Tailplane lift

[DB6]

Talking recently with someone at work about tailplane lift. He reckoned that all aircraft (conventional) had tailplanes that exerted a downforce/negative lift. I had thought that some did, some didn't i.e. C of G could be either ahead of or aft of the mainplane.
He also thought that the tailplane would stall first to provide a nose drop at the stall. however I thought this was because of the C of lift moving back as the stall progresses, tailplane will be unstalled and fully useable during the stall.
I would welcome informed opinions.
(I know I could go and look it up, but the discussions are normally quite illuminating on PPRuNe, and I often learn more).

[fantom]

I was told by a Typhoon test pilot (what does he know?) that only the Typhoon tail produces lift.

What I do know is that my wonderful A330 stuffs 5 tonnes of fuel into the tail for the cruise to force the tail to produce less down-force (I didn't say 'lift'). This reduces the 'apparent all-up weight'. That reduces drag.

So, aft CG reduces drag.

It's got nothing to do with stalling.

QED

[italia458]

I don't have any data to back this up but my understanding is that to provide positive static and/or dynamic stability, most aircraft incorporate a design that has a tail downforce. I think generally that it wouldn't matter where the CG was located (either at the aft or forward limit) there would still be downforce on the tail. With an aft CG, however, there will be less downforce required to balance the aircraft and therefore you will have better cruise performance. Basically, the tail downforce acts like weight and has to be compensated for by lift (or a vertical component of thrust). To do that, you need to increase your angle of attack to create more lift which also creates more drag and slows you down. It's generally better to have a forward CG for recovery from unusual attitudes, ie: stalls or spins, and better to have an aft CG for cruise performance.

As you increase your angle of attack on the wings, your angle of attack on the tailplane decreases. You'll notice that as you decrease your speed towards the stall, you'll require more back elevator - it will be essentially an exponential increase in elevator deflection. This is partly to do with the decrease in speed (lift being a function of velocity squared) and partly to do with the decrease in angle of attack of the tailplane.

It is true that if you stall the tailplane, the nose will pitch down. With any stall, there is generally a buffeting action that occurs immediately prior to the stall. In a normal wing stall, you will feel the wings buffeting the whole aircraft. In a tailplane stall, the buffeting will mostly be felt through the control column. I would have to disagree that the tailplane is stalling first, causing the nose to drop. Of all the planes that I've stalled, I've never gotten the tailplane to even buffet. As a note, the tailplane stall recovery is completely opposite. To recover, you must pull back on the elevator assertively to break the stall.

So, what causes the nose down pitching moment at the point of stall? A lot of stuff.
1) The center of pressure will move backwards at the point of stall.
2) The amount of lift sharply changes - some people believe that lift drops off completely at the point of stall and that's why what happens at a stall happens. But, looking at a Cl vs AoA graph, you can see that that is not the case.

Take a look at this graph: http://classicairshows.com/Education/Aerodynamics/AeroDynamicsImages/Airfoil2CL.gif

You can see that at the Clmax, the lift drops off sharply - but it's still making TONS of lift! This airfoil stalls at what looks like 22 degrees AoA. Even at 8 degrees beyond the Clmax, it's creating lots of lift. But that's not analyzing the whole picture and that's why you initially don't think that lift still is high.

http://i.imgur.com/cKpPK.png

I've drawn a red line where, if you were able to continually increase your AoA and maintain the same lift, the Cl vs AoA line would follow. Essentially, about 1 degree beyond the critical AoA, the green line shows how much lift you've lost. So instead of being at 23 degrees AoA with a coefficient of lift of 1.8, you're stalled with a coefficient of lift of 1.15. At 1 degree beyond the critical AoA, you've lost 36% of the lift you should have at that AoA. The blue line shows roughly 6 degrees beyond the critical AoA. Not only are you losing a significant amount of lift, looking at the drag line you can see that at the critical AoA the drag rises sharply and continues rising at a high rate beyond the critical AoA. So, at the point of stall you're losing a significant amount of lift and, with the significant increase in drag, you're slowing down quickly - this is how you 'fall out of the sky'!

Going back to my points...
3) When you stall, your wings are essentially useless at this moment (not quite, but work with me here!) and you'd actually be better off without them. So, imagine that you're flying at a high angle of attack in level flight and your wings vanish. Just strictly based on the CG of the airplane, it will want to point the nose down to the ground. Not only that, but you've got that tailplane still there and if the CG was completely neutral, it would be the one to make sure that the nose pointed directly at the earth. It's now a missile heading directly for the earth. But before it can get to that point, the wings reappear and create lift effectively, ensuring that you don't nose dive the earth. Going back to the previous picture, you can see that a very small AoA increase at the Clmax point can have a drastic effect on the lift and drag. The same can be said for a very small AoA decrease at a point beyond the Clmax. So just as soon as your nose starts dropping down towards the earth like a missile, the wings come back to work and ensure that you don't dive straight down! You'll notice that the deeper the stall you enter, the more nose down the airplane will go and the above analogy should explain why.

Not all airplanes will handle exactly like that but for the majority of airplanes, that's pretty much the way she works - as I understand it.

Hopefully I answered your questions. If you have more let me know!

If you find some inaccuracies here also let me know.

[italia458]

I was told by a Typhoon test pilot (what does he know?) that only the Typhoon tail produces lift.

Not surprised. But I would also be surprised if an airplane like the F-22 didn't produce upwards lift on its tail. Generally, the more maneuverable an airplane, the more unstable. And if I ever did see a maneuverable airplane, the F-22 is one! It's amazing what fly-by-wire technology can do these days.

There are some difficulties with creating upwards lift on the tail. I'd be interested to hear from anyone who's familiar with the aerodynamics of an airplane with that type of configuration.

Edit: It's got nothing to do with stalling.

QED

It has to do with QED? - Quantum Electrodynamics?! :8

[FE Hoppy]

I spent a great deal of my youth making free flight balsa models. I can assure you that an aircraft can be perfectly stable producing lift from the tail.

It's all about the moments baby!

[Clandestino]

Theorethicaly, it would be possible to have stable aeroplane with Cp ahead of Cg and tailplane generating upforce to balance wings' pitch-up momentum, provided every change of AoA leads to greater change of pitching momentum produced by tailplane than wings'. Whether such design can really be made (or was ever made), I have no idea.

On conventional design with tailplane producing downforce, its stall leads to rapid pitchdown with vertical dive easily and quickly attained, so it has to be designed to be immune to stall. On some specific designs, heavy ice accretion on tailplane's leading edge can compromise this. Non-FBW canards, such as VariEze, absolutely must have foreplanes that stall before mainplane, otherwise canards still producing lift combined with stalled wing would produce further pitchup - not quite desired effect when stalling.

Moving the CG aft by transferring fuel into tailplane reduces pitch-down momentum of the wings by reducing its arm. With lesser downforce from tail, less lift is required to balance it (decreasing a bit the stall speed), also trim drag is decreased. Downside is with shorter Cp to Cg arm, restoring pitch momentum with change of AoA is also smaller, making the aeroplane less stable - that's why it is desirable to empty trim tanks before approach.

[bookworm]

Talking recently with someone at work about tailplane lift. He reckoned that all aircraft (conventional) had tailplanes that exerted a downforce/negative lift. I had thought that some did, some didn't i.e. C of G could be either ahead of or aft of the mainplane.

