Does lift go up in Ground Effect?

[shdw]

http://www.airlinepilotforums.com/fl...rainers-4.html

Post #32 is where he starts to answer this for us...or does he?

[tgrayson]

Originally Posted by shdw

http://www.airlinepilotforums.com/fl...rainers-4.html

Post #32 is where he starts to answer this for us...or does he?

Cubdriver seems to be using a different definition of lift; he seems to be using it to mean the aerodynamic force generated by the wing, some of it in the vertical dimension and some in the horizontal. By that (nonstandard) definition, there will be a net reduction in lift, because the pilot must reduce the AoA in ground effect in order to keep from ballooning.

Absent any AoA reduction in ground effect, there will be an increase in the effective AoA due to the reduction of wing tip vortices; without pilot action, the aircraft will balloon. Cubdriver seems to acknowledge this.

Note that once the pilot controls the ballooning, lift is restored to a lift=weight relationship and all that remains is a reduction in drag, which is the most significant result of ground effect. That's why you float.

I'm not sure there is any real disagreement in that thread.

[shdw]

Originally Posted by tgrayson

I'm not sure there is any real disagreement in that thread.

It bread the misunderstanding of Cl max and others perceiving what he said as lift doesn't go up and Cl can't change. Here:

There seems to be a misconception that an increase in a specifc C/L is a total increase in lift... this is not true... total lift can only be increased when C/Lmax is increased

It is a good post to learn from, he gets things close to right but hammers in certain points that leave the students mind wondering the wrong way.

[tgrayson]

Originally Posted by shdw

There seems to be a misconception that an increase in a specifc C/L is a total increase in lift... this is not true... total lift can only be increased when C/Lmax is increased

I don't have the context for that, but I agree that an increase in Cl doesn't necessarily imply an increase in lift. As we slow the airplane in level flight, Cl goes up, but lift remains constant. But at a constant airspeed, an increase in Cl does produce an increase in lift, although it's temporary, of course, because airspeed will start bleeding off to correspond to the cl.

As for "total lift can only be increase when C/Lmax is increased", I'm not sure what that means. What is "Total Lift"? Are they referring to the aerodynamic force generated by the wing?

This is why using standard terminology is so important. Lift is defined to be the aerodynamic force perpendicular to the flight path and drag is defined to be the aerodynamic force parallel to the flight path. To use something like "Total Lift" to be the vector sum of those forces is a bit confusing, particularly when they don't explain how they're using it.

[shdw]

Here are the other two maybe this will help:

Page 5:

http://www.airlinepilotforums.com/fl...rainers-5.html


He started a new thread on it:

http://www.airlinepilotforums.com/te...continued.html

[tgrayson]

Originally Posted by shdw

Here are the other two maybe this will help:

ryan1234 started off okay, but then went off into Looney Tunes land. You won't reach any understanding with him.

[shdw]

Originally Posted by tgrayson

ryan1234 started off okay, but then went off into Looney Tunes land. You won't reach any understanding with him.

Cubdriver agreed with him earlier on, I fear I won't convince either of them. Why the narrow minds...

[at1024]

How would you change CLmax? Every time I've ever looked at a CL graph it's always a fixed angle of attack the generates the greatest amount of lift.

[shdw]

Originally Posted by at1024

How would you change CLmax? Every time I've ever looked at a CL graph it's always a fixed angle of attack the generates the greatest amount of lift.

Cl max can vary with airfoil design. I don't have access to the formula, I have one advanced book that likely has it but I wouldn't have the slightest idea where. Basically Cl max is figured out for each individual wing (right?) during testing. If I remember right it can vary 14-17 degrees and I have read 18 but I think you get the point, wing design varies Cl Max.

[at1024]

Originally Posted by shdw

Cl max can vary with airfoil design. I don't have access to the formula, I have one advanced book that likely has it but I wouldn't have the slightest idea where. Basically Cl max is figured out for each individual wing (right?) during testing. If I remember right it can vary 14-17 degrees and I have read 18 but I think you get the point, wing design varies Cl Max.

Oh ok I misunderstood the quote, I thought he was talking about CLmax for one certain airfoil design and configuration. Obviously flaps would change that.

[Doug Taylor]

Does lift go up in ground effect?

Well, I generally try to spend as little time in ground effect as possible!

[tgrayson]

Originally Posted by shdw

they are just the other factors that cause lift: impact, bernoulli, and deflected air.

