Questions on High Speed Stall

[corpjet]

I have a couple of questions regarding the high speed stall. I don't know if this is the right link or not, but here it goes:
Q1: Where is it more probable to have a High Speed Stall. High altitudes or Low altitudes?
Q2: With maximum weight, where is it more probable to have a HSS.
High altitudes and low speeds? or
High Altitudes and High speeds?
Thanks everybody

[WingletTurbo]

I don't think altitude really matters. A high speed stall can occur at any speed at any height. If you’re flying at, lets say 70 KIAS and pull all the way back to give full elevator deflection, the wing will rise immediately past the critical angle of attack therefore disrupting the smooth airflow over the wing. The definition of a stall is the same regardless of the speed. Of course, at a greater weight your plane will stall faster than a lighter plane, just because it has to flying closer to the CAOA to maintain altitude... but the subject of weight is also the same subject why Va increases with weight..

Noticed your from the U.S.. that question will definitely be on your oral exam :)

Good luck!

[chornedsnorkack]

Corpjet, the previous two answers cover low speed stalls, not high speed stalls.
A low speed stall is when the airflow over your airfoil separates due to a too high angle of attack.
It is a function of angle of attack and not of speed. Many people mix this up with speed, because for a certain weight, speed (IAS) and angle of attack are in relation to each other, e.g. the slower you fly the more angle of attack you need to produce a certain amount of lift. But separation is a function of angle of attack and not of speed.
With higher altitude (= reduced air-density), you need a higher TAS to have the same IAS, thus the higher you fly, the higher will be your TAS to carry a certain weight (good for traveling).
A high speed stall is when the airflow over your wing reaches the speed of sound (M = 1 at this point of the wing) due to acceleration around the wings shape. A shockwave emerges and the airflow separates. [Note: you might have heard of the term Mcrit or critical machnumber. This is the aircrafts theoretical speed when its airflow over the wing reaches M = 1]
The higher you fly the lower is your TAS needed to reach M = 1 (due to lower temperature). The higher you fly, the easier you end up in a high speed stall.
So now we combine low and high speed stall: at a certain altitude, the minimum TAS you need to avoid a low speed stall will be the maximum TAS you can fly before entering a high speed stall. This point is called "coffins corner" and theoretically the highest altitude an aircraft can fly.


Just what is wrong about having airflow separate? Especially at high Mach?

Look at it this way: the lower surface of the wing intercepts airflow, which is forced to accelerate around the wing. The wing creates lift at all angles of attack except approximately 0 degrees, approximately 90 degrees, approximately 180 degrees and approximately 270 degrees.

At approximately 0 degrees, there is no lift because there is exactly as much air accelerated above wing as below wing. At 90 degrees, the same applies.

Now, from 0 to certain stalling angle of attack, the larger the angle of attack, the larger the lift. This should provide the necessary vertical and therefore lateral stability.

So what if the speed of airflow exceeds the speed of sound over the wing, or under the wing, or in free airflow? Yes, shockwaves are generated, and the airflow separates easily over the wing - but the airflow under the wing has nowhere to separate and has to support the wing for simple Newton second law reasons, whether its speed is 0,8 M, 1,8 M or 18 M...

[GE 90]

Just what is wrong about having airflow separate? Especially at high Mach?
Look at it this way: the lower surface of the wing intercepts airflow, which is forced to accelerate around the wing. The wing creates lift at all angles of attack except approximately 0 degrees, approximately 90 degrees, approximately 180 degrees and approximately 270 degrees.
At approximately 0 degrees, there is no lift because there is exactly as much air accelerated above wing as below wing. At 90 degrees, the same applies.
Now, from 0 to certain stalling angle of attack, the larger the angle of attack, the larger the lift. This should provide the necessary vertical and therefore lateral stability.
So what if the speed of airflow exceeds the speed of sound over the wing, or under the wing, or in free airflow? Yes, shockwaves are generated, and the airflow separates easily over the wing - but the airflow under the wing has nowhere to separate and has to support the wing for simple Newton second law reasons, whether its speed is 0,8 M, 1,8 M or 18 M...

