I have heard of the faster flow slower flow theory but what about the symmetrical airfoil's. The flow of air should be traveling at the same speed so how is lift generated there?

to get lift there is just a little incidence in it ore the nose of the plane is up vary slightly to have more airflow under the wing then on top.
Seth

For a fully illustrated and correct explaination of how lift is created see:
http://www.monmouth.com/~jsd/how/
It takes Bernnouli, Newton, Coanda and circulation laws and theories, all working together, to full explain how lift is produced.

When a wing with a symmetrical airfoil is at zero degrees angle of attack, it produces no lift. The stagnation point where the flow splits to go over and under the airfoil is at the leading edge of the airfoil.

As the angle of attack increase from zero, the stagnation point moves down and aft along the bottom near the leading edge of the airfoil so that the flow over the bottom has to negotiate less of the curvature near the leading edge and the flow over the top has to turn back and flow around the leading edge before it can negotiate the top of the airfoil. The greater the angle of attack the farther the stagnation point moves and the greater the coefficient of lift. This behaivor can be clearly seen in smoke equipped wind tunnels. The flow over the top of the airfoil was greatly accelerated as it negotiated the small radius of curvature near the leading edge (Bernnouli).

The flow over the wing tends to cling to the wing surface rather than seperating (Coanda).

If you subtract the air velocity before encountering the wing, from the velocities of flow in the vicinity of the wing, you have a circulatory flow left. This circulation, or vortex is bound to the wing along the wing span but at the wing tips it turns aft to become the tip vortices that feed the sides of the downwash sheet behind the wing (circulation theory).

When the flows combine again at the trailing edge, the flow from the top is going faster than the flow from the bottom, resulting in a net deflection of the flow downward so that downwash is produced. The air particles traveling over the top beat the particles over the bottom in the race to the trailing edge. The flow well ahead of the wing had no vertical component but the flow behind the wing does and this constitutes a vertical acceleration of the air by the wing. The lift force is equal to the mass of the downwash times its vertical acceleration (Newton's Law).

Does this question sound familiar Ollie ?

Make sure you go to that link Ollie posted. Not only will it answer your question, but it will provide you with a lot of aerodynamic knowledge.

Ollie, I've never heard the description you used for Newton's law before. I always understood the airfoil more like a bent tube. Air goes straight in, and comes out at an angle. The net mass flow vector is your lift. I have also read studies of thin simple plate airfoils with knife edges. In these cases, there is little to no acceleration due to differences in curvature or path. But, they still produce lift as a function of attack angle. Any insight here?

OK, I read the link (some of it). The flat plate behaves like a regular airfoil (short story). This is a very good link. I'll read it on my lunch break for the next few years. Meanwhile, my pessimistic engineering mind will question every claim.

Ilikeplanes

The description that Ollie gave works the same for a flat plate at an angle of attack. What happens with these (while at a small angle of attack) is the following: The forward stagnation point is once again slightly below the leading edge, just like a normal airfoil. That means from just above the stagnation point, air is trying to move forward and around the leading edge, then over the upper surface of the airfoil. However, the leading edge on a flat plate presents a sharp corner, which the air cannot navigate. The air separates locally, and form a little bubble of separated flow just behind the leading edge on the upper surface of the plate. Air approaching from in front of the airfoil and above will now flow over this bubble, reattach behind the bubble on the upper surface of the plate and flow to the trailing edge the same way as on a normal airfoil. The overall effect is a flow field that looks almost the same as for a normal airfoil; the flow still accelerates significantly over the top of the airfoil and especially over the leading edge region, and lift is still produced in the same way as Ollie described. The change in momentum due to the downwash is also still present, resulting in a force according to Newton's laws.

However, as the angle of attack is increased, the separated region increases quickly in size until the flow does not reattach anymore behind it and the plate stalls. This happens at a much lower angle of attack than a "proper" airfoil, which makes the flat plate a less efficient shape for producing lift than a normal airfoil.