BTW, this quote of yours is not correct. There is no "impact" air, and Bernoulli and deflected air are flip sides of the same coin. You can count one or the other, but not both. The theoretical determination of lift on an airfoil is based on the sum of the pressure differences around the airfoil, and those pressures will be related the velocity of the air via the Bernoulli equation. This is uncontroversial and the very basis of aerodynamic theory. You might look up some of our other "What causes lift?" threads.

That's one of the reasons I suggested that ryan1234 had some mistaken ideas.

[shdw]

Originally Posted by tgrayson

The theoretical determination of lift on an airfoil is based on the sum of the pressure differences around the airfoil, and those pressures will be related the velocity of the air via the Bernoulli equation.

So what exactly is Cl measuring than?

Let me see if I have this right then, impact lift is nothing more than a pressure applied at a point on the airfoil. Lift is a measure of that application of pressure which is sometimes confused as a separate entity?

The same would hold true for downwash or deflected air? Where ever the pressure is originally applied to the airfoil is what is measured, not the force it exerts downward. Am I on the right page?

[tgrayson]

Originally Posted by shdw

So what exactly is Cl measuring than?

Probably one thing that ryan1234 had right...Cl is just a measure of how effective an airfoil shape is in generating an up force....by whatever means. The actual lift on an airfoil is measured, and then various environmental factors are divided into the force value in order to get a dimensionless number. This number is useful because you can then combine it with changed environmental factors and get an accurate lift value.

Let me see if I have this right then, impact lift is nothing more than a pressure applied at a point on the airfoil. Lift is a measure of that application of pressure which is sometimes confused as a separate entity?

On any airfoil generating positive lift, there is a large reduction in pressure over the top of the airfoil. At low AoA, the bottom of the airfoil will also experience a drop in pressure. Thing is, the top is much lower than ambient and the bottom is only slightly lower than ambient. There is a net pressure difference in the "up" direction....this is lift. Is it caused by the top or the bottom? Meaningless question. It's caused by the difference between the two.

As the AoA increases, the bottom of the airfoil starts to incur positive pressure, but the top gets even more negative. Even in this case, the bottom doesn't have "impact" pressure bacause the airflow never "impacts" the airfoil. It's diverted around the airfoil, generating the stream lines that you've seen zillions of times.

The most ludicrous thing that ryan1234 said that made me think he's just yanking your chain is that the pressure on top of the airfoil increases with AoA, rather than reduces. That is so absurdly wrong that I just can't believe he's serious.

The same would hold true for downwash or deflected air? Where ever the pressure is originally applied to the airfoil is what is measured, not the force it exerts downward. Am I on the right page?

I'm not sure what you mean by this. For the airflow to be deflected requires a force acting on the air....this force is due to the pressure differences around the airfoil, so the pressure differences must already exist for the airflow to be deflected. The deflected flow is a result of lift, not the cause, but you probably could calculate the quantity of lift based on the quantity of air deflected.

[shdw]

Originally Posted by tgrayson

The deflected flow is a result of lift, not the cause, but you probably could calculate the quantity of lift based on the quantity of air deflected.

What I meant was that the deflected air exhibits some form of pressure on its point of impact (really the point of deflection around the wing) and that pressure is calculated for? That help.

Cl is basically calculating the pressure differences than? But it is a theoretical number, not an absolute? I say this because the rest of the lift formula is air density, speed, and wing area. Cl would be the only one capable of attempting to calculate the pressure differences that I can see.

[tgrayson]

Originally Posted by shdw

What I meant was that the deflected air exhibits some form of pressure on its point of impact (really the point of deflection around the wing) and that pressure is calculated for? That help.

Let's clarify some terms. What most people mean when they say "deflected air" is the downward motion imparted to the air via the wingtip vortices. However, this downward motion is the result of the pressure distribution already present around the airfoil and so can't be the cause of it.

Now, when you say "impact" pressure, this is nothing more than the Bernoulli equation in action. If you fired a stream of fluid at a brick wall, the stream would hit the brick wall and stop. Since V drops to 0, the dynamic pressure would fall to zero, too, but the static pressure would rise to compensate, just as the Bernoulli equation expects. This is the "stagnation point". The fluid that folllows will flow around this area of high pressure, producing the gently curving stream lines. As the air flows from the high pressure stagnation point, the pressure drops.

