He stepped on the Rudder and redefined Va
It occurs to me that an important detail has been left out of the debate:
The rudder inputs were so violent that the engines ripped off the pylons
.
Again, still not in the habit of defending Brand A, but Boeing uses sheer pins to allow the engines to separate the aircraft in certain extreme circumstances - e.g. rotor seize or a wheels up landing - to avoid more serious damage. For example, the stresses from a high-power rotor seize can potentially fail the wing structure - better to let go of the engine.
If the rudder oscillations are so severe that the engine struts fail, does it really matter much if the vertical tail stays intact? It occurs to me that if the engines have departed the aircraft, you're pretty much guaranteed to have a bad day
The rudder inputs were so violent that the engines ripped off the pylons
. Again, still not in the habit of defending Brand A, but Boeing uses sheer pins to allow the engines to separate the aircraft in certain extreme circumstances - e.g. rotor seize or a wheels up landing - to avoid more serious damage. For example, the stresses from a high-power rotor seize can potentially fail the wing structure - better to let go of the engine.
If the rudder oscillations are so severe that the engine struts fail, does it really matter much if the vertical tail stays intact? It occurs to me that if the engines have departed the aircraft, you're pretty much guaranteed to have a bad day
It occurs to me that an important detail has been left out of the debate:
The rudder inputs were so violent that the engines ripped off the pylons
The rudder inputs were so violent that the engines ripped off the pylons
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Oh gee wiz...anyone in here see acro planes, student trainers, even tprops and bizjets during training, people are standing on the rudders, back and forth all day long, year after year for decades...not one tail has fallen off..
Have a read about the square cube law.
In mathematical terms, the law states that when an object increases in size, its weight multiplies faster than the strength of the structure that supports it.
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Boeing uses sheer pins to allow the engines to separate the aircraft in certain extreme circumstances - e.g. rotor seize or a wheels up landing - to avoid more serious damage. For example, the stresses from a high-power rotor seize can potentially fail the wing structure - better to let go of the engine.
After the El Al crash the 747 fuse pins were changed to steel and additional attachments were added to all 747s to prevent separation. I believe this is the case with all Boeing now.
I have one of the old fuse pins on my desk as a paper weight.
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AirRabbit
However, there are some who will refuse to spend the time to read AND understand, believing that simply reading the material or relying on their own knowledge and experience will be sufficient. It’s those persons I would prefer to NOT fly with, thank you.
However, there are some who will refuse to spend the time to read AND understand, believing that simply reading the material or relying on their own knowledge and experience will be sufficient. It’s those persons I would prefer to NOT fly with, thank you.
Teldorserious
Oh gee wiz...anyone in here see acro planes, student trainers, even tprops and bizjets during training, people are standing on the rudders, back and forth all day long, year after year for decades...not one tail has fallen off..
What gets me is that the really funny kinds of accidents that defy logic seem to happen in Airbus's...
The only double engine failure ever due to birds?
The only crash where three pilots lose complete SA in level flight completely lost, right until the they impact the water?
The only tail coming off on departure at slow speeds?
Didn't some Afgani AB go down because the FO's seat slid back? Where was the captain?
Oh gee wiz...anyone in here see acro planes, student trainers, even tprops and bizjets during training, people are standing on the rudders, back and forth all day long, year after year for decades...not one tail has fallen off..
What gets me is that the really funny kinds of accidents that defy logic seem to happen in Airbus's...
The only double engine failure ever due to birds?
The only crash where three pilots lose complete SA in level flight completely lost, right until the they impact the water?
The only tail coming off on departure at slow speeds?
Didn't some Afgani AB go down because the FO's seat slid back? Where was the captain?
I rest my case.
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SMOC - Using your logic, that airliners are weaker as they are bigger, as opposed to the way aircraft are tested to withstand x loads at y speeds...then I will run with your premise and see how it Socratically follows out..
Two planes take off into turbulance..one is a biz jet, another is an airliner.
Both are flown 1kt below Va, both are getting slammed around, both crews are doing what they can to keep the rubber side down.
In the ensuing departure, airline test pilot A, steps on the right rudder as hard as he can, all the way to the stop, then the left, then the right..the tail comes off, they don parachutes and bail out.
In the ensuing departure, bizjet test pilot B, gives a Jean Claude Van Damme kick to the right rudder, then then left, then right. Flight procedes normaly.
According to your view of how aircraft are certified the bizjet is an inherently more robust aircraft.
Two planes take off into turbulance..one is a biz jet, another is an airliner.
Both are flown 1kt below Va, both are getting slammed around, both crews are doing what they can to keep the rubber side down.
