Critical angle of attack and stall
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Critical angle of attack and stall
I am confused.Again.
According to theory,the critical angle of attack it the angle of attack:
1)where the CL (coefficient of lift) reaches its maximum value,and beyond which it will decrease
2)beyond which the separation point will syddenly move well forward
3)beyond which the center of pressure will suddenly move rearward
4)where the aircraft stalls.
We know that the aircraft stalls when the boundary layer has mostly separated from the wing (breakdown of streamline flow above wing) and there is no sufficient lift.So,due to decreased lift,it will start to sink.
But,the theory says that stall happens at the critical or stalling angle of attack where (as I wrote above) the CL is at its maximum value,and the lift produced at that moment can counteract weight.
So,stall happens at the critical angle of attack of shortly after?
Can anyone help me?
Cheers!
According to theory,the critical angle of attack it the angle of attack:
1)where the CL (coefficient of lift) reaches its maximum value,and beyond which it will decrease
2)beyond which the separation point will syddenly move well forward
3)beyond which the center of pressure will suddenly move rearward
4)where the aircraft stalls.
We know that the aircraft stalls when the boundary layer has mostly separated from the wing (breakdown of streamline flow above wing) and there is no sufficient lift.So,due to decreased lift,it will start to sink.
But,the theory says that stall happens at the critical or stalling angle of attack where (as I wrote above) the CL is at its maximum value,and the lift produced at that moment can counteract weight.
So,stall happens at the critical angle of attack of shortly after?
Can anyone help me?
Cheers!
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@Radix
Thanks for your answer!
As far as 3) is concerned: you are right,I didn't explain it correctly.I meant that when we have CLmax,we can produce as much lift to counteract weight,and I refer only to 1G straight and level flight.
Thanks again!
Thanks for your answer!
As far as 3) is concerned: you are right,I didn't explain it correctly.I meant that when we have CLmax,we can produce as much lift to counteract weight,and I refer only to 1G straight and level flight.
Thanks again!
@dims
I am not sure where you are getting this "critical angle of attack" term from.
I highly recommend a visit to the library and see some books. Of course, Wiki is likely to have good stuff on aero or some links.
You can "stall" at damned near any speed, gross weight, gee, bank angle - there must be "fifty ways". It all comes down to one thing - AoA. The classic graphic of AoA and Cl seems out dated to this old fart. The older, straight wing and large camber wings behave like that. The swept wings and airfoils since the 60's do not. You can fly well beyond the peak of the Cl point on the graph and not "stall". On some planes ( like deltas - think Concorde), you never lose a lotta roll authority or directional authority, but develop a high sink rate. Call it stall, but simply lowering the nose will regain "normal" flying characteristics real quick.
Go to the library and then come back.
I am not sure where you are getting this "critical angle of attack" term from.
I highly recommend a visit to the library and see some books. Of course, Wiki is likely to have good stuff on aero or some links.
You can "stall" at damned near any speed, gross weight, gee, bank angle - there must be "fifty ways". It all comes down to one thing - AoA. The classic graphic of AoA and Cl seems out dated to this old fart. The older, straight wing and large camber wings behave like that. The swept wings and airfoils since the 60's do not. You can fly well beyond the peak of the Cl point on the graph and not "stall". On some planes ( like deltas - think Concorde), you never lose a lotta roll authority or directional authority, but develop a high sink rate. Call it stall, but simply lowering the nose will regain "normal" flying characteristics real quick.
Go to the library and then come back.
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Critical Angle of Attack is referenced in nearly all my textbooks (usually as approx 16 degrees) and described as the point Coefficient of Lift is at it's maximum- Exceed the 'critical angle' and the lift begins to reduce.
Have a look at the lift graphs in the book, you'll see the graph progressively rising, peaking at the critical angle and then falling off sharply from there.
Have a look at the lift graphs in the book, you'll see the graph progressively rising, peaking at the critical angle and then falling off sharply from there.
