A Theory Refined
So why would the ADC tell fibs (to the FCS and rudder limiter) about the aircraft's speed? It might not be the same reason every time and I covered that in my initial hypothesis below ("just another Theory"- "Help me out here"). But in AA587's case I'd suggest that the initial wake turbulence encounter was enough to set the ball rolling. Normally, in a yaw, the fact that the port and starboard static ports are Y-connected is enough to iron out (and average) any discrepancies caused locally at each port (by the venturi effect of yaw angle, air inflow and fuselage curvature). However the design engineers only ever consider balanced flight and a few different configurations when they decide upon static port positioning and then map the PEC (position error correction) errors to be fed into the ADC. In yawed flight I'm suggesting that quite large errors can creep in and be magnified by both the ADC and FCS sampling rate (i.e. how often these gizmos ask for the speed info). As I pointed out in the earlier article, about a 9mb error (= an 8% error on an ISA day) can cause the airspeed to zero out. I'm not suggesting that that happened, just pointing out the small magnitude of P.E. error required to cause a largish speed error and an inappropriate rudder deflection.
Remember the 1996 Birgenair 757 that crashed due to static port tape being used to stop water getting into the static system during an aircraft wash (not being removed)? Well it's a little-known fact that heavy rain and a falling barometer can allow a static port to suck in water. It's usually not significant because it normally pools in water traps in the system - and gets drained eventually. In a heated system it's unlikely to freeze and cause the gross errors mentioned in my earlier article. However just consider what effect an amount of trapped water might have in the static system - particularly upon the timeliness of ADC sensing, and particularly in a yaw? The water may well flow in an adverse direction, due to the forces of the yaw and inertia, and induce a false static pressure (which is then picked up by the ADC etc etc). Do you see where I'm going here?
So if you accept my hypothesis, you will see that a clean/dry static
pressure system can induce an error into the FCS - and so can a wet one,
but probably a fiercer reaction in the latter case - because of the adverse
flow of trapped water during any yawing. Due to the unpredictable nature
of the water-flow and the ADC sensing, rudder ratio changes and resulting
erroneous FCS responses, any yaw meanderings would not be rhythmic, but
they could build up and become divergent and gross. Equally, the rudder
could become out of phase with the aircraft's physical yawing and tend
to assist the damping effect of the fin. Luck of the draw I would say
- but in AA587's case I'd suggest that the second wake turbulence encounter
came at just the wrong time in the cycle. It can probably therefore be
seen as a once off type accident. But if I am right, then there are things
that can be done to resolve any such discrepancies in the Airbus system's
logic and susceptibility to this type of error.
In answer to the question (at the end).
Pitot Pressure (as sensed by a pitot tube facing into the airflow) is a Total Pressure made up of dynamic and static pressure: T = D + S1
So obviously we need to subtract that S1 (the static pressure – which should be the ambient pressure at that height), because what we want to see on our Air Speed Indicator is D (dynamic pressure or the pressure due to our forward speed). The formula now becomes D = T – S1 (a conventional airspeed indicator does this within its internal plumbing but in a glass ship the Air Data Computer obliges).
But where do we source that S1? Well it comes from the static ports (let’s call it S2)…. which live on the port and starboard sides of the airplane’s aft fuselage (normally two small holes each side / above and below, which are heated against icing but which are open to the elements inflight and quite often left open and unplugged on the ground). If you park an airplane in the open and there’s heavy rain and it’s flowing down over those holes then capillary action can cause water to be sucked in, sometimes in fairly large quantities. I can recall an inflight emergency where I lost all pressure instruments after climbing through freezing level. They figured out later that, parked in the rain, water had been sucked up the hollow centres of the downward-facing rubber bungs (off which the water was dripping). When it later froze in the static system, of course I lost the altimeter (it froze at that height), the VSI went to zero and the Air Speed indicator just wound back to zero (from the 220knot climb speed). It was calculated later that that will happen (from that IAS) over a climbing height change of 2800feet (about = 9mbs or 8% of the Sea-level ISA pressure). The school solution is to depressurize and break the glass on the VSI and accept that there will be a fairly gross altitude error (due to using cockpit static). But that gets your ASI back in the picture (although greatly errored, trends will be good).
