AF 447 Thread No. 5
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Amen to Owain's post.
#1 - Amen to Owain's post.
#2 - Some of the contributrors here may benefit from being reminded that lift & drag are defined as the aerodynamic forces perpendicular & along, respectfully, to the free airsteam, NOT the airframe.
As far as the THS is concerned, with the airplane 'falling' @ ~60deg AOA (e.g) drag would contribute more pitch down moment than lift.
#2 - Some of the contributrors here may benefit from being reminded that lift & drag are defined as the aerodynamic forces perpendicular & along, respectfully, to the free airsteam, NOT the airframe.
As far as the THS is concerned, with the airplane 'falling' @ ~60deg AOA (e.g) drag would contribute more pitch down moment than lift.
maintaining the stall
Excellent presentation by Owain, just peachy.
I also wonder about recovery possibilities if the THS was moved to the opposite position/angle from that which existed all the way down to the ocean. Pitch moments would change, and it's possible that the nose would come down
In other words, as with the Viper, at extreme AoA, you still have nose up authority, but not down. We had the equivalent of the THS all the time - no elevators at the rear of the stabilators. All but delta designs of the lites use the all-moving stabilzer. Because we had a lotta lift from the fuselage and the strakes, our aero c.p. was quite a ways forward of the airliner designs. So we could get to a fairly stable condition, stalled, but still with nose up authority.
I would add that as the speed rapidly slowed and the jet entered the abnormal AoA regime that there could be a brief time where downwash could "help" increase the nose up moment, like the classic T-tail deep stalls. In other words, without doing anything the AoA would continue to increase, driving the main wing further into the stall until equilibrium was reached.
Lastly, I do not believe that an extreme nose down pitch attitude would be required to recover. As 'bird suggests, this is where an AoA indication could really help. We aren't talking about 100 knots of extra speed, as it looks like we already had 150 knots or so of effective forward speed. Get to 200 + knots and start pulling back to level flight. And BTW, I always thought the gee requirements for the heavies was 3 gees, plus a standard addition until structural failure. The 2.5 gee number surprised me.
I also wonder about recovery possibilities if the THS was moved to the opposite position/angle from that which existed all the way down to the ocean. Pitch moments would change, and it's possible that the nose would come down
When you look at this, it is clear that the question to be asked is not will there be enough down elevator power to give a ND recovery moment, but rather will there be enough elevator power to get and hold 60 deg AoA. To recover all that would be necessary would be to remove the up elevator, although of course some down elevator wouldn’t hurt – so long as you don’t stall the THS.
I would add that as the speed rapidly slowed and the jet entered the abnormal AoA regime that there could be a brief time where downwash could "help" increase the nose up moment, like the classic T-tail deep stalls. In other words, without doing anything the AoA would continue to increase, driving the main wing further into the stall until equilibrium was reached.
Lastly, I do not believe that an extreme nose down pitch attitude would be required to recover. As 'bird suggests, this is where an AoA indication could really help. We aren't talking about 100 knots of extra speed, as it looks like we already had 150 knots or so of effective forward speed. Get to 200 + knots and start pulling back to level flight. And BTW, I always thought the gee requirements for the heavies was 3 gees, plus a standard addition until structural failure. The 2.5 gee number surprised me.
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As far as the THS is concerned, with the airplane 'falling' @ ~60deg AOA (e.g) drag would contribute more pitch down moment than lift.
#2 - Some of the contributrors here may benefit from being reminded that lift & drag are defined as the aerodynamic forces perpendicular & along, respectfully, to the free airsteam, NOT the airframe.
As far as the THS is concerned, with the airplane 'falling' @ ~60deg AOA (e.g) drag would contribute more pitch down moment than lift.
As far as the THS is concerned, with the airplane 'falling' @ ~60deg AOA (e.g) drag would contribute more pitch down moment than lift.
If it is, then doesn't that drag force act "up" (a component of the vector, anyway) in a nose pitch down moment, when acting through the arm based on where the CG resides?
I am still left with the picture from my little FBD here that you have forces acting normal to the axis of the aircraft (vector components of vectors acting along the axis you point out) which tend to pitch the nose down, the only component acting against that being the body drag vrom CoG forward to the nose. There's more surface area aft of CG, based on my sketch. Am I missing something here?
Harry, we aren't at a high Mach number if we are stalled, are we?
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Here is a summary of what I gleaned from your post. Please let me know what I have misunderstood.
Once stalled at altitude, and with CG as estimated, (be it 29% or 37%) the center of lift of the wings will provide a self-correcting nose down pitching moment proportional to such lift as the wing is still creating.