For stability in angle of attack, it's only necessary that the proportion change in lift from the tailplane exceeds the proportion change in lift from the mainplane for a small change in AoA at both. If you imagine identical symmetric aerofoils for the mainplane and the tailplane, that criterion only requires that the tailplane is at a lower AoA than the mainplane, not that it is at a negative AoA.

For the same loading, the difference in AoA will be smallest for low AoA (high speed) cruise -- after all, you pull back (increase the tailplane lift in a downwards direction) to increase the AoA. But that means that the mainplane AoA is already at its lowest, and if the tailplane AoA has to be even lower, it's likely to be only just positive. I don't know if that ever occurs in practice.

[bubbers44]

I think only the cannard style planes can use positive lift on the tail. Stability is caused by the negative lift of the tail plane allowing the nose to drop when slowing down and up when speeding up. Yes, a rearward CG increases efficiency by requiring less negative lift of the tailplane. Aircraft use fuel transfer aft to accomplish this increasing fuel economy. I did it in the older Lear Jets 35 years ago. Fighter jets use computers to not require much or any tailplane negative lift.

[italia458]

I spent a great deal of my youth making free flight balsa models. I can assure you that an aircraft can be perfectly stable producing lift from the tail.

It's all about the moments baby!

I agree, but your basla models weren't carrying lots of cargo at varying CG positions! But I couldn't help but think it'd be hard to be sure that your tailplane was actually creating positive lift unless you had a windtunnel! I could dig up some resources on this and find a solid answer as to why they are generally designed to have the tailplane producing downwards lift but it essentially comes down to safety, utility and maneuverability.

[Mechta]

FEHoppy is right. Several classes of free flight competition models routinely use lifting section tails and have their centre of gravity around 75% chord, or in the case of the F1D class microfilm model below somewhere between the front and rear surfaces.


When the wing loading is very low, the induced drag from a lifting tail is minimal too. The QinetiQ Zephyr solar powered UAV used this to good effect as well. It should be borne that this and most models that use a lifting tail are virtually single speed aircraft.

Tandem wing aircraft such as the Henri Mignet's Flying Flea also generate lift from both surfaces, and if you know anything about that, you will know it can cause problems, as the forward wing at low angles of attack will act as a slot for the rear wing.

[Dan Winterland]

It's true to say that for a conventionally controlled aircraft (i.e. not FBW) the tailplane should generate proportionally less lift than the wing to maintain positive longitudinal stability. This can be acheived by a different aerofoil section, lower angle of attack or both. Often, the tailplane is designed to produce a downforce which makes for a very stable design.

[Fostex]

Some interesting footage of what happens when you remove the tail plane of an aircraft in flight. This was during the test flights in the lead up to the Dambusters raids.

Bouncing Bomb Boo Boo - YouTube

[Dan Winterland]

I think it's safe to say that the tailplane on that aircraft was generating a downforce!


That was the US trial of the "Highball" bouncing bomb desined for use against ships. It was smaller and more spherical than the "Upkeep" weapon used against the dams. Designed to be carried by the Mosquito -or similar size it was being trialled for use in the Pacific war. Some were sent out to the Pacific, but the war ended before they were used.

[DaveReidUK]

What I do know is that my wonderful A330 stuffs 5 tonnes of fuel into the tail for the cruise to force the tail to produce less down-force (I didn't say 'lift').

Your tail isn't producing any less down-force, it's just that more of it is the result of gravity and less from aerodynamics (hence the reduced drag).

[italia458]

I'd have to agree with the A330 quote - the tail would be producing less downforce. The tail downforce acts the same way as weight does - down and needs to be compensated for by lift. So increasing the angle of attack on the wings will create more lift but also more drag which slows the plane down. This explains why more weight or a forward CG will decrease cruise performance.

I think it's safe to say that the tailplane on that aircraft was generating a downforce!

I completely agree - but I'd also say that the CG has changed drastically! I would say both conditions contributed to the accident in this case.

[rudderrudderrat]

Hi italia458,
This explains why ... or an aft CG will decrease cruise performance.
Surely an aft CG helps improve cruise performance?

[Microburst2002]

Stability is one thing and equilibrium is another.

Negative lift at the tailplane is a requirement for the latter. It is due to the wing aerodinamic moment, which is nose down for positive lift flight. Tailplane provides nose up with negative lift, thus balancing moments.

[Mechta]

Martin Simons 'Model Aircraft Aerodynamics' 1994 edition has a good explanation on pages 163 to 165 of zero lift tails and lifting tails. If you can't be bothered to go and find where you've put your own copy, there is a pdf version on 4shared...:eek:

[italia458]

Surely an aft CG helps improve cruise performance?

It sure does! Thanks, I made the correction.

Stability is one thing and equilibrium is another.

Negative lift at the tailplane is a requirement for the latter. It is due to the wing aerodinamic moment, which is nose down for positive lift flight. Tailplane provides nose up with negative lift, thus balancing moments.

It also improves stability. An airplane disturbed and put into a dive will pick up speed. As it picks up speed (if the tailplane is producing downforce) the wings will decrease AoA and the tail will increase AoA which is what causes the pitch up to restore the airplane towards it's original path. Elevator trim doesn't trim for airspeed, it trims for angle of attack. And when you disturb the airplane from it's trimmed AoA, it will try to return to that AoA. Since the trim is trying to maintain a certain AoA, the wings will also try to return to that AoA and, in doing so, will produce more lift which will also help bring it back to it's original path.

[Microburst2002]

Aircraft stability is determined by the relative location of ac and cg, irrespective of wether tailplane lift is positive or negative.

For equilibrium, cp location with regard to cg is what matters. For an equilibrium condition, total airplane's lift moment about cg must balance any other pitching moments (due to drag an thrust).

Normally drag line of action is below cg, so it creates nose up moment. More often than not, so does thrust (good characeristic in case ong engine failure).

Wing lift acts somewhere behind the 25% chord, depending on the aoa (the less the angle, the further aft the cp is). It normally creates a pitch down moment, but it can be pitch up at high aoa, with an aft cg. Probably this is what they do in those models?

Tailplane lift must balance the net pitching moment of the other forces, plus the pitch down aerodynamic moment of the wing. For normal flight conditions in typical transport airplanes, it is a downwards force which must be balanced by wing lift just as if it was weight.

An aft cg means taht the wing lift nose down created moment will be less, so that the nose up moment required from the tail is less, so that less part of wing lift is wasted in balancing "apparent weight".

the wings will decrease AoA and the tail will increase AoA...

If wing aoa increases, so does tailplane's. Can't be otherwise.

[rudderrudderrat]

Hi Microburst2002,
If wing aoa increases, so does tailplane's. Can't be otherwise. It can be so if the tail plane is at a negative incidence (producing a down force).
That's why normal commercial aircraft are able to be trimmed angle of attack stable.

[italia458]

Microburst... there is a lot more that contributes to airplane stability than just AC (aerodynamic center) and CG! Here is a tiny excerpt from 'Aerodynamics for Naval Aviators' that talks a little about the effects of the horizontal tail.

http://i.imgur.com/MuQSF.png

You said AC at the beginning but then started talking about CP (center of pressure). It's generally the AC that's used when working with stability. The AC is the point where the wing pitching moment coefficient doesn't vary with lift coefficient. At subsonic speeds it generally remains stationary at 25% chord. "All changes in lift coefficient effectively take place at the wing aerodynamic center. Thus, if the wing experiences some change in lift coefficient, the pitching moment created will be a direct function of the relative location of the AC and CG." - Aerodynamics for Naval Aviators

If wing aoa increases, so does tailplane's. Can't be otherwise.

