Tailplane lift
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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.
If wing aoa increases, so does tailplane's. Can't be otherwise.
That's why normal commercial aircraft are able to be trimmed angle of attack stable.
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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.
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.
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.
Last edited by italia458; 12th May 2012 at 12:41.
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Originally Posted by micro
Normally drag line of action is below cg, so it creates nose up moment.
Originally Posted by italia
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.
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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...
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...
Last edited by Microburst2002; 13th May 2012 at 13:07.
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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.
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?
Last edited by italia458; 14th May 2012 at 01:50.
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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
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
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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.
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.
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.
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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.
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Indeed, bubbers. Hence my suggestion that to avoid the back axle effect italia defined his/her "axes and values"
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.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.
Last edited by BOAC; 15th May 2012 at 07:04.
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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.
When AoA increases the incidence on both wing and stabiliser increases,
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!
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?
Last edited by TURIN; 15th May 2012 at 21:54. Reason: spilling.
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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!
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?
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.
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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.
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.
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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)
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
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.
Microburst
What is different in the circulation of airflow created by AoA with respect to that created by camber?
lift curve slope*(AoA - AoA zero lift)
or for the Clark Y:
lift curve slope*(AoA + 3.35)
There is a relation Camber-circulation-Aerodynamic moment that I don't quite understand.
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
Last edited by Owain Glyndwr; 16th May 2012 at 14:35.
"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 disagree. Most people use the spoken word to explain anything. Mathematics confuses almost everybody at one level or another. See above.