In the ensuing departure, airline test pilot A, steps on the right rudder as hard as he can, all the way to the stop, then the left, then the right..the tail comes off, they don parachutes and bail out.
In the ensuing departure, bizjet test pilot B, gives a Jean Claude Van Damme kick to the right rudder, then then left, then right. Flight procedes normaly.
According to your view of how aircraft are certified the bizjet is an inherently more robust aircraft.
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Blantoon, don't come in here with 4 posts and start giving the riot act. I am all ears if you have something to offer up.
If the tails of airliners come off under Va, then technically Va doesn't exist, or it needs to be lowered, or recertified with a placard that states.
'Pilots are to only use the rudder one way per episode to avoid catastrophic results'
If the tails of airliners come off under Va, then technically Va doesn't exist, or it needs to be lowered, or recertified with a placard that states.
'Pilots are to only use the rudder one way per episode to avoid catastrophic results'
The structural certification standard never envisioned rapid reversals of the flight controls. Va does not imply a strength criteria, it merely is the point at which the wings stalls at the G limit, thus relieving the wing loading. No plane is tested for rapid reversals; similarily asymmetric G load limits are published for FA 25 aircraft, but the exact.
GF
GF
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Originally Posted by AirRabbit
During the last 10 days of 2006 through the first 8 or 10 days of 2007,
sharing movements of the pedals used with much superstition against endemic dutch roll or any not understood oscillation - PIO or other oscillation. Isn't?
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This is directed to anyone who may be interested and specifically to Mr. Teldorserious.
You know … throughout my participation on this forum, while it may have only been 8 years or so, I believe I have always tried to maintain some sense of professionalism with the others who participate here – even when they were off into some of their own beliefs – as limited as they were (and recognized as such by many other participants). So, if you would not object, I would like to suggest that you might consider reading the section of the rules that govern the certification of various kinds of airplanes.
In the US, those rules are written, published, and enforced by the FAA – in other countries those responsibilities fall to other organizations. As for the US, the section of the rules that govern the certification of transport category airplanes (like the A-300) is Part 25. This particular FAR Part addresses large civil airplanes and large civil helicopters. Transport category aircraft include jet-powered airplanes with 10 or more seats or with a maximum takeoff weight (MTOW) greater than 12,500 lb (5,670 kg); Propeller-driven airplanes with more than 19 seats or with a MTOW greater than 19,000 lb (8,618 kg); and Helicopters with a MTOW greater than 7,000 lb (3,175 kg). In the example you posted (an airliner and a “bizjet”) if the bizjet met the requirements to be certificated under the rules applicable to Transport Category jet powered airplanes, then, yes, both airplanes would be held to the same standard. You might wish to review the “yaw” standards posted in §25.351 - Yaw maneuver conditions. Here are those requirements:
The airplane must be designed for loads resulting from the yaw maneuver conditions specified in paragraphs (a) through (d) of this section at speeds from VMC to VD . Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the airplane inertia forces. In computing the tail loads the yawing velocity may be assumed to be zero.
(a) With the airplane in unaccelerated flight at zero yaw, it is assumed that the cockpit rudder control is suddenly displaced to achieve the resulting rudder deflection, as limited by:
(1) The control system on control surface stops; or
(2) A limit pilot force of 300 pounds from VMC to VA and 200 pounds from VC /MC to VD /MD , with a linear variation between VA and VC /MC .
(b) With the cockpit rudder control deflected so as always to maintain the maximum rudder deflection available within the limitations specified in paragraph (a) of this section, it is assumed that the airplane yaws to the overswing sideslip angle.
(c) With the airplane yawed to the static equilibrium sideslip angle, it is assumed that the cockpit rudder control is held so as to achieve the maximum rudder deflection available within the limitations specified in paragraph (a) of this section.
(d) With the airplane yawed to the static equilibrium sideslip angle of paragraph (c) of this section, it is assumed that the cockpit rudder control is suddenly returned to neutral.
I would call your attention to this last paragraph … where the airplane would be yawed to the static equilibrium sideslip angle and then the rudder control is suddenly returned to neutral. And we all should note that neutral is not full opposite rudder. While I am sure there are those on this forum who could provide chapter and verse and the formulas involved to calculate the inertia of the mass being “yawed” and what kind of force is generated on the vertical structure if full opposite yaw control (rudder pedal) were applied while the mass was moving in the original direction. If the airplane structure could withstand such reckless use of the rudder controls, all the better for the airplane – but to meet regulatory requirements, it is not required. There is no allowance for any transport category airplane anywhere (at least that I am aware of) that would not have a rudder application limit – and a cavalier “Van Damme kick to the rudder” at any time or place would likely result in that pilot’s immediate dismissal – assuming he and his airplane made it safely back to terra firma.