Now in an ideal world S2 will always be equal to S1. But we don’t live in an ideal world so there will always be some discrepancies:
D = T – S2 becomes an errored speed indication because it’s now a greater quantity than D = T – S1.
So it is starting to look like a real pakapoo ticket is it not? The business of getting a true S2 just became tantamount to impossible when we entered that first wake and began the tail-wag. When we entered it the second time, it was with a thoroughly confused system that was doing the wild thing with the rudder and, as luck would have it, the second encounter just happened to be at the worst possible angle and at the worst possible time. The rudder was already imposing a high load upon the fin and suddenly the additional load snapped an attachment lug and started the fin “working”. As the fin started its lateral dance of detachment, the rudder was excitated into a frenzy of fin-swaying corrections, leading to the death-rattle heard on the CVR.
Addressing what you’ve asked below:
“The static pressure drops to zero” isn’t really the case at all. What happens when water freezes in the static system is that the ambient pressure at that height is trapped and displayed (and that’s why the altimeter freezes at that height, even though the aircraft continues to climb). 2800ft later, the altimeter reads the same height - but the subtraction of that greater S2 (trapped pressure) causes the Air-speed indicator to wind back to zero. Over that 2800ft of climb the ambient pressure drops off about 9mbs (roughly 3mbs Hg/100ft). So if there was to be a momentary 9mb drop in the S2 (for any reason including the adverse flow of trapped static system water) at a speed of 220kts, the sensed speed would be nil, zero, zip. Now obviously that’s never the case because it’s a very dynamic situation we’re talking about here – but it does give you an idea of the magnitude of the pressure errors as sensed at the static ports – which could drastically influence the airspeeds being fed into the flight-control system. The speeds sensed could be at any one point of time either above or below the actual aircraft speed – thus causing the continually inappropriate FCS-dictated rudder responses (in reply to the gyro-detected yawing moments). And there’s the rub.
Question: I don't quite understand one thing in your "Just another theory". It deals with pitot tubes and static ports, which is to say dynamic and static pressure. If static pressure drops to zero, why would the speed reading decrease? If the speed is dynamic pressure minus static pressure, then lower static pressure would suggest a greater speed to me than a lesser one. At greater perceived speed, the rudder would be more (not less) limited in its arc of travel.
Answer: Just about right. They hit the first wake and the yaw gyros sensed a particular yaw angle and so the FCS computes a rudder deflection that should eliminate it. Unfortunately, in accordance with my theory, the Air Data Computer is being fed erroneous static pressures and that causes the calculated IAS to be wrong at any one moment (high sampling data-rate here). That miscalculation of airspeed causes the rudder limiter to allow the wrong deflection (i.e. one inappropriate to the aircraft’s actual speed). So correction number one didn’t work out - and we shall now try for rudder throw number two…. and so on.
For the same reasons, rudder deflection number two just causes a tail wag the other way and the yawing continues in this vein. If it was all working properly, then the rudder would centre just as the yaw came off the airplane. Alas, that’s not to happen, and it just happens to be at the apex of one of those amplitude-increasing divergent tail wags when they hit the second wake at just the wrong angle – and likely broke just the one fin attachment lug. Because of the characteristic of composites whereby the strength of a structure is within the whole, the whole fin started “working” –as driven by the rudder. The wobblier the fin became as further lugs fractured, the harder the system worked the rudder – to compensate.
The whole point about water in the static system is not that it would be frozen at that height (because of course it would not be) – but water has inertia and we’re talking about sensing small increments of pressure here. Try sucking water up a straw. You’ll then appreciate the reverse effect that water (flowing under its own inertia due to the yawing moment) can have upon sensed static pressure within the system. Quite considerable.
The whole discussion about frozen water was simply to enable an example of how very small pressure changes can exert a considerable effect upon the indicated airspeed. Here we are talking about the high sampling rates and instantaneous readings peculiar to computer-driven flight control systems.