Once stalled at altitude, and with CG as estimated, (be it 29% or 37%) the center of lift of the wings will provide a self-correcting nose down pitching moment proportional to such lift as the wing is still creating.
Per the FBD I just sketched on my napkin, the arm that the force acts through is of the length somewhere between 70-29 to 70-37, (as outer boundaries). Based on where the THS is, I'd guess its relative number for arm calculation is about 96 or 97. (Am I close?) as compared to the center of pressure on the wing. (@takata: thanks for the posted CoG chart).
Thought: IF THS and elevator (as a lift producing system) have lost control authority or were stalled, THEN there would initially be no force from the back end, or a very small force, countering the "self correcting" pitch down moment of the stalled wing.
Thought number 1:
As the nose attitude gets closer to level, wouldn't the C of P start to move forward from 70 towards a smaller value, and gradually reduce the length of the arm, and thus the moment, of the correcting tendency?
As the nose attitude gets closer to level, wouldn't the C of P start to move forward from 70 towards a smaller value, and gradually reduce the length of the arm, and thus the moment, of the correcting tendency?
Thought number 2:
THS is an airfoil, so even if stalled, it produces some amount of lift and thus provides, through that longer arm, some counter to the correcting tendency of the wing whether or not it is stalled.
THS is an airfoil, so even if stalled, it produces some amount of lift and thus provides, through that longer arm, some counter to the correcting tendency of the wing whether or not it is stalled.
Your line of thought presents me with the provisional conclusion that the THS was not stalled, since the nose stayed up (per the BEA report) and didn't (as far as we know) oscillate up and down as it might if the THS were stalled.
What you described is a "natural" pitching down movement (??) of the stalled wing, which seems to have been countered by the longer arm being acted on by lift from an unstalled tail/THS.
Am I close?
Am I close?
If that's about right, it leaves me with a non trivial concern:
if the nose stayed up due an input or command other than pilot control inputs, the nose being held up by (THS lift) x (arm) prevented stall recovery for about 30 thousand feet worth of travel down to the surface.
if the nose stayed up due an input or command other than pilot control inputs, the nose being held up by (THS lift) x (arm) prevented stall recovery for about 30 thousand feet worth of travel down to the surface.
The above sort of reasserts a fairly obvious point: an ounce of stall prevention avoids about 200 tons of attempted cure. 
Owain:
I went back to sketch, and yes, I should have been looking more like 300 or more. Came back to your post, thanks for the 400.
By lost control authority I meant 'move your flight control X amount' (by telling the Flght Control computers to move it as far as it will go) and get less response than what will change your pitch, or get no change.
Go to full deflection (or largest allowed command) and get less change than needed, or none.
That to me is loss of control authority. I may be using the term in a non standard way.
Part of what was in the back of my mind as I typed that was that each knot lost costs you lift/force as the square of the velocity lost, roughly. I move the stick and expect a response, don't get one, so keep moving the stick more to try and get it to respond at all ... mind goes back to the old slow flight demonstrations, and how "sloppy" the control responses are compared to cruise flight.
Rudder travel limits are a built in control authority limit. (To avoid damage to tail). That's an intentional limit to control authority. What I was thinking about was inadvertent loss of control authority in the pitch channel due to loads/forces being insufficient for the amount of force needed to change pitch.
Also in the back of my mind as I wrote that is a loss of control authority problem in helicopter tail rotors, where you put controls to full deflection but are unable to overcome the torque due to limitations on control surface travel (or tail rotor blade stall). In that case, your authority over your nose, (yaw) is lost, since it goes one way while you are trying to get it to go the other way, or just stop it from moving.
Thanks for the clarification. 
I went back to sketch, and yes, I should have been looking more like 300 or more. Came back to your post, thanks for the 400.
By lost control authority I meant 'move your flight control X amount' (by telling the Flght Control computers to move it as far as it will go) and get less response than what will change your pitch, or get no change.
Go to full deflection (or largest allowed command) and get less change than needed, or none.
That to me is loss of control authority. I may be using the term in a non standard way.
Part of what was in the back of my mind as I typed that was that each knot lost costs you lift/force as the square of the velocity lost, roughly. I move the stick and expect a response, don't get one, so keep moving the stick more to try and get it to respond at all ... mind goes back to the old slow flight demonstrations, and how "sloppy" the control responses are compared to cruise flight.
Rudder travel limits are a built in control authority limit. (To avoid damage to tail). That's an intentional limit to control authority. What I was thinking about was inadvertent loss of control authority in the pitch channel due to loads/forces being insufficient for the amount of force needed to change pitch.
