Tailplane lift
 

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).

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

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...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.

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

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!

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.

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.

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.

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.

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.

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.

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.

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.

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).

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 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.

Hi italia458,
Surely an aft CG helps improve cruise performance?

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.

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:

It sure does! Thanks, I made the correction.

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.

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".

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

Hi Microburst2002,
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.

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

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.

- try drawing that out on paper and maybe edit your post? - 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..........................

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

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...

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.

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.

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

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.

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

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

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.

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.

.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.

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.

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:

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!

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:

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.

Turin
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

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/...kYzerolift.jpg

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

"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. :)