It is possible to curve the plate a little, giving it camber which will increase the maximum angle of attack a little but now the lower part of the airfoil will stall again due to separation on the bottom of the airfoil at low angles of attack. In general, thin airfoils tend to have a smaller angle of attack range than thick airfoils.

I highly recommend the website Ollie pointed out - it is a very good introduction to aerodynamics.

Regards,
Ben

The confusion stems for many generations of elementary text books that so simplified the explaination of how lift is produced that people started arguing about which theory is correct when, in actuality, all four therories are correct, necessary and do not contradict each other in any way. These oversimplifications have blinded rather than enlightened their readers.

A flat plate is a special case of a symmetrical airfoil. When a flat plate is rotated from zero to a positive angle of attack the stagnation point moves from the leading edge, down and aft along the lower surface so that the flow over the top must first change direction almost 180 degrees before negotiating the leading edge. I know this fact is counter intuitive but it has been experimentally verified.

There is a boundary layer beginning at the surface of the airfoil where the flow begins to shear. The air at the surface does not flow but sticks to the surface (adhesion) and there is some thickness involved before the air gets up to the local velocity (viscosity). When there is no mixing between layers in the boundary layer the flow is said to be laminar. When there is mixing between the layers the flow is said to be turbulent. When a boundary layer transitions from laminar to turbulent the boundary layer thickens. The thickness of the boundary layer varies some along the chord and also with the size of the chord and the airspeed. The boundary layer imparts a virtual thickness to the airfoil. Because of the boundary layer the flat plate can't have a virtual thickness of zero. At model sizes and speeds the boundary layer is much thicker than at full scale. This is part of the reason that thinner airfoils work better at model reynolds numbers and thicker airfoils may work better at full scale. By the time you get down to insect sizes and speeds the boundary layer is so thick compared to the chord that insects have to produce lift by an entirely different mechanism such as clap-snap-twist.

P. S. In my previous post I forgot to mention that the circulation produces an upwash ahead of the wing.

Ben,
Sorry for stepping on your lines. I was writing while you were posting and I didn't see your post in time to keep from duplicating part of what you said.

No problem, Ollie. I don't post here much but I always enjoy reading your comprehensive and also very accurate explanations.

Cheers
Ben

So let's talk about a flat bottom wing. Because the air moves faster on top of the wing and slower on the bottom it created lift because the faster air molecules rise above the slower one's. Like hot air, hot air rises because it moves around faster and makes it's way above slower moving air or colder air?

The classification of airfoils as flat bottomed or semisymmetrical is a very crude way of classifying airfoils. The lift characteristics of an airfoil are closely related to the mean camber line and the drag characteristics are related to the thickness. A thin, flat bottomed airfoil can have less mean line camber than a thick semisymmetrical airfoil making it difficult to draw conclusions about the way the airfoil behaves from such a simple classification.

Over simplification of lift theory led to serious misunderstanding by a lot of people. Airfoil classification simply as flat bottomed or semisymmetrical is an over simplification that can lead to serious misunderstandings too.

As a case in point, a flat plate is a flat bottomed airfoil. The fact that it is also symmetrical tells us that it has no camber of the mean line and the fact that it is flat on both sides implies that it has little thickness but doesn't tell how thick so we don't know much about the drag. The lack of camber tells us that it will perform the same upright as inverted. So, it takes a lot more information than just flat bottomed or semisymmetrical to begin to draw conclusions about how an airfoil will behave.

In the 1930's NACA adopted a four digit airfoil designation system based on the camber of the airfoil's mean line and the airfoil's thickness because these numbers had primary aerodynamic significance. Later, NACA refined the designation system by increasing the number of digits to five so that the location of the high point of the mean camber line could also be included. As laminar flow airfoils were developed in the early 1940's NACA increased the number of identifying digits to six, one of which was a subscript, so that the distribution of thickness along the chord could also be identified by the designation.