Also in the back of my mind as I wrote that is a loss of control authority problem in helicopter tail rotors, where you put controls to full deflection but are unable to overcome the torque due to limitations on control surface travel (or tail rotor blade stall). In that case, your authority over your nose, (yaw) is lost, since it goes one way while you are trying to get it to go the other way, or just stop it from moving.
As I said in that earlier posting, without significant and continued up elevator application, even with the THS at -13 deg, you won't get any nose up pitch from the back end above about 26 deg AoA. Well maybe I didn't actually SAY that, but it is implicit in my argument. 26 deg is more than enough to stall the aeroplane of course, but AF447 has been linked to much larger values of AoA that, IMHO, can only be achieved with up elevator. I will leave it to others to argue whence came that up elevator.
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And BTW, I always thought the gee requirements for the heavies was 3 gees, plus a standard addition until structural failure. The 2.5 gee number surprised me.
The actual requirement is a complicated formula that depends on MTOW, but pretty well all civil airliners this equates to 2.5g as a 'limit' load with a factor of 1.5 to be applied for ultimate (breaking) load. i.e. 3.75*MTOW is the design condition. At less than MTOW of course one could pull more.
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g limit
Gums
G limit in the heavy industry:
minus 1 till plus 2.5 in clean configuration
0 (zero) till plus 2.0 with high lift devices in anything but UP.
Of course, like Owain states, Ultimate (breaking) load is 1.5 times limit load.
G limit in the heavy industry:
minus 1 till plus 2.5 in clean configuration
0 (zero) till plus 2.0 with high lift devices in anything but UP.
Of course, like Owain states, Ultimate (breaking) load is 1.5 times limit load.
The point I was trying to make was that it was the Flight Law that was determining the position of the control surfaces and not the AP.
Therefore my reference to Normal Law to make this difference a bit clearer.
In this case after AP disconnect the system would have soon/immediately/at the same time switched to Roll Direct Law and thus very likely to control surfcaes neutral.
Control surfaces neutral means just that. It doesn't mean the aircraft will try to level the wings. But it means the ailerons will not actively contribute to further roll after change to Roll Direct Law (AP doesn't enter into the equation at all), only inertia and/or other forces will.
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Hi Owain,
You did Sir! great posts, and please, don't stop making many other positive contributions like that...
Originally Posted by Owain Glyndwr
I do think that I may be able to make a positive contribution towards understanding the aerodynamics
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About the lost infos due to Unreliable Airspeed.
The characteristic speed function (see above) computed by the FE (Flight Envelope part of the FMGC) is lost, meaning that Speed limits, computed by the PRIMs (FCPC), are also lost, including Valpha_prot (Normal) and VSw (Alternate).
Two other functions, displayed by the Primary Flight Display (PFD) were also lost (flagged, see ACARS): Both Flight Director (FD) and Flight Path Vector (FPV) were flagged. This is very bad as the FPV would also provide a visual form of AOA via the Flight Path Angle:
The characteristic speed function (see above) computed by the FE (Flight Envelope part of the FMGC) is lost, meaning that Speed limits, computed by the PRIMs (FCPC), are also lost, including Valpha_prot (Normal) and VSw (Alternate).
Two other functions, displayed by the Primary Flight Display (PFD) were also lost (flagged, see ACARS): Both Flight Director (FD) and Flight Path Vector (FPV) were flagged. This is very bad as the FPV would also provide a visual form of AOA via the Flight Path Angle:
Last edited by takata; 14th Jul 2011 at 20:27.
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Originally Posted by takata
Speed limits are lost, including Valpha_prot (Normal) and VSw (Alternate).
Last edited by HazelNuts39; 14th Jul 2011 at 20:23.
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Hi HN,
Thank you for pointing that for clarification. This is what I meant by specifying the Law related to each. VAlpha_prot (Normal) is replaced by VStall_warning (VSw) in Alternate... But, it should never be displayed on the PFD if the change of Law is due to an UAS event. Instead, there is a flag -SPEED LIMIT- meaning that the limits (high an low) are lost.
Thank you for pointing that for clarification. This is what I meant by specifying the Law related to each. VAlpha_prot (Normal) is replaced by VStall_warning (VSw) in Alternate... But, it should never be displayed on the PFD if the change of Law is due to an UAS event. Instead, there is a flag -SPEED LIMIT- meaning that the limits (high an low) are lost.
Hi Gums,
I think that perhaps on the graph of the pitching moment of the Viper the points where the stabilator starts stalling may be detectable. Do you know its max deflections? was it a symetric airfoil?
I think that perhaps on the graph of the pitching moment of the Viper the points where the stabilator starts stalling may be detectable. Do you know its max deflections? was it a symetric airfoil?