It’s not really appropriate to start spouting real life figures because then some smarty pants will just grasp at an error of magnitude. It’s not about magnitude - it’s about the timing, sampling rates and the dynamics and development of the situation (best described as a state of flux and incredibly bad luck).
It’s probably not a valid theory – but then again it may be (but either way, it’s certainly filling a giant vacuum at present).
I think it's clear. As the airspeed winds back to zero, the system thinks the plane is going slower, so it allows the rudder limiter more authority. But the movements wind up out-of-phase, leading to the dynamically destructive scenario you portrayed. Very subtle, my man.
|My interspersed bits in
blue and magenta. Rainman may well comment further.
From Rainman (a flight controls engineer)
Not really sure what he's describing here because
I very much doubt that there are any bike chains in this system.
Question along this line: Is the 10 DEG rudder limit applied to the entire
posted 14 February 2002 03:16
posted 14 February 2002 13:15
Just another Theory (the original submission)
Preliminary information based on FDR data and an analysis of the manner in which rudder position data is filtered by the airplane’s systems indicates that within about 7 seconds, the rudder traveled 11° right for 0.5 second, 10.5° left for 0.3 second, between 11° and 10.5° right for about 2 seconds, 10° left for about 1 second, and, finally, 9.5° right before the data became unreliable. [FOUR complete reversals inside 7 seconds].
These are the last 7 seconds the fin and rudder were well enough attached to give reliable FDR readings. The FDR shows FOUR complete rudder reversals inside "7 seconds" but the sum of the intervals given only comes to 3.8 seconds and we are not given the travel time of the last rudder movement to the right. The letter does show that the last reliable FDR reading shows the a/c in a left yaw of 8-10° and that that combination exceeded the structural strength.
We are only given the last yaw angle -- I would really like to see yaw vs. rudder position in these 7 seconds. In my opinion, the crew would have to be Tour de France material to work the pedals that fast against a 32 pound force. This inclines my suspicions to the flight control system.
a. AA587 + eight prior uncommanded yaw incidents (U/Y) + a bent FEDEX rudder actuator rod (with associated fin disbonding)
b. An anecdotal reputation (particularly amongst F/A's) for the A300 being a tailwagger.
c. early design, early software
d. Composite fin more prone (than metal) to sudden failures due to its axial strength, load-bearing capabilities and "strength only through integrity" characteristic.
e. five recorded rapid high-throw rudder deflections at 255kts during and immediately after the second wake encounter (before fin failure) on AA587 (assuming this was the filtered DFDR read-out). The A300-600 rudder can move at a rapid 39 deg/sec.
f. neither autopilot nor Control Wheel Steering engaged.
g. There was an audible "rattle" from aft (as recorded on the AA587 CVR) after the second wake encounter (just prior to fin failure).
2. Assumptions: (intuitive, logical, recommended and standard practices)
A pilot would not (normally) make a large rudder input at 255kts (about
1.9Vs) [see Boeing/Airbus Guidance at:
The FAA guidance for pilots is transparently a bit of an irrelevant butt-covering exercise - but it does give us the heads up that an oscillatory rudder (see Farley above) certainly can break a fin. Happy with that - because that oscillatory motion is equivalent to aerodynamic or system induced "flutter". Flutter has always been associated with imminent catastrophic failure - as it often tends to be oscillatory and divergent. The nature of flutter is such that it all happens too quickly to enable recognition and speed (or configuration) changes. It's best avoided altogether (through design).
b. Pilots had no time in which to react - either appropriately or inappropriately (about 7 secs for recognise and respond/react). They would not have recognised the fin loss and the full power response probably did not materially affect the outcome.
c. The uncommanded yaw can happen on approach (i.e. slow and configured) or at FL310 so that common factor gives us a further clue. It doesn't appear to be a factor on/near the ground (i.e. you need a speed range).
d. As not all uncommanded yaw instances resulted in disastrous failure, there must be a trigger and exacerbating factors (i.e. present for AA587). We know of one - wake turbulence. Possibly another - the preflight yaw damper reset by maintenance. We can possibly refine this to "a characteristic (of design) plus an unserviceability" in order to complete the chain. Or maybe it's all predicated by design plus circumstance......no u/s required.