Not sure what kind of wing you are thinking of. Lift is created at 0 degrees AOA by the action of air passing over the top surface at a higher speed than under it. This causes a lower pressure area to form above the wing thus sucking the a/c into the air. If you increase the AOA until the air separates the lift is lost so it is only the engines thrust pushing up that will act against the weight. If thrust is insufficent then you will descend. (Possibly backwards as happened to a Trident many years ago.)
At close to Mach1 the air over the top of the wing will be supersonic and so will form a shockwave. As the air hits the shockwave it is squashed and so creates a high pressure area. (The reverse of what you want on top of the wing.) Newton's laws have very little to do with lift compared to Berniolie (Sorry having spelling breakdown.) and his gas flow laws. It is also one of the reasons for "control reversal" when the pioneers were trying to break the sound barrier.
Feel like I am teaching most of you to suck eggs but perhaps I have misunderstood chornedsnorkack's post.

[chornedsnorkack]

Not sure what kind of wing you are thinking of. Lift is created at 0 degrees AOA by the action of air passing over the top surface at a higher speed than under it.
True for a cambered airfoil.

For a symmetrical airfoil, at 0 degrees AOA, the air passes over top and bottom surface at equal speeds, so lift is 0 by symmetry. For an asymmetrical airfoil, lift does not have to be zero at exactly 0 degrees AoA - and for cambered airfoils, it is positive. However, for every cambered airfoil, there is an angle of attack where lift does equal zero, this being less than 360 degrees - but normally much closer to 360 than 270 degrees. Anyway, AOA of cambered airfoil is often defined so as to be 0 at zero lift... (though it seems to me that such definition cannot be independent of Mach).
This causes a lower pressure area to form above the wing thus sucking the a/c into the air.
And at the same time, a high pressure area forms under the wing, pushing the a/c into the air.
If you increase the AOA until the air separates the lift is lost so it is only the engines thrust pushing up that will act against the weight.
But there is still the high pressure area under wing. The only time the engine thrust remains the only force counteracting weight is when the AoA is about 90 degrees, so that the lift is directed horizontally.
At close to Mach1 the air over the top of the wing will be supersonic and so will form a shockwave. As the air hits the shockwave it is squashed and so creates a high pressure area. (The reverse of what you want on top of the wing.)

The air hitting the actual underside of the wing is still squashed against the underside and forms high pressure area. Which is what you want under the wing.

[DC-Mainliner]

True for a cambered airfoil.
For a symmetrical airfoil, at 0 degrees AOA, the air passes over top and bottom surface at equal speeds, so lift is 0 by symmetry. For an asymmetrical airfoil, lift does not have to be zero at exactly 0 degrees AoA - and for cambered airfoils, it is positive. However, for every cambered airfoil, there is an angle of attack where lift does equal zero, this being less than 360 degrees - but normally much closer to 360 than 270 degrees. Anyway, AOA of cambered airfoil is often defined so as to be 0 at zero lift... (though it seems to me that such definition cannot be independent of Mach).
And at the same time, a high pressure area forms under the wing, pushing the a/c into the air.
But there is still the high pressure area under wing. The only time the engine thrust remains the only force counteracting weight is when the AoA is about 90 degrees, so that the lift is directed horizontally.
The air hitting the actual underside of the wing is still squashed against the underside and forms high pressure area. Which is what you want under the wing.

Hi there-

Your thoughts describe a portion of newtonian lift, but a symetrical airfoil at high altitude at zero angle of attack in equilibrium is not likely :suspect:

The book, "Fly the wing", in the high speed, high altitude section of the book, really explains the physics of high speed flight in a straight forward, yet more detailed manner than many sources.

Studi has it right for high altitude flight.

I suspect part of the original question asks the following clarification that has already been offered: For any speed flight , low or high, the wing will ALWAYS enter a traditional stall at the critical angle of attack at lower and mid altitudes.

At high altitude, the "perception" of an aerodynamic wing stall is blurred slightly by a few factors. First, a relatively "high" angle of attack, but not necessarily the "critical" angle of attack, may force the air over the wing to reach such a speed where sufficient shock waves form on the upper surface of the wing to seperate the local airflow. The significant buffeting, change of center of pressure, shift in downwash angle and lift distrubution change.. they closely give you the "indications" of a traditional, immenent stall. The difference in this case is that a shock wave is disrupting the airflow at a angle of attack that is often below the critical angle of attack.

The above scenario occurs at a low envelope speed, meaning it is the "low speed buffet" of the high altitude envelope. This is from the aircraft being too heavy for the altitude - and the required lift needed by the wing (usually higher in the root area where airfoil shape is typically the most dramatically cambered) accelerates the air to a supersonic speed in a more dramatic way than typical. (local, small areas of mach flow normally exists at normal cruise flight but here we have a much greater area of supersonic flow and widespread seperation).