I don't believe so. When the tailplane is producing a downforce, which in most airplanes it is, when you increase your AoA on the wing, the tailplane's AoA will decrease. The best way to see this is to visualize the airplane in level flight, slowing down, and increasing its wing's AoA. As the airplane is pitching up, the air striking the tailplane is hitting it more on the underside. This is reducing the AoA on the tailplane. Virtually everything is the opposite when it comes to a tailplane that produces downward lift.

You might have heard that, in icing conditions, putting flaps down can really aggravate the condition and possibly enter the airplane into a tailplane stall. That's because it's deflecting the air downwards, even more, right before the tail. The more downwards air from the wing, the higher the AoA on the tailplane. It's also why the tailplane stall recovery procedure is to pull back on the yoke and pretty much undo what you just did - ie: take the flaps up.

Edit: Microburst, you mentioned that stability and equilibrium are different. I do agree that they are different, but not separate. Equilibrium is a part of stability.

[BOAC]

Normally drag line of action is below cg, so it creates nose up moment. - try drawing that out on paper and maybe edit your post?When the tailplane is producing a downforce, which in most airplanes it is, when you increase your AoA on the wing, the tailplane's AoA will decrease. - i think it would help you and others to avoid MFS's 'back axle' effect if you defined your axes and values, since we are talking maths/physics here. Let's make it simple - define positive AoA for a flat plate section and then 'increasing' AoA in terms that will work in equations. Use 'traditional' a/c stability axes for simplicity? Then we can discuss raising the nose above the horizon while S&L - but inverted..........................

[pigboat]

An example of a tailplane that produces downward lift. The top of the horizontal stab of the F-27 is practically flat.

http://aircraftwalkaround.hobbyvista.com/f27/f27_40.jpg

[Microburst2002]

Italia

I must have written my posts very poorly when you are explaining to me what I have just said.

I know well what the ac is and that is precisely why I say that it is the position of the cg with respect to it that determines de degree of stability of the airplane. By the way, I have read, reread and even lovingly caressed that blessed book for many years...

There are many factor affecting stability, but once the airplane has been designed, it is the cg location what will make it more or less stable. Not the angle of incidence of its tailplane.

The conditions for equilibrium are one thing and the conditions for stability are a different thing. The cp is where we consider that Lift is acting. The moment about the cg created by Lift is determined from the cp. Unlike the ac, cp position varies with CL, being zero at zero lift and about 25% at maximun CL.

BOAC You are right, It was a lapse, Drag acts above the CG, of course. Thanks for that, it took me a while to spot it.

Italia

Of course, if we talk about negative AoA, it decreases when the wing's positive AoA increases, but the effect is always in the same sense as the wing. Less negative lift has the same effect as more positive lift, right?

The taiplane is also referred to as the stabilizer for obvious reasons, but it is not due to the sense in which it develops Lift. It stabilizes because when the airplane increases its AoA, the nose up moment of the tailplane about the cg will decrease (or the nose down moment will increase). The effect is the same: opposing to the pitch up. All due to its being behind the CG. A canard plane will always be unstabilising, no matter if its lift is positive or negative.

More nose down or less nose up, the stabilizing effect is the same. If stability depended on the angle of incidence of the tailplane, it would vary every single time we moved an all moving slab or a trimmable horizontal stabilizer. The truth is that with that we would just change is the estate of equilibrium.

BOAC, sorry, I am a sweet water sailor. Inverted flight belongs to emergency, for me...

[italia458]

Microburst...

Yes, that A for NA text is fantastic! I would love to study it cover to cover but I sadly don't have the time right now.

Of course, if we talk about negative AoA, it decreases when the wing's positive AoA increases, but the effect is always in the same sense as the wing. Less negative lift has the same effect as more positive lift, right?

I think this would actually be "negative lift" and not "negative AoA" to be correct? I could have a positively cambered horizontal stabilizer at a negative AoA that still produces positive lift. In that case, increasing the AoA on the wings would also increase the AoA on the horizontal stabilizer.

[Microburst2002]

Very true.

By the way, this whole subject has reminded me of an old confusion that I have regarding the aerodinamic moment.

I am not sure if it is the torque moment created by the wing along with lift and drag, or if it is a mere mathematical subproduct of considering lift as acting on the ac instead of on the cp.

I think I raised a thread on that subject but it didn't end well

[Owain Glyndwr]

The conditions for equilibrium are one thing and the conditions for stability are a different thing. The cp is where we consider that Lift is acting. The moment about the cg created by Lift is determined from the cp. Unlike the ac, cp position varies with CL, being zero at zero lift and about 25% at maximun CL.

By the way, this whole subject has reminded me of an old confusion that I have regarding the aerodinamic moment.

I am not sure if it is the torque moment created by the wing along with lift and drag, or if it is a mere mathematical subproduct of considering lift as acting on the ac instead of on the cp.

If I may chip in ....

At zero lift a normally cambered airfoil will have a negative (nose down) pitching moment which is a couple. This means that the cp at zero lift is out at infinity somewhere (not zero). As AoA is increased the additional lift can be taken to act at 25% chord. This is where it acts from theory and where it acts when measured in a wind tunnel. A symmetric airfoil will have its cp at 25% chord at all AoAs, even zero, because the pitching moment at zero lift is also zero. Cambered airfoils will have their cps varying with AoA but tending towards 25% chord at high CL as you say.

So the aerodynamic moment at any CL is created by the wing as a combination of lift and zero lift pitching moment. For cambered airfoils the associated cp will be somewhere aft of 25% chord. The aerodynamic centre is defined as the point at which changes in lift (as a consequence of AoA changes) act. This, for low speed airfoils at least, is 25% chord.

As you say, the conditions for equilibrium and stability are different things. For equilibrium (trim) we need to consider the moment arm from the CG to the cp; for stability we need to consider changes from the trimmed state so we are interested in the moment arm from CG to aerodynamic centre.

When AoA increases the incidence on both wing and stabiliser increases, but the wing generates an increased downwash on the tail so the tail AoA increases by less than the wing AoA - think in terms of 3 deg at the tail for 5 deg on the wing.

It should be clear from this that with the CG ahead of the cp, you will need downward lift on the tail to get equilibrium. Moving the CG aft will reduce the lift required until eventually, if you bring the CG aft of the cp, you can trim with upwards (positive) lift on the tail which then helps to offload the wing (for a given weight) and reduce drag.

BUT, moving the CG aft will also reduce the moment arm to the ac and thus reduce stability. Without going into theory, although the wing ac is at 25% chord, the aircraft (wing plus tail) ac is somewhere around 55% chord and this is what matters. So if the cp is at, say 30% chord you can fly with a CG at say 32%, get the trim drag benefits and still have a stable aircraft. You don't need FBW to get this, but it helps.

[italia458]

I think this section of A for NA will be helpful.

https://www.box.com/s/0a1746cf52f6fec42be0

[bubbers44]

Yes that is true but the wing is positive lift and the tailplane is negative lift or down force. Looking at it your way, It can be confusing for the people that think conventional planes have positive lift on the tailplane. It is down or negative lift. The increased airspeed and slight increase in negative AOA makes conventional AC stable.