b. The rudder trim switch had been subject of two AD's, the second one of which called for a lengthened switch shaft and a wiring change. That gave three rudder-trim switch maint mischoice possibilities (plus the wiring change). There was also an aileron trim AD for inadvertent actuation (both switches being close adjacent on the 408VU panel - increasing the possibility of a maint "Murphy")
c. Intermittent wiring problems have been known to cause similar flight control erratics....but oscillatory?? Maybe not, that's more likely a systemic interaction.
d. The Loral DFDR had been previously found unsatisfactory by investigators because it recorded only a filtered sampling of data (as per that displayed in the cockpit). The FCS is in fact capable of moving the rudder more than twice in the time that the DFDR records one motion. This may help explain the "rattle" (lateral fin-rocking) recorded on the DFDR and yet indicate that the pilots could not have possibly moved the rudder as fast as the FCS was capable of doing (39 deg/sec) - but without it being recorded.(See this article regarding AA587’s DFDR deficiencies (low sampling rate). And see also [url=http://www.srg.caa.co.uk/publications/CAP455_airworthiness_notices.pdf]Loral_800_"obsolescence"
4. Theory: Considering all of the above and extrapolating into plausible explanations, here's an hypothesis:
a. Static port positions are chosen by designers in order to minimize local pressure disturbances throughout the speed and configuration range - so that inputs to the Air Data Computer (ADC) are as solid and error-free as possible. However they tend NOT to consider what happens under yawed flight. In a steady-state yaw (asymmetric, engine-out, OEI) the port/starboard discrepancy is soaked up by the fact that there are ports on both sides of the fuselage and the two pressures are joined in a Y Junction before being presented to the ADC. If, however, the yaw angles become much greater, there is both an asymmetric blanking and a pressure change (venturi) effect caused by both the rapid reversal of yaw and the extent and periodicity of that yaw. The actual data (incremental pressure changes) presented at the static port has a long way to travel and it can become both erroneous and out-of-phase - simply due to the dynamics of the situation (and perhaps even due to the sampling rate of the ADC / and its subsequent input rate to the FCS). PEC (position error correction for static errors) is normally applied to the ADC only for balanced flight in the envelope - for normal configurations.
b. Consider further the effect of yaw if the static ports are placed in an optimal error-free position for balanced flight - but just happen to be at a point of curvature on the fuselage where the venturi effect on the sensed pressure, under yaw, will be very significant (i.e. perhaps producing large local pressure drops).
c. Airspeed is Dynamic Pressure (but is sensed at the pitot as TOTAL Pressure (static plus dynamic). The static port pressure is deducted to give you the indicated airspeed (i.e. the dynamic only). If you need a yard-stick for the significance of static port pressure errors on the sensed airspeed, take note that trapped water freeze-blocking a static line (in a climb above freezing level) will cause an ASI to wind down from 220kts to zero in just under 2800ft of climb (equivalent to 9mbs Hg). It would also cause the VSI to zero and the altimeter to freeze - but that's irrelevant here. I'm just pointing out that quite a small static port pressure discrepancy can have a large effect on ADC-sensed airspeed. Those sensed airspeeds control yaw-damper action and rudder ratio limiting - at any one point in time.
d. So why (and how) might this be significant? In prior instances of U/Y there may have been a minor out-of-phase disagreement between the FCS and the ADC. As pointed out, this could be caused by the rudder limiter's throw (and/or yaw damper input) being misset by the flawed ADC airspeed info. A yaw requiring an 18% displacement throw of the rudder might end up being much greater because the rudder limiter is permitting (courtesy of the ADC and FCS software) up to 16 degrees full throw (instead of the 10 deg appropriate to the actual airspeed at that instant). That much greater (or simply inappropriate) corrective yaw causes another ADC sensing error and the tail goes into "wag" mode (familiar to A300 backenders). Eventually the damping effect of that large fin normally causes these rudder-induced overshoot oscillations to dissipate. But you need a trigger for this to "kick off" - some initial yaw. atmospheric disturbance, pilot "stretching" with feet on rudder pedals (done it myself), climb/descent through an inversion, wind-shear, turbulent air, wake turbulence etc etc. Help me out here.