The high speed buffet is where the aircraft reaches buffeting, generally, more consistantly from shock waves all over the airframe, like the tail, the fuselage, the wings, etc. This is from the aircraft simply flying too fast for the aerodynamic design.

The sources that speak about this subject in depth go into the shock wave induced stall buffet in detail. however, the majority of sources that do not break down the "whys and hows" simply call it a high altitude stall buffet. :8

[BOAC]

I think we are in danger of confusion here - what exactly do you mean by a 'high-speed stall'? The 'Shock stall' referred to above is not the 'high-speed stall'. The HSS is when the angle of attack of the wing reaches stalling angle at a speed above the straight and level stall speed. This occurs when the aircraft is being 'accelerated' ie 'g' is being pulled, eg in a steep turn or aeros, because the 'weight' of the aircraft is increased by the 'g' thus raising the stall speed.

IF this is what you are asking about, my answers are:
Q1 Low altitude (more 'g' available)
Q2 High and High ( " " )

NB Different answers for a 'shock stall':eek:

[WingletTurbo]

Are high speed stalls also known as accelerated stalls? (Not sure if were using different terminology because of language differances) Because if the answer is yes...

Then my question for studi:

How can a high speed stall happen on a cessna if the speed of the air over the wing DOES NOT reach the speed of sound? I still stand by my answer if were talking about the same thing.

EDIT.

BOAC gave a better explanation. Just like when your in a 60 degree turn... load factor increases therefore increasing stall speed.

-WT

[JackOffallTrades]

I've never read so many different, widely varied answers to a question on simple aerodynamics in my life!

Perhaps it's time to read the books again chaps. :{

[BOAC]

Go on, then, Jack - your answers?

[DC-Mainliner]

Studi has it right for high altitude flight.
I suspect part of the original question asks the following clarification that has already been offered: For any speed flight , low or high, the wing will ALWAYS enter a traditional stall at the critical angle of attack at lower and mid altitudes.


Yup. This sort of just started to become a discussion on the previous posts. But what good is a question and answer if it can't develop into a discussion? :E

I agree. Light Cessnas - a "high speed" stall is probably an "accelerated stall" occurring at any airspeed. Just raise the alpha (AOA) to the critical angle of attack and you have a stall - Voila. A "1G" stall is at Vs, a 2G stall is at a higher "accelerated" stall speed.

BTW, a "high speed stall" is really not a good term to use unless it is defined as in a certain flight envelope. I think that is what led to the high altitude discussions.

[High Wing Drifter]

JackofAllTrades,

Master of one perhaps? Simple aerodynamics? Transonic to supersonic is about as difficult as it gets :uhoh:

WingletTurbo,
Are high speed stalls also known as accelerated stalls?No they are completely different. The high speed stall, also called the "mach tuck", "mach stall", etc is not a stall at all, it just gives the appearance of a stall (nose down). It is specifically due to the centre of pressure moving fore and aft (and very aft at about Mach .90) as the aeroplane transitions from about Mach .7 to 1.0. That is what Studi described.

I believe this where the all moving tailplane comes in as it is powerful enoough to out trim the effects. Not sure, but I think Concorde used to pump fuel fore an aft to help control it too.

Accelerated stalls are due to increasing the angle of attack beyond the stalling angle at speeds above stall speed due to severe turbulence or over zelous use of the elevator.

[chuks]

The way it used to be taught on a basic level for little airplanes in the States was that you had a certain stall speed for maximum weight in straight-and-level flight. Let's say that was 50 KIAS. It would have been mentioned that this was due to airflow separation and that stalls could occur at any speed and attitude but here we are just discussing very basic instruction.

When you banked the aircraft then you saw the stall occur at a higher speed than that for straight and level. For instance at 60 degrees of bank that would have been 1.44 (the square root of the 2-G loading in a 60-degree banked turn) times 50, or 72. So, because the stall speed was high relative to the unaccelerated stall speed it was called a high-speed (more properly 'accelerated') stall.

The student was taught to recognise this sort of stall by entering a banked turn and then slowing to the buffet with recovery made by first levelling the wings. Since the aircraft was already above the unaccelerated stall speed that did the trick! It was a simple, safe way to demonstrate an accelerated stall.

Unfortunately, lots of students just remembered stalls as a speed-related, rather than an angle-of-attack-related phenomenon since that was the simplified way it was shown to them.

Shock stalls or high-speed buffet would not have been discussed at this level since light training aircraft don't get into that regime so that it cannot be demonstrated. I was interested to see it for the first time in a Cessna 441, since I had pretty much forgotten about it. The numbers were right, so where was this buffet coming from? Then I remembered that I was up at 25 000 feet, not down around 4 000 feet! Oh.