[BOAC]

Indeed, bubbers. Hence my suggestion that to avoid the back axle effect italia defined his/her "axes and values"

[italia458]

Indeed, bubbers. Hence my suggestion that to avoid the back axle effect italia defined his/her "axes and values"

BOAC... Not being disrespectful, but I honestly don't know why you think I wasn't clear in defining up and down. I could have said something like down is defined as being towards the center of the earth when the airplane is in an attitude that is parallel to the surface of the earth, but I thought it was clear enough without some complicated definition.

[BOAC]

.and 'increasing AoA'? In which direction is that? In which direction is 'positive lift' for a wing at 90 degrees to the earth's horizontal, or inverted? What is 'increasing AoA' for a section with a negative AoA - increasing negative or increasing positive? It does matter.

[italia458]

Fair enough.. I thought it was easy to follow what I was saying but I see what you're saying. In the cases I was talking about, 'increasing AoA' meant that if it was positive, it was getting more positive and if it was negative, it was getting more negative - in both cases, the angle is getting bigger relative to itself. Since I was talking about a normal airplane that is right-side-up, I assumed wing lift to be acting upwards away from the earth and tailplane lift acting downwards towards the earth.

[TURIN]

When AoA increases the incidence on both wing and stabiliser increases,

I do not understand how this statement can be true of a 'conventional', stable, aircraft design.


Please indulge me for a moment while I work this through as I know from previous experience that most Ppruners are a damn sight cleverer than me!:O

Assuming that the mainplane is approx horizontal to the horizon and is at a positive AoA, IE the L/edge is higher than the T/edge relative to airflow then it will produce positive lift. IE UP. The tailplane is also approx horizontal but with the L/edge lower than the T/edge it is still at a positive AoA relative to itself. It will be producing negative lift. IE DOWN.

Therefore, if the a/c nose rises relative to the airflow, the mainplane AoA will increase and the tailplane AoA will decrease.

The result will be that the nose drops because the tailplane is producing LESS downforce (negative lift). After a cycle or two steady state level flight is restored.

I am assuming that the CofG is forward of the CofP.

I know there are a numerous other factors to consider but in the ABC of aerodynamics for idiots (such as myself) this is correct. Yes? :confused:

[BOAC]

Turin - the essence is that a change from a certain negative AoA to a lesser negative AoA is, de facto, an increase in AoA towards a positive val;ue in real terms. That is why it is important to understand our 'framework'. It is basic mathematics. Think for one minute how you would describe the situation IF the AoA of the tail 'changed' so as to be 'positive' (ie nose above tail) to the airflow, and then the a/c pitched further nose up - would you still maintain the tailplane AoA decreased? You would find a mathematical approach very difficult if you did. Every angle and movement/rotation must have a reference to the a/c frame. EG Start at -10 degrees and finish at -5 degrees - increase or decrease? Once you start saying the negative angle 'decreases' you are in 'double negative land'. Zero is a problematical number!

[TURIN]

Think for one minute how you would describe the situation IF the AoA of the tail 'changed' so as to be 'positive' (ie nose above tail) to the airflow, and then the a/c pitched further nose up - would you still maintain the tailplane AoA decreased?

No I wouldn't.

AofA (I thought) is the angle between chord line and relative airflow.

If the angle between chord and airflow reduces then AofA reduces and vice-versa.

If the angle reduces it cannot be an increase in AofA.

I can see how mathematically it is important to define negative and positive to explain or calculate effect but for descriptive purposes we must at least try and use a common language.

Beer is usually an excellent translator I find. :ok:

[Microburst2002]

I feel guilty...

But Let's go on with the subject of camber and cp and why symmetrical airfoils always have cp in 25% (always speaking subsonic).

What is different in the circulation of airflow created by AoA with respect to that created by camber?

There is a relation Camber-circulation-Aerodynamic moment that I don't quite understand.

[Owain Glyndwr]

Turin
I can see how mathematically it is important to define negative and positive to explain or calculate effect but for descriptive purposes we must at least try and use a common language. Most people find mathematics a very useful common language in these circumstances. It introduces a discipline that avoids the sort of confusion your approach leads to.:=

Microburst

What is different in the circulation of airflow created by AoA with respect to that created by camber?Not really sure how to reply to that one. Short answer is that there is no real difference in that the variation of lift with changes in AoA is the same on a cambered airfoil as it is on a symmetrical one with the same planform. What camber does is change the datum point from which these changes originate so that (for example) zero lift on the old aeromodeller's favourite section the Clark Y occurs at -3.35 degrees AoA. Lift coefficient is then:

lift curve slope*(AoA - AoA zero lift)
or for the Clark Y:
lift curve slope*(AoA + 3.35)

http://i1081.photobucket.com/albums/j351/OwainGlyndwr/ClarkYzerolift.jpg

There is a relation Camber-circulation-Aerodynamic moment that I don't quite understand. The other major difference is that at zero lift on a cambered section there is some downward lift on the front bit of the chord and upward lift on the rear bits. The overall lift is zero, but the two components combine to give a nose down pitching moment (zero lift pitching moment, or Cmo) which is a couple since there is no net force involved. Once you move away from the zero lift AoA, i.e. -3.35 degrees in this case, the additional lift acts at 25% chord just like it does on an uncambered section. The cp will then be Cmo/CL chord lengths away from the 25% chord point.

Any better?

PS: Sorry about the scruffy shading - in a bit of a hurry. Also I seem to have lost the bit that says the solid line is upper surface pressure and the dashed line the lower

[TURIN]

"Most people find mathematics a very useful common language in these circumstances. It introduces a discipline that avoids the sort of confusion etc etc. "

I have to disagree. Most people use the spoken word to explain anything. Mathematics confuses almost everybody at one level or another. See above. :)

[Microburst2002]

Owain

Thanks for your reply.

After a while the graph seems to have some sense. Lately I have lost my ability to interpret graphs. I was good at that some time ago, but I'm growing old... What is on the vertical axis, exactly? Anyway I think I understand what you mean. Comes to my mind a graph with those lobes of pressure in the upper and lower surfaces.

So symmetrical airfoils never create any couple, but cambered ones always do, Right?. I will accept that as an empirical fact, because a theoretical explanation of it can be extremely complicated, I deem.

Can we consider the effect of the airstream on the wing as producing a total reaction plus a pure moment (a couple) that we call Aerodynamic Moment?

In case this is affirmative, the point at which the Lift is considered to be acting (i.e. the CP)... Does it take into account that couple, too? or you have to consider the moment of Lift (acting on the CP) about the CG and then add "the couple" or aerodynamic moment to find the total pitching moment effect?

PD
I think that maths are a kind of language but it is is good to frequently translate it into normal words

[Owain Glyndwr]

Microburst

but I'm growing old...

Know the feeling:ok:

What is on the vertical axis, exactly? Anyway I think I understand what you mean. Comes to my mind a graph with those lobes of pressure in the upper and lower surfaces.

Vertical axis is pressure difference relative to ambient pressure divided by freestream dynamic pressure - negative values are suctions, positive values are pressure above ambient. So the +1.0 is where the local pressure equals stagnation pressure. It is in fact one of those graphs with lobes of pressure on upper and lower surfaces. You just have to be aware that at negative AoA the pressure on the lower surface LE is the one that is negative.

So symmetrical airfoils never create any couple, but cambered ones always do, Right?. I will accept that as an empirical fact, because a theoretical explanation of it can be extremely complicated, I deem.