@ Laz
O.K., "critical angle of attack" seems to be the max AoA before lift begins to reduce according to you and Dims and whatever reference you guys are using.. Oooops, not lift, but Cl.
The newer wings have a lot more gentle curve up near max Cl. The old graphics with the sharp break are not there except for very old airfolil designs, most likely used by general aviation planes. For example, some newer planes reach the max Cl at "x" degrees, but do not get a lot lower Cl for 5 or 10 degrees more AoA. They just have more drag and sink.
The Viper I flew could get up to 25 degrees AoA with great performance and not a lot of induced drag. Leading edge flaps and the "flaperons" on the trailing edge changed the camber of the wing. So I would not place a firm value on AoA for the "critical AoA'", like the 16 degree number. And as I asserted, the deltas had a peak Cl on the graph, but simply started to sink when you got a lot higher AoA. Some planes can not even fly with AoA above 10 degrees or less without stalling. All depends on the airfoil used.
O.K., "critical angle of attack" seems to be the max AoA before lift begins to reduce according to you and Dims and whatever reference you guys are using.. Oooops, not lift, but Cl.
The newer wings have a lot more gentle curve up near max Cl. The old graphics with the sharp break are not there except for very old airfolil designs, most likely used by general aviation planes. For example, some newer planes reach the max Cl at "x" degrees, but do not get a lot lower Cl for 5 or 10 degrees more AoA. They just have more drag and sink.
The Viper I flew could get up to 25 degrees AoA with great performance and not a lot of induced drag. Leading edge flaps and the "flaperons" on the trailing edge changed the camber of the wing. So I would not place a firm value on AoA for the "critical AoA'", like the 16 degree number. And as I asserted, the deltas had a peak Cl on the graph, but simply started to sink when you got a lot higher AoA. Some planes can not even fly with AoA above 10 degrees or less without stalling. All depends on the airfoil used.
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A bit off topic but I was wondering, when an airfoil is at the aoa for max CL (critical aoa term used here in this thread), can we expect to already see some flow separation on the upper surface of the wing? Or is it generally the case that the flow is still completely attached, in order to get this maximum CL value.
anybody?
anybody?
Except maybe for an ideal laminar flow wing. This exists only on paper though.
In reality it is pretty safe to assume you are correct.
The polars @dimsp81 is probably referring to are usually for an infinite and ideal wing without twist and taper, sweep, slats and other geometrical properties or also physical properties like rivets or dents or other imperfections of a real wing which are significantly influencing separation behaviour.
So this critical AoA thing is nice but to which section in which condition of a real wing does it refer?
Due to these mentioned geometrical features the separation will be a progressive process with increasing AoA but starting at very low AoA.
Last edited by henra; 9th May 2013 at 09:04.
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OK, flow never completely attached, I get that. any guesses on how far back it may separate for a generic type airliner wing at max CL aoa?
I'm just thinking, if it was say at approximately the 3/4 chord point, ie 3/4 the way back approximately, that would mean the wing was getting it's max CL value with the air only attached for 3/4 of the area or so. That would be an interesting thought to me.
Maybe that 3/4 is too far forward. But is separation at 99% chord too far back for the max CL point?
I'm just thinking, if it was say at approximately the 3/4 chord point, ie 3/4 the way back approximately, that would mean the wing was getting it's max CL value with the air only attached for 3/4 of the area or so. That would be an interesting thought to me.
Maybe that 3/4 is too far forward. But is separation at 99% chord too far back for the max CL point?
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I am referring to a tupical finite wing,only for educational ppl purposes.I am not interested for a special case of wing.
Sorry If I confused you with the critical angle term.It is a term that trevor thom uses in his books.Just call it stalling or stall angle of attack,and it is usually 16 degrees.
Sorry If I confused you with the critical angle term.It is a term that trevor thom uses in his books.Just call it stalling or stall angle of attack,and it is usually 16 degrees.