But why did AA587 self-destruct? I think it may be as simple as the first
wake encounter starting the tail wagging the dog and the next wake encounter
really confusing the issue and leading to an instantaneous fin overload
(the rattle of that semi-detached fin indicating that lugs had sheared
and that the fin was on its way). These fins and rudders on the big twins
are quite powerful because they have to accommodate OEI any old time.
Courtesy of the
f. Why the bent FEDEX rudder actuator rod? Sometimes the fin wins and the periodicity of the feedback to the rudder is sufficient to bend the actuator.
g. Or perhaps it wasn’t the FCS computer’s fault at all, perhaps it was a cross-wiring or a reversed hook-up of the static ports or a hydraulic servo valve. Whatever it was, if it was a hardware flaw, it may have been that way for a while. Sometimes you need a significant divergence (such as that caused by wake turbulence), before you cross a trigger-point for feedback mayhem. Think of the microphone analogy. If you place an open mike in front of a speaker that’s in the same circuit, as long as the volume is way down there’ll be no feedback squeal. Wind it up a little and it will suddenly be deafening.
5. No doubt someone will now tell me that yaw-induced static port pressure errors aren't of sufficient magnitude to start this ball rolling....
or that this is all accommodated in a cunning ADC/FCS program patch of narrow notch (or broadband) filtering.. But as I said, it's just a theory that seems to fit the bill (and, if correct, one flaw which should be easily fixed). I don't think that it's a fin strength composite) problem. Whether or not this static port theory holds water, it's still likely to be an FCS computer-derived (and driven) flutter that arises from a sensing problem. Just keep in mind that the rudder has two movements. Rudder deflection resolves the yaw (or is intended to) and then (in an ideal world) centering sends the rudder to parade rest, calmly zero'd in trail. But in my theory, there was always a bunfight going on back there and a simple recentering was rarely on the cards; sometimes the yawing gave cause for alarm, but normally the static port error wasn't significant because the yaw wasn't massive and so the natural damping of that large fin meant there were only a few wags at most. But the potential was (and is still) there for an A300 to eventually meet its wake encounter match. One is left wondering what the "break-out" yaw angle/rate is (that might "dud" the static port pressures and set this ball rolling).
And while we're at it, also think about the effect of some trapped water in one side of the static system - and what effect its inertia might have (on the ADC sensed pressure) during yaw gyrations.
Belgique - Masterful!
Your lucid and clearly communicated static port theory is the first I have seen that explicitly addresses the "second failure mode" of the independent rudder-travel limiting mechanism seemingly failing to keep rudder deflections inside the stipulated limits for the aircraft speed.
The AA587 aircraft might have ridden through the wake events OK if the extreme deflections of rudder had been limited according to the speed vs throw-angle rules - even if the First Failure Mechanism for rudder oscillation were present due to design instability, wiring faults, or both.
One suspects the static system value data has considerable mechanical and/or electronic "buffering" in it, so that responses to individual port changes are averaged over a time interval of seconds, at least. Because it is the baseline reference for barometric calculations, fast response to static pressure variations would just add noise to the gauges and controls.
If this is the case, a pressure-change from the first wake event (and related slip?) might have created enough distortion of the averaged static pressure that the computed airspeed for rudder limiter purposes was scaled back toward the 165K mark where up to full rudder travel is allowed. The inter-wake interval leaves time for the (probably fairly slow) mechanical travel interlocks to reposition for a substantially wider rudder travel window. Then comes wake bump number 2, which sets off the First Failure Mechanism on a cycle of uncommanded rudder oscillations. Because the rudder travel window is inappropriately wide at that moment, those swings do the tail in, in short order.
I sure wish someone would start measuring how "normal" A300's wiggle their tails in the course of a working day....per my comment on the "Pilots want A300's grounded " thread.