The main problem with using the terminology "impact" lift is that it calls to mind the idea of a stream of air bullets impacting the bottom of the airfoil over its entire surface; this metaphor fails for two reasons: 1) the actual impact is limited to the stagnation point, not the entire bottom of the airfoil, and 2) it doesn't take into account the reduced pressure on the top of the airfoil, which is essential to producing large quantities of lift.

Cl is basically calculating the pressure differences then?

Not really. It's just a number that relates lift to dynamic pressure and wing area. You can plug that number into other formulae and get actual physical quantities, as long as you supply some other data along with it. Sorta like glide ratio or propeller efficiency.

But it is a theoretical number, not an absolute? I say this because the rest of the lift formula is air density, speed, and wing area. Cl would be the only one capable of attempting to calculate the pressure differences that I can see.

I don't know what you mean regarding the "theoretical" vs "absolute". The lift equation is mechanism-agnostic, meaning that it has no assumptions regarding how lift is produced, only that it exists. However, since we *know* from external sources that lift is generated via pressure differences, we can draw the conclusion that for 2,000 lbs of lift, the net pressure difference between the top and bottom of the airfoil total to be 2,000 lbs.

[shdw]

Tgray, or some other dynamics guru please take a look at this: http://www.airlinepilotforums.com/te...tml#post655599

Specifically the last two paragraphs. I have been rereading this over and over and it still doesn't make sense, I wonder if anyone can shine some light on this in more lemans terms, thanks.

[SteveC]

Originally Posted by shdw

Tgray, or some other dynamics guru please take a look at this: http://www.airlinepilotforums.com/te...tml#post655599

Specifically the last two paragraphs. I have been rereading this over and over and it still doesn't make sense, I wonder if anyone can shine some light on this in more lemans terms, thanks.

Why not ask over there?

[tgrayson]

Originally Posted by shdw

Tgray, or some other dynamics guru please take a look at this: http://www.airlinepilotforums.com/te...tml#post655599

Specifically the last two paragraphs. I have been rereading this over and over and it still doesn't make sense, I wonder if anyone can shine some light on this in more lemans terms, thanks.

That post is your own post. Which of your own paragraphs confuse you?

(Ryan, BTW, is confusing absolute AoA with effective AoA.)

[shdw]

Originally Posted by tgrayson

That post is your own post. Which of your own paragraphs confuse you?

(Ryan, BTW, is confusing absolute AoA with effective AoA.)

Also would you mind making some attempt to explain this: http://www.desktopaero.com/appliedae...erivation.html

That is what I don't get.

Steve, I did and it has been almost a week with no reply so I figured they didn't know or didn't care.

[tgrayson]

Originally Posted by shdw

That is what I don't get.

I know very little about Fourier analysis, so I can't give you a blow by blow. But the basic idea is that of a "vortex sheet", an infinite number of vortices that have an axis of rotation along the wing span. The math sums up their effects. The up and down portions cancel, but the part tangential to the wing do not, resulting in the velocities across the top and bottom of the airfoils.

John D. Anderson's "Fundamentals of Aerodynamics devotes a dozen or so pages to the math, if you're interested. But at the end, he says

Therefore, the local velocities tangential to the airfoil surface are equal to the local values of γ (gamma). In turn, the local pressure distribution can be obtained from Bernoulli's equation.

The significance of the above depends on his previous emphasis, such as

In Section 1.5, we emphasized that the resultant aerodynamic force on a body immersed in a flow is due the the net integrated effect of the pressure and shear stress distributions over the body surface. Moreover, in "Section 4.1, we noted that lift on an airfoil is primarily due to the surface pressure distribution, and that shear stress has virtually no effect on lift. It is easy to see why. Look at the airfoil shapes in Figures 4.12 and 4.13, for example. Recall that the pressure acts normal to the surface, and for these airfoils the direction of this normal pressure is essentially in the vertical direction, that is, the lift direction. In contrast, the shear stress acts tangential to the surface, and for these airfoils, the direction of this tangential shear stress is mainly in the horizontal direction, that is, the drag direction. Hence, pressure is the dominant player in the generation of lift and shear stress has a negligible effect on lift. It is for this reason that the lift on an airfoil below the stall can be accurately predicted by inviscid theories such as that discussed in this chapter.

Rather than debate theory with your buddies, you may find experimental data more user-friendly. I highly recommend "Theory of Wing Sections," which has lots of good stuff and is cheap.