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A problem here is the tendency to want to make things black and white when the real world is many shades of grey.
A few thoughts ..
When we refer to a stalling angle of around 16deg, that is appropriate for an older, traditional aerofoil profile and is the typical dumbed down story for new pilots.
As with any form of training or information transfer, the detail level needs to be appropriate to the needs and comprehension of the trainee. For the new chum, the story needs to be simplistic until the basic ideas are comprehended and then the student can be introduced to more complex and detailed stories about what might be happening.
Doesn't matter whether we are talking about Scouts or similar learning the absolute basics for a badge of some sort through to trainee pilots then progressing to airline folk or, for that matter, first aiders progressing over a period of time to become brain surgeons .. same principle applies.
Keep in mind that any of the nice pictures you see will be generated from (typically tunnel) data for quasi-steady flow states. Once you get out into the real world it starts to get a little muddied.
It is a term that trevor thom uses in his books
Critical angle is fine to think about .. whatever it might be for a particular aerofoil section .. somewhat below that and things are comfortable .. progressing to more interesting things as you increase to higher angles. Please don't suggest that 16 deg is the number for all wings .. that's nonsense. The basic textbook Cl by alpha characteristic is sort of typical but varies a LOT according to the profile and, also, pitch rate (especially for helicopters).
I haven't spoken to Trev for many years now .. he did a lot of good work in the original development of his books (with a bit of help from a few of us along the way). Flew with him a few times in years long gone by ... thoroughly nice chap. I would imagine that the present business, run by Dave Robson, still is a purveyor of fine books albeit, I imagine, with a bit more of an emphasis on technical rigour. Flew with Dave only the once and that, many, many years ago.
A few thoughts ..
When we refer to a stalling angle of around 16deg, that is appropriate for an older, traditional aerofoil profile and is the typical dumbed down story for new pilots.
As with any form of training or information transfer, the detail level needs to be appropriate to the needs and comprehension of the trainee. For the new chum, the story needs to be simplistic until the basic ideas are comprehended and then the student can be introduced to more complex and detailed stories about what might be happening.
Doesn't matter whether we are talking about Scouts or similar learning the absolute basics for a badge of some sort through to trainee pilots then progressing to airline folk or, for that matter, first aiders progressing over a period of time to become brain surgeons .. same principle applies.
Keep in mind that any of the nice pictures you see will be generated from (typically tunnel) data for quasi-steady flow states. Once you get out into the real world it starts to get a little muddied.
It is a term that trevor thom uses in his books
Critical angle is fine to think about .. whatever it might be for a particular aerofoil section .. somewhat below that and things are comfortable .. progressing to more interesting things as you increase to higher angles. Please don't suggest that 16 deg is the number for all wings .. that's nonsense. The basic textbook Cl by alpha characteristic is sort of typical but varies a LOT according to the profile and, also, pitch rate (especially for helicopters).
I haven't spoken to Trev for many years now .. he did a lot of good work in the original development of his books (with a bit of help from a few of us along the way). Flew with him a few times in years long gone by ... thoroughly nice chap. I would imagine that the present business, run by Dave Robson, still is a purveyor of fine books albeit, I imagine, with a bit more of an emphasis on technical rigour. Flew with Dave only the once and that, many, many years ago.
Also just to point out that this-
Isn't necessarily so. Post stall, you still have a Cl, so you can still generate Lift that equals weight, if you have sufficient speed.
As far as 3) is concerned: you are right,I didn't explain it correctly.I meant that when we have CLmax,we can produce as much lift to counteract weight,and I refer only to 1G straight and level flight.
A problem here is the tendency to want to make things black and white when the real world is many shades of grey.
A few thoughts ..
When we refer to a stalling angle of around 16deg, that is appropriate for an older, traditional aerofoil profile and is the typical dumbed down story for new pilots.