Tails of transport category airplanes do not “come off” at speeds under Va … or over Va … if the airplane is flown the way it was intended (and certificated) to be flown. In this particular case (AA587), the pilot flying successfully transitioned the first of two vortex encounters quite nicely – in fact, looking at the FDR, it was almost, if not, Textbook. The second vortex encounter was wholly different in that same pilot’s response. I have no idea if you have personally looked at any of the data applicable to this specific accident – if not, it would serve you well to do so … before you get to the level of conclusions that you’ve apparently reached without that kind of research/reading. However, and to alleviate your having to stomp around in knee-deep records, let me offer the following:
I’ll start by focusing on the initiation of the 2nd vortex encounter (and lest you wonder, my observations are directly from the FDR-generated display provided by the NTSB for this accident). The first control movement looks like it begins with a very small aileron correction to the left as the airplane looked to be ever so slightly beyond the 22-degree bank that seemed to be what the F/O was happy with during the departure turn as directed by ATC. Almost immediately we see a right rudder pedal deflection to the stop and a control wheel input of about 80 degrees to the right. The attitude indicator just prior to this looked to be relatively steady. Of course, with full right rudder and darn near full aileron to the right, the airplane begins a roll to the right (back toward wings level). However, when the airplane reaches about 20 degrees of left bank, while rolling to the right we see a simultaneous control wheel and rudder surface movement to the left. The rudder surface actually looks to exceed the pedal limit (but I don’t know if this is an anomaly with the indicator or not) and the control wheel gets to or awfully near to full control wheel displacement. This means at least 160 degrees of wheel change and probably something like 8 – 10 degrees of rudder surface change – all in 1 second, to the left as the airplane is rolling to the right. With all this opposite control surface input, the roll to the right is almost stopped (at about 10 degrees of left bank).
To add to the excitement, the full left aileron position is not maintained, nor is the full left rudder pedal position. The wheel is brought back to something like 10 degrees to the right simultaneously with application of full right rudder pedal deflection, again in 1 second, probably reaching constant rate saturation. While full right rudder is maintained, the control wheel is moved back to about 10 degrees left – again, in 1 second. As the rudder pedal deflection is maintained (very likely getting close to a stable sideslip), the control wheel is moved back to the right to just about full wheel travel and the rudder pedal exceeds the pedal limits (again, I don’t know if this is a display anomaly or if the actual limit was exceeded) – again taking a total of 1 second. Immediately, the rudder pedal is repositioned to full left deflection, and, in fact, goes well beyond the limits (again depending on the accuracy of the display), simultaneously the control wheel is deflected full left … again taking only 1 second. As the control wheel is moved back to the right (to about 45 degrees left), the rudder pedal deflection goes full right and the surface position presentation disappears, while the pedal position continues to show full right deflection. There is little doubt that this is where the data feed was stopped – probably because of the departure of the vertical stabilizer and rudder. The control wheel goes back to about neutral and back again to a right control wheel deflection of about 45 degrees.
The pitch attitude when this second event began was about 10 degrees nose up and the airspeed was 238 knots. By the time the rudder surface position display blanks out, there were 7 control wheel reversals and 5 rudder pedal reversals, all in about 7 seconds … and in this 7-second time frame the pitch goes to zero degrees and the airspeed increases to 251 knots.
Again, as I noted earlier, when an airplane is in a maximum equilibrium yaw, a sudden commanded full, or nearly full, opposite rudder movement against that sideslip can generate loads that exceed the “limit loads” and possibly the “ultimate loads” and can easily result in structural failure. I think the professional investigators reached this conclusion and if you have access to that NTSB produced video, you would likely see the same events and come to the same conclusion.
You know … throughout my participation on this forum, while it may have only been 8 years or so, I believe I have always tried to maintain some sense of professionalism with the others who participate here – even when they were off into some of their own beliefs – as limited as they were (and recognized as such by many other participants). So, if you would not object, I would like to suggest that you might consider reading the section of the rules that govern the certification of various kinds of airplanes.