MICHAEL A. DORNHEIM/LOS ANGELES
Rudder limiter protects aircraft when flying straight. But pilots should know: In a side slip, 'you're on your own.'
Simple rudder motions on the Airbus A300-600R, the aircraft type in the American Airlines Flight 587 crash, can create forces exceeding ultimate load on the vertical stabilizer and possibly break it off, according to an engineering analysis by Aviation Week & Space Technology. The study was prompted by the loss of the A300 last Nov. 12 after takeoff from New York John F. Kennedy International Airport, but the principle applies to transport aircraft in general.
The critical condition is to put the aircraft into a sideslip, and then apply full rudder against the slip. The Federal Aviation Regulations (FARs) only require that the fin be strong enough to take neutral rudder in a slip, and the load increase from applying opposite rudder can be more than the 50% margin to which aircraft are designed. Applying rudder against a slip may exceed the ultimate load of the fin, depending upon aircraft characteristics.
Some pilots believe that limiting mechanisms prevent the rudder from causing structural overload and allow essentially carefree use of the rudder, but this is not so. "I think you can break an aircraft in any axis if you work on the controls," said a flight controls expert at a major airframe company. "Operating aircraft relies on basic airmanship."
The Flight 587 flight data recorder (FDR) shows three lateral accelerations of 0.3g and 0.4g right, and 0.4g left, in the approximately 7 sec. before it appears that the fin came off. Aviation Week estimates show that the aircraft may have been in a full slip to produce the high accelerations. During the same period the FDR shows the rudder making about five deflections of 5 to 10-11 deg., culminating in a rudder reversal immediately before the fin apparently came off. The 10-11-deg. deflection is the maximum allowed by the A300's rudder limiter, suggesting it was working correctly.
The large rudder deflections and sideslip angles indicated by the FDR are the combination of conditions that may overload the fin. It is not clear why the rudder made these motions. The National Transportation Safety Board (NTSB) said last week that it was still investigating both mechanical and pilot sources for the motions.
Though the A300 has a rudder limiter to protect the tail structure, it allows fin side loads of about 80,000 lb. at 250 kt., close to the approximately 240-kt. airspeed when the rudder went haywire and the fin came off. That is 60% higher than our estimate of the design load and possibly at a level where the fin could fail. These loads are greater than other conditions that might design the tail, such as the roughly 20,000 lb. needed to balance an engine failure at takeoff and the 40,000-45,000 lb. that we estimated to be the maximum gust load.
Airbus declined to comment on our estimates due to the ongoing NTSB investigation, but said the questions raised "are of relevance."
THE FARS DON'T COVER the opposite rudder case because regulations have been written with consideration for the intended use of the airplane, said an FAA official with the Transport Airplane Directorate in Seattle. "Transport category airplanes, by virtue of their passenger carrying function, are not intended to be subjected to violent maneuvering conditions. The suitability of the current regulations as being adequate to ensure an acceptable minimum level of safety is based, in part, upon the historical record of vertical tail failure incidents."
That record appears to show that Flight 587 is the first time in civilian jet transport history that a fin has come completely off due to aerodynamic loads. It is on the order of the one-in-a-billion-flight-hours safety rate targeted by the FARs.
Avoiding abrupt rudder usage is considered standard airmanship by the industry, an Airbus official said. Both Boeing and Airbus have highlighted concerns about rudder usage in their upset recovery literature, warning that too much rudder could lead to loss of control or even structural failure.
Fin loads are mostly described in two paragraphs in FAR Part 25, which regulates airworthiness of transport aircraft. Paragraph 25.351 covers yaw maneuver conditions, and 25.341 covers gust and turbulence loads. Our analysis of the A300 found that yaw maneuver is the critical case.
Paragraph 25.351 spells out a simple maneuver and requires that the manufacturer analyze the loads at four conditions (see graphic, p. 24). The maneuver is to:
An FAA official said each condition tends to load different parts of the fin, such as the front spar, rear spar, hinges, rudder, etc., but condition D can create the highest fin bending loads, as far as the regulations are concerned. But he noted that some aircraft may have different critical conditions.