A few thoughts ..
When we refer to a stalling angle of around 16deg, that is appropriate for an older, traditional aerofoil profile and is the typical dumbed down story for new pilots.
Knowldege of the different AoA at which Clmax occurs depending on the exact profile chosen is even quite old for that matter. The good old NACA Report 460 dating back to the thirties shows nicely how different the chrarcteristic of Cl can be around clmax and at what a wide range of AoA this Clmax can occur.
And those are all only old NACA 4 digit profiles.
http://naca.central.cranfield.ac.uk/...report-460.pdf
And it doesn't include Leading Edge devices or other geometrical properties of a real world wing which can enourmously change Alpha(Clmax).
So even for a theoretical infinite wing in a wind tunnel a statement that a wing stalls at 16° AoA is nonsense.
It might be helpful as a rule of thumb but in reality even that is not worth much.
If the Buffeting starts, it might be wise to stop pulling, even if your AoA gauge only shows 14°.
I am glad to see JT and Henra post here.
I apologize to Dims for suggesting he visit the library. It was crass.
The best thing I saw was Henra's warning about "feeling" the buffet. And then JT's observation that a specific AoA may or may not define max Cl or a stall.
I was personally blessed to learn in Aeronica's, Luscombs and Taylorcraft. Then to USAF, and Tweets and the T-33, and then a mix of bent wings and straight wings..... The big thing in all of them was "touch" or "feel". The plane talks to you, and those who do not listen or cannot listen will pay the price, regardless of what the indicators display or the control system is designed for. See the AF447 threads.
Good thread, IMHO.
I apologize to Dims for suggesting he visit the library. It was crass.
The best thing I saw was Henra's warning about "feeling" the buffet. And then JT's observation that a specific AoA may or may not define max Cl or a stall.
I was personally blessed to learn in Aeronica's, Luscombs and Taylorcraft. Then to USAF, and Tweets and the T-33, and then a mix of bent wings and straight wings..... The big thing in all of them was "touch" or "feel". The plane talks to you, and those who do not listen or cannot listen will pay the price, regardless of what the indicators display or the control system is designed for. See the AF447 threads.
Good thread, IMHO.
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Just chipping in with my 2c...
There is a lot of talk of AoA, specific angles, and initiation of stall recovery here. The question is, are we discussing stall of an aerofoil or an aircraft?
As above, a fairly standard aerofoil section will stall at or around 16 degrees AoA, this is repeatable and well understood.
Once the aerofoil is attached to an aircraft, and that aircraft is in motion all factors change. Its also of paramount importance to understand that the AoA of the wing, is completely different to the aircraft 'pitch', they are two mutually exclusive states.
Once you suss out the concept of the Relative Air Flow, you'll find all this to be much more clear. Best of luck!
There is a lot of talk of AoA, specific angles, and initiation of stall recovery here. The question is, are we discussing stall of an aerofoil or an aircraft?
As above, a fairly standard aerofoil section will stall at or around 16 degrees AoA, this is repeatable and well understood.
Once the aerofoil is attached to an aircraft, and that aircraft is in motion all factors change. Its also of paramount importance to understand that the AoA of the wing, is completely different to the aircraft 'pitch', they are two mutually exclusive states.
Once you suss out the concept of the Relative Air Flow, you'll find all this to be much more clear. Best of luck!
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Gee, I would agree in principle, but even minor cross-control (uncoordinated flight) could rob the pilot of a warning buffet. I learned that lesson in a few memorable seconds in an old PT-22, fifty years ago. My instructor asked for a stall with about 1/3 rudder, and we did the prettiest snap roll 
with NO buffet warning, only a knot or two above Vso.
The big thing in all of them was "touch" or "feel". The plane talks to you, and those who do not listen or cannot listen will pay the price, regardless of what the indicators display or the control system is designed for. See the AF447 threads.
with NO buffet warning, only a knot or two above Vso.