[WingletTurbo]

JackofAllTrades,

Master of one perhaps? Simple aerodynamics? Transonic to supersonic is about as difficult as it gets :uhoh:

WingletTurbo,
No they are completely different. The high speed stall, also called the "mach tuck", "mach stall", etc is not a stall at all, it just gives the appearance of a stall (nose down). It is specifically due to the centre of pressure moving fore and aft (and very aft at about Mach .90) as the aeroplane transitions from about Mach .7 to 1.0. That is what Studi described.

I believe this where the all moving tailplane comes in as it is powerful enoough to out trim the effects. Not sure, but I think Concorde used to pump fuel fore an aft to help control it too.

Accelerated stalls are due to increasing the angle of attack beyond the stalling angle at speeds above stall speed due to severe turbulence or over zelous use of the elevator.


Thank you for the clarification! :)

[Loose rivets]

A week or so ago there was a link to a subject, old training films and the like I think, but in there somewhere there was training for (Saber) swept wing that showed all of the characteristics against graphs and pictorials.

It made quite a point about the variation with height for all kinds of maneuvers

Sorry I can't locate it quickly, but the old style American training was very practical and even I could understand it.

[411A]

>>I believe this where the all moving tailplane comes in as it is powerful enough to out trim the effects.<<

Indeed so.
And, which large transport jet uses same?

TriStar, of course. The -500 model even better, with MDLC and ACS.
Simply a better design.

Ahhhh, Lockheed:D

Sorry folks, just could not resist:E
Will fly the old girl for the next few months, and a pleasure it is...without a doubt.

[chornedsnorkack]

Hi there-
Your thoughts describe a portion of newtonian lift, but a symetrical airfoil at high altitude at zero angle of attack in equilibrium is not likely :suspect:

Well, it won´t be in equilibrium any more than accelerated stall is an equilibrium condition. If the plane pitches nose up, it can reach stalling AoA. If the plane pitches nose down, it can reach zero AoA, or even AoA less than 360 degrees, in which case the lift is respectively 0g or negative.

Obviously, the airflow around any wing, whatever AoA, will be affected by shockwaves by the time freestream velocity reaches Mach 1,0.

[DC-Mainliner]

Well, it won´t be in equilibrium any more than accelerated stall is an equilibrium condition. If the plane pitches nose up, it can reach stalling AoA. If the plane pitches nose down, it can reach zero AoA, or even AoA less than 360 degrees, in which case the lift is respectively 0g or negative.
Obviously, the airflow around any wing, whatever AoA, will be affected by shockwaves by the time freestream velocity reaches Mach 1,0.

Freestream velocity at mach 1 is not worth discussing in trans sonic aircraft. Much, much more will occur on the wing and aircraft structure itself at much lower speeds. Depending on the nature of the wing design, Mach Crit can occur anywhere from M0.7 to 0.75 free stream speed and onward where local flows produce shock waves. Since cruise speeds are in excess of mach crit on trans sonic aircraft, quite a lot goes on within the aircraft for the majority of the time it is airborne.

At the risk of sounding redundant, look at J.P. Davies nice little book, "Handling the Big Jets" where pages 109 through 140 apply. Chapter 6 has great bits too. He does a realy good job, as a test pilot on all sorts of aircraft, relaying the essence of a stall within comprable and valid circumstances. It is well tested and valid data.

Then check out "Fly the Wing" by Jim Webb, chapter 12, "Stalls" (basic) and then compare this to chapter 4, "High Altitude Machs". Chapter 4 is good stuff for jet aircraft, the subject of most here at pprune.

Some of what you say is valid, but it doesn't pertain to the discussion. The above two sources are well respected answers to issues raised in this discussion.

[Craggenmore]

With higher altitude (= reduced air-density), you need a higher TAS to have the same IAS, thus the higher you fly, the higher will be your TAS to carry a certain weight (good for traveling).

A high speed stall is when the airflow over your wing reaches the speed of sound (M = 1 at this point of the wing) due to acceleration around the wings shape. A shockwave emerges and the airflow separates. [Note: you might have heard of the term Mcrit or critical machnumber. This is the aircrafts theoretical speed when its airflow over the wing reaches M = 1]

The higher you fly the lower is your TAS needed to reach M = 1 (due to lower temperature). The higher you fly, the easier you end up in a high speed stall.