Yes, and I am darned glad you don't want a theoretical explanation ;)

Can we consider the effect of the airstream on the wing as producing a total reaction plus a pure moment (a couple)

Yes

that we call Aerodynamic Moment?

Maybe I've misread your intent, but aerodynamic moment is usually used for the combined effects of reaction (lift) and couple, referred to some chosen reference point e.g. 25% chord or aircraft CG rather than just the couple.

In case this is affirmative, the point at which the Lift is considered to be acting (i.e. the CP)... Does it take into account that couple, too? or you have to consider the moment of Lift (acting on the CP) about the CG and then add "the couple" or aerodynamic moment to find the total pitching moment effect?

The point at which the lift is considered to act (cp) takes the couple into account, which is why the cp is calculated as Cmo/CL chords away from the 25% chord point (assuming that to be the chosen reference).

I think that maths are a kind of language but it is is good to frequently translate it into normal words

Agreed, but I think you need to retain the discipline of mathematical conventions in your "normal" wording.

[Microburst2002]

I think I start to understand

The moment exists physically, which I was not sure. I mean, if the wing was made of paper or something frail it would be bent and maybe even folded or broken because of the couple. And its effect is all included in the cp.

Now, regarding the tail lift as being negative or positive, I had the notion that it had to be always negative because of the aerodynamic moment, and that a tailplane was necessary (an then the germans invented the flying wing)

But what I think is just that for the airplane to be in trim, the tailplane lift has to able to balance all the other pitching moments, whichever they are. These are basically those due to the wing plus fuselage lift at the wing-fuselage cp, the thrust and the drag.

The Drag-Thrust couple usually gives a nose up (positive?) moment, and the Lift gives a nose down (negative?). The tail lift has to balance the resultant of those moments. I guess that the Lift effect is much greater than the T-D couple? It seems so if Lift is many times bigger than Trust for Drag, but it depends on where the wing-fuselage cp lies. I deem it does lie well aft of the CG?. That would be why the tailplane has to produce a pitch up (positive?) moment with negative lift. But can the cp be ahead of the CG for "normal" angles of attack?

[TURIN]

To answer your last point. Yes, but the a/c would not be naturally stable.

[Owain Glyndwr]

Now, regarding the tail lift as being negative or positive, I had the notion that it had to be always negative because of the aerodynamic moment, and that a tailplane was necessary (an then the germans invented the flying wing)

Hmmm ... but on a certain very fast tailless aircraft the wing/body cp and cruise CG (with fuel transfer) were arranged so that the elevons were deflected slightly TE down (positive lift) to keep the drag down.

But what I think is just that for the airplane to be in trim, the tailplane lift has to able to balance all the other pitching moments, whichever they are. These are basically those due to the wing plus fuselage lift at the wing-fuselage cp, the thrust and the drag.

The Drag-Thrust couple usually gives a nose up (positive?) moment, and the Lift gives a nose down (negative?). The tail lift has to balance the resultant of those moments. I guess that the Lift effect is much greater than the T-D couple? It seems so if Lift is many times bigger than Trust for Drag, but it depends on where the wing-fuselage cp lies. I deem it does lie well aft of the CG?. That would be why the tailplane has to produce a pitch up (positive?) moment with negative lift. But can the cp be ahead of the CG for "normal" angles of attack?

Maybe if we put in some numbers? these are not specific to any one type, but they are realistic.

Modern airliner wings have the section (camber and thickness distributions) and twist (local AoA variations) carefully tailored to maximise performance (NOT to minimise drag since wing loads/weight come into the equation). In addition their planform is usually swept and has a compound taper. Consequently it is very difficult to pick a 'real' chord to represent things and it is usual to calculate a 'mean aerodynamic chord' (mac or amc) based on planform and use that as a reference 'axis'.

The wing body CP in cruise conditions varies between 30~40% mac depending on lift coefficient and Mach number, with the higher end corresponding to the maximum cruise Mach (Mmo). If you back off to best economic cruise conditions then the CP will be around 35% mac.

On an aircraft without fuel tanks in the tail the CG limits will range from 15~20% mac at the forward end to about 30~35% mac aft. If the aircraft has a tail tank then the aft limit will stretch to 40% mac, maybe a touch more.

Just glancing at those values you can see that in most cases the CP will be behind the CG and you will need negative tail lift to trim, but if you have and use the tail fuel tank then the CG can be at or behind the wing/body CP and the tail lift will be either very small or positive. Factor in the nose-up pitch from the drag/thrust couple and you will definitely have positive tail lift to trim.

Just to round things off, the complete aircraft (tail on) aerodynamic centre will be around 50~55% mac so the aircraft would be stable even at the aft CG limit (as it must be of course to meet regulations)

[XPMorten]

Maybe also mention that on most low tail aircraft, there will be downwash from the wing
hitting the stabilizer creating a downforce. The higher airspeed, the stronger the downwash
and more downforce.

[cwatters]

I used to fly competition RC model aircraft. The class rules specified a maximium wing loading limit so there was an incentive to build them light so they were smaller. On one occasion a competitor appeard to have taken this a bit too far when the tail boom broke on a straight and level but very high speed run. Due to the nature of the break it appeared the down load on the tail was the cause, probably compounded by damage that occured on a previous flight.

[Owain Glyndwr]

Maybe also mention that on most low tail aircraft, there will be downwash from the wing
hitting the stabilizer creating a downforce. The higher airspeed, the stronger the downwash
and more downforce.Not exactly true - the higher the airspeed the lower the wing AoA and the lower the downwash angle at the tail (which is usually about 40% of the wing AoA)
Plus of course the fact that downwash doesn't change the force on the tail needed to balance the airplane - it just changes the tail/body setting necessary to generate that force.

[TURIN]

My understanding of tail (trim tank) fuel is to reduce the amount of trim and therefore drag, not to swap pitch up trim for pitch down trim. The a/c still needs to be stable. Especially air transport category, to comply with regulations. Think about failure modes. If engines are lost with the a/c in non normal trim, hand flying, whether it be fly by wire in direct law or traditional control systems, will be very tough for an average pilot. Military a/c performance is a completely different kettle of fish.

[XPMorten]

OG,

Not exactly true - the higher the airspeed the lower the wing AoA and the lower the downwash angle at the tail (which is usually about 40% of the wing AoA)
Plus of course the fact that downwash doesn't change the force on the tail needed to balance the airplane - it just changes the tail/body setting necessary to generate that force.

