Tailplane lift
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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
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
Last edited by Microburst2002; 18th May 2012 at 10:31.
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Microburst
Know the feeling
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.
Yes, and I am darned glad you don't want a theoretical explanation
Yes
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.
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).
Agreed, but I think you need to retain the discipline of mathematical conventions in your "normal" wording.
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?
I think that maths are a kind of language but it is is good to frequently translate it into normal words
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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?
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?
To answer your last point. Yes, but the a/c would not be naturally stable.
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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?
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?
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)
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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.
hitting the stabilizer creating a downforce. The higher airspeed, the stronger the downwash
and more downforce.
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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.
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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.
hitting the stabilizer creating a downforce. The higher airspeed, the stronger the downwash
and more downforce.
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.
Last edited by Owain Glyndwr; 19th May 2012 at 10:05.
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.
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OG,
Pilot's Handbook of Aeronautical Knowledge Chapter 3 - American Flyers
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/aviati..._3-2_img_6.jpg
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.
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.
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/aviati..._3-2_img_6.jpg
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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.
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.
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.
Last edited by Owain Glyndwr; 20th May 2012 at 11:03. Reason: tidying up
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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"
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
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
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.
I think that is the reason for our apparent difference of opinion
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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 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
Last edited by XPMorten; 19th May 2012 at 13:22.
Moderator
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.
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.
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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
an increased downforce on the tail. F=m*a
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.
Last edited by Owain Glyndwr; 19th May 2012 at 14:03.
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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!
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!
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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.
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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?
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?
Last edited by Microburst2002; 20th May 2012 at 09:20.
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Hi Owain Glyndwr,
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
since changes in control hinge moments due to aerodynamic changes don't get reflected back into stick position and thence control movements.
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
Last edited by rudderrudderrat; 20th May 2012 at 09:37. Reason: still can't spell
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]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]
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.