Conditions A, B, C and D are also marked on the graph of tail loads (below). The units on the left-hand scale are tail lift coefficient, referenced to the estimated fin area of 470 sq. ft. The center line shows how tail lift varies with sideslip, with the rudder fixed in the neutral position. The slope of the line is negative per aeronautical convention. At 15 deg. it starts to wash out due to nearing maximum lift and fuselage interference.
The upper and lower lines show the effect of applying 10 deg. of left and right rudder. The A300 rudder is relatively powerful because it is large--about 34% of the total fin chord. Rudder effectiveness also washes out with increasing sideslip, and this affects the critical anti-slip rudder (the bottom line) more than pro-slip rudder. The 10-deg. rudder deflection is approximately the maximum allowed by the rudder limiter at 250 kt. We assumed that 10-deg. of rudder would give a 10-deg. sideslip, which is typical behavior.
The cases of applying rudder against the slip are shown as conditions X and Y. Condition X is a rudder reversal in a steady sideslip, and Y is a reversal timed when sideslip reaches peak overswing, and gives the highest loads. Both would be achieved by applying left rudder long enough to build up sideslip angle, then quickly hitting right rudder. The red line is 1.5 times the design load as defined by condition D, i.e., the ultimate load. Both conditions X and Y exceed the ultimate load.
ULTIMATE LOAD, PER FAR 25, only needs to be tolerated for 3 sec. and can result in permanent deformation. There appears to be no requirement on what the strength must be after surviving ultimate load.
The chart is a presentation of static conditions, and real aircraft behavior is dynamic. The 39 deg. per sec. rudder only takes 0.5 sec. to move from stop to stop at 250 kt., during which we estimated dynamic sideslip angle would reduce by 1-1.5 deg. at condition X, and slightly more at condition Y. With this small reduction the loads are still at or beyond 1.5 times condition D. Also, reversing the rudder while swinging outbound through point C would produce high loads because fuselage rate would act to increase effective sideslip angle.
Because of the rudder deflections shown, this chart is valid only for one speed, and the actual critical load could occur at another speed; for example, a lower speed with a larger deflection. However, we believe it correctly portrays the concept, and the fact that the relatively fixed engine-out and peak gust loads are not critical means condition D is likely critical, which supports the importance of conditions X and Y.
Given the limited amount of FDR data released by the NTSB it is not clear if forces approaching X and Y were achieved, but the high sideslip and rapid full rudder motions are ripe for this possibility. The exact motions may never be known because the FDR only measured the rudder twice per sec., while it can move at 39 deg. per sec.--it could go from neutral to the stop and back between samples. And as the NTSB noted last week, fast rudder motions were distorted by being filtered.
The right axis on the graph converts the lift coefficient into actual force at 250 kt., to give a feel for the magnitudes and provide a comparison with other loads on the tail. Conditions X and Y are essentially the equivalent of hanging a fully-loaded 18-wheel truck from the fin. With that consideration, the thick fin attachment points start to look not so thick.
Compliance with the FAR 25.351 yaw maneuver may be shown analytically, and the airframe manufacturers may not actually conduct a full-force maneuver. They are leery of taking an aircraft to 100% design load in flight, and may limit loads to the neighborhood of 80%. The flight test data is to validate a model, which combined with ground test results shows the aircraft complies with regulations.
MANUFACTURERS ARE FREE to have their own requirements that exceed the FARs, and Airbus declined to comment on its techniques. Boeing says its requirements are proprietary, but one engineer familiar with the company believed the requirement was that rudder reversal in a sideslip, i.e., condition X, not exceed ultimate load. This may not require extra metal, because meeting design load with other conditions was usually more critical.
The A300 rudder has ±30 deg. of authority at speeds below 165 kt., and the limiter progressively cuts this back to 3.5 deg. at maximum speed. It may be tempting to further limit the rudder at higher speeds, but it needs enough authority to handle engine failure with some margin, and serve as a yaw damper. There also are unusual conditions such as multiple leading edge flap failure that may require a large amount of rudder to counteract.
|to Hot off the Press|