Pilot's Handbook of Aeronautical Knowledge Chapter 3 - American Flyers (http://www.americanflyers.net/aviationlibrary/pilots_handbook/chapter_3.htm)

Even though the horizontal stabilizer may be level when the airplane is in level flight, there is a downwash of air from the wings. This downwash strikes the top of the stabilizer and produces a downward pressure, which at a certain speed will be just enough to balance the “lever.” The faster the airplane is flying, the greater this downwash and the greater the downward force on the horizontal stabilizer (except “T” tails). [Figure 3-13] In air-planes with fixed position horizontal stabilizers, the airplane manufacturer sets the stabilizer at an angle that will provide the best stability (or balance) during flight at the design cruising speed and power setting. [Figure 3-14]

If the airplane’s speed decreases, the speed of the air-flow over the wing is decreased. As a result of this decreased flow of air over the wing, the downwash is reduced, causing a lesser downward force on the horizontal stabilizer. In turn, the characteristic nose heaviness is accentuated, causing the airplane’s nose to pitch down more. This places the airplane in a nose-low attitude, lessening the wing’s angle of attack and drag and allowing the airspeed to increase. As the airplane continues in the nose-low attitude and its speed increases, the downward force on the horizontal stabilizer is once again increased.

fig 3.13 and 3.14
http://www.americanflyers.net/aviationlibrary/pilots_handbook/images/chapter_3-2_img_6.jpg

[Owain Glyndwr]

My understanding of tail (trim tank) fuel is to reduce the amount of trim and therefore drag, not to swap pitch up trim for pitch down trim. Sure, tail fuel is used to reduce trimmed drag, but do you really understand how it does so?
Negative lift (pitch up trim) means that to balance the overall lift/weight equation the wing must produce more lift than just the aircraft weight to offset the negative tail contribution and that means more drag due to lift. If you move the CG aft so that the amount of negative tail lift is reduced then you will reduce the wing lift requirement and get a lower trimmed drag. Going further to the point where the trim is pitch down (positive tail lift) is merely an extension of that process. So swapping pitch up trim by pitch down trim is exactly what you need to do to reduce trimmed drag.

The a/c still needs to be stable. Especially air transport category, to comply with regulations. Think about failure modes.Read again what I said a couple of posts ago:
QUOTE]Just to round things off, the complete aircraft (tail on) aerodynamic centre will be around 50~55% mac so the aircraft would be stable even at the aft CG limit (as it must be of course to meet regulations) [/QUOTE]

Not much to be doing with failure modes either - if the CG is ahead of the aerodynamic centre the aircraft will be statically stable whatever the FBW is doing.

If engines are lost with the a/c in non normal trim, hand flying, whether it be fly by wire in direct law or traditional control systems, will be very tough for an average pilot.This has nothing to do with engine failure (except if you are fool enough to take off with full aft CG and get an engine failure near Vmca, in which case the lateral control might be a problem, but even then the aircraft has to be certificated for engine failure at aft CG). And if the aircraft is certificated for flight with an aft CG then that is NOT non normal trim. Moreover it has to be flown without difficulty in that state by an average pilot - that is part of the airworthiness code.

[Owain Glyndwr]

XPM

Even though the horizontal stabilizer may be level when the airplane is in level flight, there is a downwash of air from the wings. This downwash strikes the top of the stabilizer and produces a downward pressure, which at a certain speed will be just enough to balance the “lever.” The faster the airplane is flying, the greater this downwash and the greater the downward force on the horizontal stabilizer

Hmm - different definition of 'downwash' is causing a problem. In this FAA document they seem to be using downwash to describe a force where I am using it to describe an angle of flow.

Consider for example this extract from Richard Shevell's "Fundamentals of Flight"
The downwash at the tail comes almost entirely from the trailing vortex system of the wing ...... In practice one wing semispan behind the wing qualifies as "far behind", so that 2CL/(Pi*A.R) gives a rather good approximation to the downwash at the tail. The downwash actually varies both with distance behind the wing and height of the tail above the wake.

In that expression CL is the lift coefficient and AR is the wing aspect ratio. I hope we can agree that if airspeed is increased the lift coefficient is reduced, so that downwash, to my definition, would be decreased. This is just the opposite of what the FAA are saying, but Shevell (and I) are talking about the incident flow angle onto the tail not the force generated on the tail by the downwash, which varies with speed squared as well as flow angle.

I think that is the reason for our apparent difference of opinion :ok:

[XPMorten]

In that expression CL is the lift coefficient and AR is the wing aspect ratio. I hope we can agree that if airspeed is increased the lift coefficient is reduced, so that downwash, to my definition, would be decreased

Cl does decrease, but total lift force is constant regardless of airspeed (assuming level flight)
This means that to maintain lift at higher airspeeds (lower aoa and Cl), more air mass gets accelerated faster (vertically) giving
an increased downforce on the tail. F=m*a

[john_tullamarine]

An interesting discussion is developing.

For those who might think that tailplanes are not associated with downforce, a simple Google search for "friendly fire Caribou accident" will lead to a very well-known photo of a Caribou's last seconds after having the empennage shot off by friendly fire on short final into a forward strip in Vietnam during the mid-60s.

Once you have figured out the perspective, you will see that the aircraft has pitched nose down and through the vertical .. in the space of a second or two.

[Owain Glyndwr]

Cl does decrease, but total lift force is constant regardless of airspeed (assuming level flight) Agreed

This means that to maintain lift at higher airspeeds (lower aoa and Cl), more air mass gets accelerated faster (vertically) giving
an increased downforce on the tail. F=m*a Except that it doesn't work like that. Force is equal to the rate of change of (vertical) momentum; that is it equals the mass flow over the wing times the vertical velocity imparted to that air by its passage over the wing. [Just rewrite your F=m*a (or F=m*(dV/dt) to read F= (dm/dt)*deltaV - still the same equation but different emphasis]
As you point out, at a higher airspeed that is more mass flow and since the lift is constant that means that the vertical velocity when the air leaves the wing must be less. But that vertical velocity is the downwash ...

If we are just talking about the changes in force on the tail that result in a stable airplane, then yes, lower AoA goes with a bigger tail downforce because that is what pitches the airplane back towards its stable condition - I don't have a problem with that, but it it the change in tail AoA that comes from the reduction in aircraft AoA combined with the change in downwash angle that produces the additional downforce (at constant airspeed that is).

If we are talking about changes to the original trim coming from an increase in speed, then both wing and tail forces are going to be increased and unless the AoA is reduced the aircraft will develop some normal acceleration.

[Microburst2002]

Owain, thanks a lot

Now I undestand it very well. Some nubers help very much indeed.

Regarding FBW, with these airplanes it is difficult to talk about stability. I mean, they cannot be left either stick fixed nor stick free, to begin with. The system is always making inputs!

[Owain Glyndwr]

Yeah, it is difficult to see exactly what is meant by stick fixed/free stability these days - but then stick free stability hasn't meant much since aircraft have been fitted with hydraulically driven controls, since changes in control hinge moments due to aerodynamic changes don't get reflected back into stick position and thence control movements.;)

[Microburst2002]

The way I look at it, the most important requirement is controlability (pilot capability to control)

If the airplane was unstable it would not have controlability. If it was too stablebit would neither have controlability because of lack of manoeuvrability.

But if the airplane had a FBW flight control system such that it can always meet pilot demand no matter how unstable the basic airplane might be, what sets the lower limit of stability?

Wait, Now thinking I guess that certification requirements must account for failure of the FCS. Is that so? The aft CG limit of an A320 or a B777 is determined because of the direct law possibility?

[rudderrudderrat]

Hi Owain Glyndwr,
since changes in control hinge moments due to aerodynamic changes don't get reflected back into stick position and thence control movements.
That's odd because, from B707s ({#1} no hydraulic power controls) to 747s, the control yoke has always caused the same deflection in the flight control surface position despite any aerodynamic changes. The effort to move the yoke may have changed due to the artificial "feel" (Q pot on the 707; - Air Data Computer adjustment to spring stiffness on others), but we always knew how much the control surface was displaced by simply looking at the yoke (similar to the power steering on your car).

Stick free has simply been the same since Cherokee days. Let go and see what happens after the aircraft has been displaced from trimmed speed.

I agree it is completely different on Airbus FBW. The side stick gives no clue as to what the control surface is doing, and the only time "stick free" makes sense is in Direct Law.

