and engineers. Mysteries to field pilots, like Blade Stall, are easily measured and diagnosed as they start to occur in test aircraft, and watching these conditions develop is the craft of a modern test team. On all the test aircraft I flew, we had cockpit indications of the degree of stall, giving plenty of warning prior to any large stall event.
It is highly unlikely that the cause will be determined by a group of guys throwing darts at this forum, much more likely that a tragic set of individually obscure circumstances produce an unlikely chain of events that led to the mishap.
The craft of development helicopter flying is similar to clearing a mine field, where the object is to find the hazards before they trigger massive problems, using tools that find the problems before they become catastrophic. It takes about 1600 flight test hours to certify a helicopter type, and every one of those hours has a finite probability of catastrophe, even the mundane maintenance test flights and avionics check flights.
I know the Bell team well, they are pros, as good as anyone at this game. This has probably shaken them up because we all believe with confidence that our systems work well, and that data, careful test procedures and training will win every time. None the less, the nature of the unknown is to be unknown, and every now and then one of those mines goes off.
This accident is probably not an measure of the professionalism of the team, or of the innate airworthiness of the aircraft, it is a measure of the uncertainty of exploration, and the cost of being surprised.
A quote that helps explain it comes from racing driver Jackie Stewart, "Was I ever scared while driving? If I wasn't scared, I wasn't driving fast enough."
It is highly unlikely that the cause will be determined by a group of guys throwing darts at this forum, much more likely that a tragic set of individually obscure circumstances produce an unlikely chain of events that led to the mishap.
The craft of development helicopter flying is similar to clearing a mine field, where the object is to find the hazards before they trigger massive problems, using tools that find the problems before they become catastrophic. It takes about 1600 flight test hours to certify a helicopter type, and every one of those hours has a finite probability of catastrophe, even the mundane maintenance test flights and avionics check flights.
I know the Bell team well, they are pros, as good as anyone at this game. This has probably shaken them up because we all believe with confidence that our systems work well, and that data, careful test procedures and training will win every time. None the less, the nature of the unknown is to be unknown, and every now and then one of those mines goes off.
This accident is probably not an measure of the professionalism of the team, or of the innate airworthiness of the aircraft, it is a measure of the uncertainty of exploration, and the cost of being surprised.
A quote that helps explain it comes from racing driver Jackie Stewart, "Was I ever scared while driving? If I wasn't scared, I wasn't driving fast enough."