In the US, those rules are written, published, and enforced by the FAA – in other countries those responsibilities fall to other organizations. As for the US, the section of the rules that govern the certification of transport category airplanes (like the A-300) is Part 25. This particular FAR Part addresses large civil airplanes and large civil helicopters. Transport category aircraft include jet-powered airplanes with 10 or more seats or with a maximum takeoff weight (MTOW) greater than 12,500 lb (5,670 kg); Propeller-driven airplanes with more than 19 seats or with a MTOW greater than 19,000 lb (8,618 kg); and Helicopters with a MTOW greater than 7,000 lb (3,175 kg). In the example you posted (an airliner and a “bizjet”) if the bizjet met the requirements to be certificated under the rules applicable to Transport Category jet powered airplanes, then, yes, both airplanes would be held to the same standard. You might wish to review the “yaw” standards posted in §25.351 - Yaw maneuver conditions. Here are those requirements:
The airplane must be designed for loads resulting from the yaw maneuver conditions specified in paragraphs (a) through (d) of this section at speeds from VMC to VD . Unbalanced aerodynamic moments about the center of gravity must be reacted in a rational or conservative manner considering the airplane inertia forces. In computing the tail loads the yawing velocity may be assumed to be zero.
(a) With the airplane in unaccelerated flight at zero yaw, it is assumed that the cockpit rudder control is suddenly displaced to achieve the resulting rudder deflection, as limited by:
(1) The control system on control surface stops; or
(2) A limit pilot force of 300 pounds from VMC to VA and 200 pounds from VC /MC to VD /MD , with a linear variation between VA and VC /MC .
(b) With the cockpit rudder control deflected so as always to maintain the maximum rudder deflection available within the limitations specified in paragraph (a) of this section, it is assumed that the airplane yaws to the overswing sideslip angle.
(c) With the airplane yawed to the static equilibrium sideslip angle, it is assumed that the cockpit rudder control is held so as to achieve the maximum rudder deflection available within the limitations specified in paragraph (a) of this section.
(d) With the airplane yawed to the static equilibrium sideslip angle of paragraph (c) of this section, it is assumed that the cockpit rudder control is suddenly returned to neutral.
I would call your attention to this last paragraph … where the airplane would be yawed to the static equilibrium sideslip angle and then the rudder control is suddenly returned to neutral. And we all should note that neutral is not full opposite rudder. While I am sure there are those on this forum who could provide chapter and verse and the formulas involved to calculate the inertia of the mass being “yawed” and what kind of force is generated on the vertical structure if full opposite yaw control (rudder pedal) were applied while the mass was moving in the original direction. If the airplane structure could withstand such reckless use of the rudder controls, all the better for the airplane – but to meet regulatory requirements, it is not required. There is no allowance for any transport category airplane anywhere (at least that I am aware of) that would not have a rudder application limit – and a cavalier “Van Damme kick to the rudder” at any time or place would likely result in that pilot’s immediate dismissal – assuming he and his airplane made it safely back to terra firma.
Tails of transport category airplanes do not “come off” at speeds under Va … or over Va … if the airplane is flown the way it was intended (and certificated) to be flown. In this particular case (AA587), the pilot flying successfully transitioned the first of two vortex encounters quite nicely – in fact, looking at the FDR, it was almost, if not, Textbook. The second vortex encounter was wholly different in that same pilot’s response. I have no idea if you have personally looked at any of the data applicable to this specific accident – if not, it would serve you well to do so … before you get to the level of conclusions that you’ve apparently reached without that kind of research/reading. However, and to alleviate your having to stomp around in knee-deep records, let me offer the following:
I’ll start by focusing on the initiation of the 2nd vortex encounter (and lest you wonder, my observations are directly from the FDR-generated display provided by the NTSB for this accident). The first control movement looks like it begins with a very small aileron correction to the left as the airplane looked to be ever so slightly beyond the 22-degree bank that seemed to be what the F/O was happy with during the departure turn as directed by ATC. Almost immediately we see a right rudder pedal deflection to the stop and a control wheel input of about 80 degrees to the right. The attitude indicator just prior to this looked to be relatively steady. Of course, with full right rudder and darn near full aileron to the right, the airplane begins a roll to the right (back toward wings level). However, when the airplane reaches about 20 degrees of left bank, while rolling to the right we see a simultaneous control wheel and rudder surface movement to the left. The rudder surface actually looks to exceed the pedal limit (but I don’t know if this is an anomaly with the indicator or not) and the control wheel gets to or awfully near to full control wheel displacement. This means at least 160 degrees of wheel change and probably something like 8 – 10 degrees of rudder surface change – all in 1 second, to the left as the airplane is rolling to the right. With all this opposite control surface input, the roll to the right is almost stopped (at about 10 degrees of left bank).