{#1} We had "rudder boost" (hydraulic assistance to rudder for take off and landing only) to reduce VMCA / VMCG

[TURIN]

Read again what I said a couple of posts ago:
QUOTE]Just to round things off, the complete aircraft (tail on) aerodynamic centre will be around 50~55% mac so the aircraft would be stable even at the aft CG limit (as it must be of course to meet regulations)

Not much to be doing with failure modes either - if the CG is ahead of the aerodynamic centre the aircraft will be statically stable whatever the FBW is doing.

[/QUOTE]

I think we are in agreement Owain.

However, reading some of the other posts here (and elsewhere) the suggestion by some is that the CofG maybe AFT of MAC.

It can be so for certain high performance a/c but not air transport catagory.
Failure modes have to be taken into account. Unfortunately, things go wrong.
In an ideal world all a/c would be designed with all surfaces lifting (Canard?) but we are not in that world.

Good thread this. :ok:

[Owain Glyndwr]

rrr

That's odd because, from B707s ({#1} no hydraulic power controls) to 747s, the control yoke has always caused the same deflection in the flight control surface position despite any aerodynamic changes. The effort to move the yoke may have changed due to the artificial "feel" (Q pot on the 707; - Air Data Computer adjustment to spring stiffness on others), but we always knew how much the control surface was displaced by simply looking at the yoke (similar to the power steering on your car).That's because you are going one way down the control path, I'm going the other!

Let me quote again from my "Bible" - Dick Shevell's book (p.315 in Iss 2)

"Stick fixed stability means that the pilot maintains the control column in a fixed position so that the control surfaces, specifically the elevators, are not permitted to move as the airplane changes its angle of attack. Stick-free stability describes the value of dCm/dCL with the elevators assuming the position, at each airplane angle of attack, for which the elevator aerodynamic hinge moment is zero. This will be the angle at which the elevators will float with no control force. Stick free stability is important for airplanes with direct cable control of the surface and/or its control tabs mounted on the surface. Aircraft with full irreversible powered controls will have stick fixed stability even when the pilot is flying "hands off" provided the pilot has trimmed the column to zero force at the initial flying speed"

In other words stick free stability depends on the feedback from control aerodynamics, whereas you, I think, are talking about the direct yoke/control surface path? Your B707 would have had stick free stability, but not the B747. And on my definition, stick free stability is meaningless on the A320 even in direct law, because it still uses irreversible powered controls.

On aircraft without any artificial feel (any left? - well maybe your Cherokee and similar) the control forces felt at the column would be proportional to the aerodynamic hinge moment on the control, but otherwise you get what the designer thinks you want (which I agree may not be the same as you actually do want ;))

[Owain Glyndwr]

Wait, Now thinking I guess that certification requirements must account for failure of the FCS. Is that so? The aft CG limit of an A320 or a B777 is determined because of the direct law possibility?

Correct :ok:

[Owain Glyndwr]

It can be so for certain high performance a/c but not air transport category.
Failure modes have to be taken into account. Unfortunately, things go wrong.
In an ideal world all a/c would be designed with all surfaces lifting (Canard?) but we are not in that world.

Yeah, aircraft designed for high manoeuvrability might be designed to be unstable and rely on artificial stability.

Don't start me off on canards - they have their disadvantages :D

[rudderrudderrat]

Hi Owain Glyndw,
That's because you are going one way down the control path, I'm going the other!
You are correct.
Many thanks for all your lucid explanations.

[Microburst2002]

Owain

Is this Dick Shevell book apt for non engineerrs?

What's the name of the book? I hope I can get a pdf copy

[Owain Glyndwr]

Microburst,

As you probably guessed I am (was) an engineer, so perhaps not the best person to judge whether Shevell's book is apt for non-engineers. Dick (I met him once) was Chief Aerodynamicist at Douglas and then went to Stanford, so his book is full of realistic data. Looking through it it seems to be about 95% good, understandable text and explanations, but some maths is inevitable - only you can judge if it is too much, but I think you could just take the maths for granted and read the text.

Anyway, it is : Fundamentals of Flight, by Richard S Shevell, published by Prentice Hall, ISBN 0-13-339060-8. I don't know whether it is still in print though, and I doubt you will get a .pdf version - it is over 400 pages.

[Microburst2002]

Thank you, sir

[italia458]

TURIN....

"Most people find mathematics a very useful common language in these circumstances. It introduces a discipline that avoids the sort of confusion etc etc. "

I have to disagree. Most people use the spoken word to explain anything. Mathematics confuses almost everybody at one level or another. See above.

I have to completely disagree with your stance. Not to be personal but your attitude contributes to the lack of scientific literacy throughout North America and many other places. I can't stand people who justify using ambiguous and/or incorrect terminology or explanations, because in reality they don't understand the topic themselves. Frankly, if you don't understand it, you shouldn't be the one telling everyone how it should be described! You should look into Richard Feynman - he was a physicist who is known for his excellent explanations yet he also shares the same point of view as I do on this matter. If you don't want to listen to me, at least listen to a man who definitely had a good idea of what he was talking about. [/rant]

To add onto what John Tulla posted, for those of you who still don't quite understand/believe that a tail almost always will produce downlift, look at this video at 0:40.

Dam busters or bouncing bomb - YouTube

[TURIN]

Oh dear! Italia 458, As I'm not from North America I can't say either way if your rant is accurate our not.
However, one thing is certain. You need to settle down. My quoted post was meant to be light hearted. Life is too short. If you really can't stand me because I don't use high level maths to understand/explain a simple opening post such as the one above that is your problem. I'll stick to normal, conversational English if you don't mind. I seem to hang on to more friends that way. Be well.:)

[italia458]

TURIN...

My quoted post was meant to be light hearted.

What exactly do you mean by 'light hearted'? Is that supposed to mean that you actually didn't mean a word of what you said?

If you really can't stand me because I don't use high level maths...

I never asked you to use "high level maths". But what I would ask is that you use "accurate" descriptions and examples to explain something. Math is a good way to do that - I suggest you learn a little bit sometime! :ok:

I'll stick to normal, conversational English if you don't mind. I seem to hang on to more friends that way.

Hmm... that attitude seems to go hand-in-hand with the attitude that you should explain everything as simply as possible, even if it's not correct. So if your friends are wrong about something you just disregard it? My friends are generally happy to have an engaging discussion and so I don't have a problem with pointing out a potential error in our discussion.

[TURIN]

Like I said. Life's too short to get into pointless discussion such as this. I enter into pprune to learn and hopefully pass on some of my experience that others may also learn. If you want to get into a fight I suggest you join a gymnasium. I'm not interested.
Good night.

[PJ2]

Microburst2002, Owain Glyndwr;

Fascinating thread - thank you for this interesting exchange between the two of you. Microburst2002, I can highly recommend Owain's contributions on the AF447 threads. You can use mm43's PPRuNe search tool (http://countjustonce.com/pprune/) to search/find anything in PPRuNe.

I searched for a pdf copy of Shevell's work and while doing so I found the following related pdf on Aircraft Design - Synthesis and Analysis (http://www.ultraligero.net/Cursos/diseno/Diseno_de_aviones_sintesis_y_analisis.pdf). One of the authors is Ilan Kroo, Professor of Aerodynamics and Astronautics at Stanford but the introduction states, interestingly:
Richard Shevell was the original author of several of these chapters. He worked in aerodynamics and design at Douglas Aircraft Company or 30 years, was head of advanced design during the development of the DC-9 and DC-10, and taught at Stanford University after that for 20 years. To a large extent, this is his course.