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Originally Posted by HazelNuts39
can you please provide a reference?
My notes are:
"Apparition du SPD LIM RED FLAG sur l'échelle des vitesses des PFD.
Les parties FLIGHT ENVELOPPE des deux FMGEC sont INOP.
Perte des informations:
- VLS, S, F, GREEN DOT, Vtrend, Vmax, VFEnext, Vsw"
I noted it from Air Caraïbes report, should be also in BEA #1 or #2, but it is of course somewhere else in the FCOM as I did read it more than once.
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PJ2,
Thanks for the paper you posted on CG. It answered my question/questions.
I also want to say to all the posters in the past day on this and related subjects,
It has been most educational to me, thanks
Thanks for the paper you posted on CG. It answered my question/questions.
I also want to say to all the posters in the past day on this and related subjects,
It has been most educational to me, thanks
A33Zab, #254
Thanks again for the illustration. From that, I assume from that pitch
and roll are shown yellow and blue gear train and that primary computer
pots are red, green and blue, while secondaries are red and blue only.
If this is correct, there is in fact no redundancy at hardware level in
terms of (ideal?) multiple pots per axis, per computer. For this configuration,
redundancy can only be done in software, between computers, using
messaging. I guess it makes sense. It reduces the overall number of
pots, still provides the required redundancy, but potentially introduces
a far more complex set of failure modes...
Thanks again for the illustration. From that, I assume from that pitch
and roll are shown yellow and blue gear train and that primary computer
pots are red, green and blue, while secondaries are red and blue only.
If this is correct, there is in fact no redundancy at hardware level in
terms of (ideal?) multiple pots per axis, per computer. For this configuration,
redundancy can only be done in software, between computers, using
messaging. I guess it makes sense. It reduces the overall number of
pots, still provides the required redundancy, but potentially introduces
a far more complex set of failure modes...
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HazelNuts39,
Concerning "SPD LIM RED FLAG", note that I'm also refering to those ACARS in the sequence:
2:10:29 -.1/WRN/0210 228300 2 06 > FLAG ON CAPT PFD SPD LIMIT
2:10:41 -.1/WRN/0210 228301 2 06 > FLAG ON F/O PFD SPD LIMIT
Those two FLAGS, despite the lag in recieving time, will be displayed on the PFD (and all characteristic speeds removed) as soon AP OFF + ALT2 switch (immediately or up to 10 seconds after AP OFF).
In fact, out of 15 ACARS time stamped 0210:
- 10 are Real Time ECAM/FLAGS
- 1 is a Real Time Position Report (kicking between those two SPD LIMIT flags above)
- 2 are System Maintenance Advisory (not displayed in flight)
- 2 are the related Faults to the above sequence which are delayed by the opening of the 1 min correlation window.
Now, if all of them kicked at the exact same time (0210:05, but it is not the case), they will be queued and it will take for the SATCOM nearly the same ammount of time up to the last fault recieved at 0211:55. The difference is only a small lag of 8-10 seconds between two FLAGs that were displayed at the same time in F/O, and CAPT PFDs... hence, this small lag is due to satellite's bandwith rather than system lag.
The processing time was nearly constant at 6 seconds +/- 1 second per ACARS, consequently, only one more message could have been sent during the same time window (01'50'') without such little lag.
Concerning "SPD LIM RED FLAG", note that I'm also refering to those ACARS in the sequence:
2:10:29 -.1/WRN/0210 228300 2 06 > FLAG ON CAPT PFD SPD LIMIT
2:10:41 -.1/WRN/0210 228301 2 06 > FLAG ON F/O PFD SPD LIMIT
Those two FLAGS, despite the lag in recieving time, will be displayed on the PFD (and all characteristic speeds removed) as soon AP OFF + ALT2 switch (immediately or up to 10 seconds after AP OFF).
In fact, out of 15 ACARS time stamped 0210:
- 10 are Real Time ECAM/FLAGS
- 1 is a Real Time Position Report (kicking between those two SPD LIMIT flags above)
- 2 are System Maintenance Advisory (not displayed in flight)
- 2 are the related Faults to the above sequence which are delayed by the opening of the 1 min correlation window.
Now, if all of them kicked at the exact same time (0210:05, but it is not the case), they will be queued and it will take for the SATCOM nearly the same ammount of time up to the last fault recieved at 0211:55. The difference is only a small lag of 8-10 seconds between two FLAGs that were displayed at the same time in F/O, and CAPT PFDs... hence, this small lag is due to satellite's bandwith rather than system lag.
The processing time was nearly constant at 6 seconds +/- 1 second per ACARS, consequently, only one more message could have been sent during the same time window (01'50'') without such little lag.