To add to the excitement, the full left aileron position is not maintained, nor is the full left rudder pedal position. The wheel is brought back to something like 10 degrees to the right simultaneously with application of full right rudder pedal deflection, again in 1 second, probably reaching constant rate saturation. While full right rudder is maintained, the control wheel is moved back to about 10 degrees left – again, in 1 second. As the rudder pedal deflection is maintained (very likely getting close to a stable sideslip), the control wheel is moved back to the right to just about full wheel travel and the rudder pedal exceeds the pedal limits (again, I don’t know if this is a display anomaly or if the actual limit was exceeded) – again taking a total of 1 second. Immediately, the rudder pedal is repositioned to full left deflection, and, in fact, goes well beyond the limits (again depending on the accuracy of the display), simultaneously the control wheel is deflected full left … again taking only 1 second. As the control wheel is moved back to the right (to about 45 degrees left), the rudder pedal deflection goes full right and the surface position presentation disappears, while the pedal position continues to show full right deflection. There is little doubt that this is where the data feed was stopped – probably because of the departure of the vertical stabilizer and rudder. The control wheel goes back to about neutral and back again to a right control wheel deflection of about 45 degrees.
The pitch attitude when this second event began was about 10 degrees nose up and the airspeed was 238 knots. By the time the rudder surface position display blanks out, there were 7 control wheel reversals and 5 rudder pedal reversals, all in about 7 seconds … and in this 7-second time frame the pitch goes to zero degrees and the airspeed increases to 251 knots.
Again, as I noted earlier, when an airplane is in a maximum equilibrium yaw, a sudden commanded full, or nearly full, opposite rudder movement against that sideslip can generate loads that exceed the “limit loads” and possibly the “ultimate loads” and can easily result in structural failure. I think the professional investigators reached this conclusion and if you have access to that NTSB produced video, you would likely see the same events and come to the same conclusion.
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I don't think dutch roll had anything to do what we are talking about. I think rudder movements caused the AA 587 crash and I will bet it had nothing to do with FO rudder movements. Yes, the rudders moved but what moved them? Where does the FDR sense the movement? I think it is the rudder actuator so what caused the actuator to move?
My first thought would be yaw damper, not a totally stupid pilot stomping on his rudders. I think both pilots were along for the ride because the captain would have prevented FO oscillations as all of us would. The yaw damper malfunction has happened before as in my previous post.
The FO didn't cause it in my opinion. NTSB has political pressure to lean to what works their way.
My first thought would be yaw damper, not a totally stupid pilot stomping on his rudders. I think both pilots were along for the ride because the captain would have prevented FO oscillations as all of us would. The yaw damper malfunction has happened before as in my previous post.
The FO didn't cause it in my opinion. NTSB has political pressure to lean to what works their way.
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Originally Posted by galaxy flyer
Va does not imply a strength criteria, it merely is the point at which the wings stalls at the G limit
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An interesting discussion, if only to demonstrate the longevity of misinformation circulating on the 'net 
AirRabbit
Well said
bubbers44
If you read the NTSB report and the performance group appendix you will see that AA installed additional instrumentation specifically to record cockpit control positions and forces. On p75 of the appendix you will find rudder position and pedal position; on page 77 you will find pedal forces. All of these correlate with each other, so I find it difficult to see how you can claim that FO input had nothing to do with this accident. How does the yaw damper provide forces at the pedal???
Teledorserious
To go back to your original post, could you please point me towards the actual FAA change to Va? I failed to find it.
AirRabbit
Well said
bubbers44
Yes, the rudders moved but what moved them? Where does the FDR sense the movement? I think it is the rudder actuator so what caused the actuator to move?
My first thought would be yaw damper, not a totally stupid pilot stomping on his rudders.
My first thought would be yaw damper, not a totally stupid pilot stomping on his rudders.
Teledorserious
To go back to your original post, could you please point me towards the actual FAA change to Va? I failed to find it.
Air Rabbit. An excellent post.
Unfortunately, there are none so blind as them as can't see.
The conspiracy theorists will have it their way no matter what the evidence.
Bubbers, take note.
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Unfortunately, there are none so blind as them as can't see.
The conspiracy theorists will have it their way no matter what the evidence.
Bubbers, take note.
Posted from Pprune.org App for Android
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For Bubbers,
The yaw damper would not have moved the pedals. It's a series yaw damper so rudder surface deflection occurs independent of pedal input. Secondly, yaw damper authority is a fraction of the full rudder travel.
The yaw damper would not have moved the pedals. It's a series yaw damper so rudder surface deflection occurs independent of pedal input. Secondly, yaw damper authority is a fraction of the full rudder travel.
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mono, the A300 yaw damper might not move the pedals but my friend felt they were going to crash trying to land at MIA because of the violent yawing causing them to go around. I wasn't there, just heard the story.
To me it seems more probable than pilot input why AA587 lost it's VS.
To me it seems more probable than pilot input why AA587 lost it's VS.