[Owain Glyndwr]

Microburst,

PJ2 seems to have found the .pdf source I recommended to you via a PM ;)

It is very good.

[Microburst2002]

I don't know what you guys are talking about, honestly.

Processing data....

:}

[Owain Glyndwr]

Microburst,

If "you guys" are PJ2 and myself, I was referring to the alternative source of material similar to that contained in Shevell's book (which you were interested in obtaining) and which can be found at the URL PJ2 gave.

[Microburst2002]

Owain

Thankyou very much for your message, i received it well.

I just wanted to humoristically thankyou two for it without "incriminating" myself for breaching some rule :)

But it didnt go well... :(

[Owain Glyndwr]

Microburst,

Guess I'm taking things too seriously ;)

[megan]

I understand the tailplane lift downwards for stability purposes, but for manoeuvring, or for any other reason, would tailplane lift ever be upwards in flight on your typical FAR 23 or 25 aircraft - Airbus/Boeing computerised relaxed stability aside. I say in flight, because I know about lifting the tail of a tail dragger on take off.

Thanks Gals/Guys.

[Wizofoz]

megan,

Is short, no. It's a design requirement that full elevator in available up till Vmo without overstressing the airframe, and that would not be sufficient to generate positive lift on the tail.

BTW, it isn't the negative lift that is needed for stability- just that the lift component of the tail changes sufficiently with changes of AofA to ensure a corrective moment, while not stalling.

It is the last bit that means that most tails create negative lift, but it is not an actual requirement.

[megan]

I realise it's not a requirement, but as wondering if any aircraft actually produced positive (upward) lift on the tailplane in order to reach its negative "g" limit as per its V-n diagram.

[PDR1]

Yes, many/most aircraft with any aerobatic capability would be an obvious start (consider an aeroplane flying inverted, or doing an outside loop, or an inverted spin/flick). Look at the range of deflection of a typical all-moving tailplane which can be anything from +/-6degrees to +/-12degrees or more. If it is suggested that full "down" would still have the tailplane at a positive AoA it would mean that the S&L trimmed position would be -3 to -6 degrees. Do a few quick sums on the associated trim drag and it's easy to see that such a situation would never be tolerable/affordable for real operations.

And the very example you cite - pushing the stick to lift the tail of (say) a cub or chippie on take off. Clearly the tailplane must be lifting here, and if it can do it at such a low airspeed then it must have much more lifting power at flying speeds.

[john_tullamarine]

It's a design requirement that full elevator in available up till Vmo without overstressing the airframe

It's a tad late in the day for me .. you couldn't, by any chance, cite a reference for this ?

[Wizofoz]

No I can't because it's wrong.

I meant Va

[megan]

I've found an answer that satisfies me, thanks to those who responded. :ok:

FLIGHT, 23 February 1985, B. R. A. Burns, technical manager of BAe's Experimental Aircraft Programme

Tailplane download

In a stable tailed aeroplane the forward c.g. limit is determined by the ability of the tailplane to produce a download to balance wing lift, which acts behind the c.g. The tail must also counteract the "no lift" pitching moment generated by wing camber, which is greatest with flaps down. (A cambered wing produces lift at zero angle of attack (AoA). At zero lift, negative AoA, the wing carries a download at the front and an upload at the rear, producing a nose-down moment.)

The c.g. limit moves forward with increasing tail size at a rate proportional to tailplane maximum download capability. For this reason some aeroplanes feature devices to augment tailplane negative lift—leading-edge blowing on the Buccaneer and slats on the Phantom are examples.

The aft c.g. limit is determined by stability, and in a stable aeroplane is located a few per cent of wing chord ahead of the neutral stability point, where c.g. and a.c. coincide. The aft c.g. limit moves rearwards with increasing tailplane size, but only at a comparatively shallow rate because its effectiveness as a stabiliser is diminished by wing downwash, typically by 50 per cent.

It is a misconception that tailed aero-planes always carry tailplane downloads. They usually do, with flaps down and at forward c.g. positions, but with flaps up at the c.g. aft, tail loads at high lift are frequently positive (up), although the tail's maximum lifting capability is rarely approached.

As our tail sizing diagram shows (Fig A), there is a steep increase in c.g. range with increasing tail size as the limits move apart. In the example illustrated, a 40 per cent increase in tail size doubles the c.g. range from 10 per cent to 20 per cent of wing chord. Also the c.g. range is located well within the wing chord, so that the trimming load on the tail (and trim drag) is small.

[john_tullamarine]

I meant Va

Ah .. much relief ... thanks, good sir. Thought that my mental archives were going to have an haemorrhage there for a bit ..

[zzuf]

Well, AC 23-19A has a different take on this.

"48. What is the design maneuvering speed VA?

a. The design maneuvering speed is a value chosen by the applicant. It may not
be less than Vs√ n and need not be greater than Vc, but it could be greater if the applicant chose the higher value. The loads resulting from full control surface deflections at VA are used to design the empennage and ailerons in part 23, §§ 23.423, 23.441, and 23.455.

b. VA should not be interpreted as a speed that would permit the pilot
unrestricted flight-control movement without exceeding airplane structural limits, nor
should it be interpreted as a gust penetration speed. Only if VA = Vs √n will the airplane
stall in a nose-up pitching maneuver at, or near, limit load factor. For airplanes where
VA>VS√n, the pilot would have to check the maneuver; otherwise the airplane would
exceed the limit load factor.

c. Amendment 23-45 added the operating maneuvering speed, VO, in § 23.1507.
VO is established not greater than VS√n, and it is a speed where the airplane will stall in a nose-up pitching maneuver before exceeding the airplane structural limits."

While the FAA did not introduce a Vo for FAR Part 25 aircraft (after American Airlines Flight 587 12Nov2001), the full control deflections are about rudder, elevator and aileron structure. Careful reading of FAR 25.331 shows that at Va, with maximum elevator deflection, loads which occur beyond the limit load need not be considered.

[Pugilistic Animus]

https://www.pprune.org/private-flying/300139-minimum-radius-turn.html#post3704889

In this post I derive the formula for Va

[zzuf]

Why do you need to derive VA?
The FAR's are quite clear. The applicant can choose any speed above VA min.
Where in the FAR's is there a requirement to associate ultimate load factor to VA at the stall?

[Pugilistic Animus]

Not the FARs...Physics

[zzuf]

I would be really greatfull if you can reference the section of FAR25, or FAR23 which associates maximum load factor, stall and Va.
It may be the case at Va min, but not at any other speed which the applicant can select as VA.

[Pugilistic Animus]

My only point was that Va is a function of n and Vs...It varies...don't feel up to reading FARs right now as I'm trying to read an interesting book about King Henry

[zzuf]

It isn't, perhaps you are thinking of VO.
Anyway too much thread drift.

[Pugilistic Animus]

Vp I never heard of Vo

[zzuf]

Ok, post 86 in this thread.

[Pugilistic Animus]

I saw it already...I mean this thread is the first time I heard of Vo

[TURIN]

And the very example you cite - pushing the stick to lift the tail of (say) a cub or chippie on take off. Clearly the tailplane must be lifting here, and if it can do it at such a low airspeed then it must have much more lifting power at flying speeds.

Not necessarily. The tailplane just needs to be producing less downforce than needed.