So now we combine low and high speed stall: at a certain altitude, the minimum TAS you need to avoid a low speed stall will be the maximum TAS you can fly before entering a high speed stall. This point is called "coffins corner" and theoretically the highest altitude an aircraft can fly.

Thanks for that studi. Ive been on the look out for a consice 'coffin corner' explanation for years! :)

Not sure, but I think Concorde used to pump fuel fore an aft to help control it too.
Yes it did. :)

[ICEMAN757200]

There is no "high Speed stall" you may be confusing it with flow separation due to shock waves or accelareted stall due to an increase of the load factor(wich happens for example if you pull too hard on the ctrl column at high airspeed).
A stall is a condition in aerodynamics (http://en.wikipedia.org/wiki/Aerodynamics) and aviation (http://en.wikipedia.org/wiki/Aviation) where the angle between the wing's chord (http://en.wikipedia.org/wiki/Chord_%28aircraft%29) line and the relative wind, defined as the angle of attack, exceeds the critical angle of attack. This angle is typically 12 to 15 degrees for many subsonic airfoils. The critical angle of attack is the angle of attack on the lift coefficient (http://en.wikipedia.org/wiki/Lift_coefficient) versus angle-of-attack curve at which the maximum lift coefficient occurs, and it defines the boundary between the wing's linear (http://en.wikipedia.org/wiki/Linear) and nonlinear (http://en.wikipedia.org/wiki/Nonlinear) regimes. Flow separation begins to occur at this point, decreasing lift, increasing drag (http://en.wikipedia.org/wiki/Drag_%28physics%29), and changing the wing's pitching moment (http://en.wikipedia.org/wiki/Pitching_moment). A fixed-wing aircraft during a stall may experience buffeting, a change in pitching moment (nose up or nose down depending on tailplane (http://en.wikipedia.org/wiki/Tailplane) configuration), and changes in most stability derivatives. Most aircraft are designed to have a gradual stall with characteristics that will warn the pilot and give the pilot time to react. For example an aircraft that does not buffet before the stall may have a stick shaker (http://en.wikipedia.org/wiki/Stick_shaker) installed to simulate the feel of a buffet by vibrating the stick fore and aft. The critical angle of attack can at 1g only be attained at low airspeed. Attempts to increase the angle of attack at higher airspeeds merely cause the aircraft to climb. Consequently at 1g, stalling occurs only when the aircraft is flying slowly.

[Dani]

Iceman, that's exactly what's the definition of high speed stall: flow separation because of mach shock waves.

Studi is basically right. Dear old D.P. "Deep" Davies ("handling the big jets") would have answered: Stall speed is always at the same EAS if you disregard some compressibility effects for high mach numbers.

Dani

[ICEMAN757200]

That´s what I know as Shock Stall, We use to think in airspeed terms when we speak about stalls but stalls are related to AOA.An airfoil can stall at any speed but the Critical AOA will be always the same for a given airfoil Lim V tends to + infinite in L = (1/2) d v2 s CL then L increases to infinite.

[ICEMAN757200]

The higher the altitude the lower the EAS at wich Mcrit is reached so the answer for Q1 is High Altitude.An increase in aircraft weight(or load factor) will result in a decrease in Mcrit thus the answer for Q2 is the same, high altitude ,but Mcrit will be reached at the lowest altitude than for any other gross weight lower than max .

[chornedsnorkack]

The critical angle of attack is the angle of attack on the lift coefficient (http://en.wikipedia.org/wiki/Lift_coefficient) versus angle-of-attack curve at which the maximum lift coefficient occurs, and it defines the boundary between the wing's linear (http://en.wikipedia.org/wiki/Linear) and nonlinear (http://en.wikipedia.org/wiki/Nonlinear) regimes. Flow separation begins to occur at this point, decreasing lift, increasing drag (http://en.wikipedia.org/wiki/Drag_%28physics%29), and changing the wing's pitching moment (http://en.wikipedia.org/wiki/Pitching_moment).

However, while one can expect that flow separation would increase drag and decrease lift, this does not mean maximum lift coefficient! The lift coefficient comes from many parts of airfoil - including the high-pressure area under the wing - so increase of lift there might more than compensate loss of lift due to flow separation...

[ICEMAN757200]

At a constant AOA , the increase of CL as speed increases from about M 0.4 into the low end of the transonic region gives a steeper lift ccurve slope, i.e. the change of CL per degree of AOA will increase.However, because of early separation resulting from the formation of shockwave,CLmax and the stalling angle will be reduced.