As Ollie has said, it's difficult to draw conclusions. For example, a 2% camber airfoil that is 18% thick might act more like a symmetrical airfoil than a 2% camber that is only 6% thick. If you look at symmetrical wings, lift is still created by pressure differential as the stagnation point moves when the angle of attack is increased. A volume of air being displaced by angle of attack alone maybe only contributes 1/3 of the lift of a symmetrical section. Again, that may vary with the thickness as well.

An airfoil and the flow of air around it are a system that produces lift. The four theories that each describe some aspect of the system behavior are integrated by the system and can't exist independently of each other except outside the system. In other words, the downwash can't exist without the circulation, the pressure distribution and the velocity distribution. The existance of each depends on the existance of all the others. So, what sense dose it make to talk about how much of the lift is contributed by which part of the system?

What's angle of attack? I have been hearing about it a lot from this book I got from my naber.

There are two angles of attack, a geometric angle of attack and an aerodynamic angle of attack.

The geometric angle of attack is the angle between two lines. One of the lines is the chord line of the airfoil which is a straight line between the nose of the leading edge and the trailing edge of the airfoil. The other line is the line of motion of the airfoil through the air.

The aerodynamic angle of attack is measured from the angle of attack at which the coefficient of lift is zero. For symmetrical airfoils, the two angle of attack are the same. For airfoils with mean line cambers, the zero lift angle of attack is negative and the greater the camber the more negative the zero lift angle of attack is.

Hi,

I've been following this thread and thanks for all I have learned. I have a practical case that I'd like to ask about.

I've been into 1/2a pylon racing for about three years and scratch build my racers. Wings are 200 sq. in using MH33 foil. Airplanes are conventional hi wing and conventional tail. V tails look great but are harder to build and trim. And I doubt that the speed range we operate in 80-90 mph will afford much benefit for the V tail. Weight ready to fly is 11-12 oz. Wing span is 32-38".

My question is: Given my application, Where would my optimum angle of attack be for the foil I'm using? I have noticed some different flight/glide characeristics between airplanes and have wondered if I'm getting the AOA right. I suspect that the geometric AOA is not the optimal. Your thoughts and comments are solicited and appreciated. Best regards, Jeff.....

Jeff,

When the plane is in level flight in a straight line at a constant airspeed, thrust is equal to drag and lift is equal to weight. The plane assumes a pitch attitude to give the angle of attack (AOA) that makes the lift equal to weight at that airspeed. If you throttle back and maintain altitude the airspeed will be lower, the pitch attitude will be more nose up and the higher AOA will still produce lift equal to weight.

The AOA is the angle that the airfoil chord line makes to the direction of motion.

In a speed model one of the objectives is to minimize the over all drag coefficient so that the available thrust results in maximum speed. The total drag coefficient has many small components. To minimize fuselage drag, the fuselage should be aligned to the flow. To minimize horizontal tail drag it should be lightly loaded. Some combination of thrust offset, CG location, wing incidence and decalage will result in minimizing the parasitic drag coefficient.

Assuming 100MPH and a weight of 12 ounces, the coefficient of lift of the wing is about 0.02. The AOA of the MH32 at a coefficient of lift 0.02 is about -1.5 degrees. That's in the straight away. BTW the minimum coefficent of drag of the MH32 is at a lift coefficient of 0.04. A thinner airfoil of lower mean camber would be better in the straight away but, worse in the turns.

In a tight (bank and yank) turn, the coefficient of lift will be about 1.0 and the AOA of the MH32 will be about 8 degrees plus about another 8 degrees induced AOA for a total AOA in a tight turn of 16 degrees.

Where to set the wing incidence relative to the fuselage centerline, depends on how much of the race is straight away and how much is tight turn. If the centerline of the fuselage is canted more than about 6 or 7 degrees, the drag goes up sharply. Depending on the race course it may be speedier to turn a little wider to prevent speed loss in the turns.

Also, a higher aspect ratio will reduce the induced AOA and the induced drag at high lift coefficients. The fuselage drag is reduced if the total AOA excursion is reduced.

In aircraft design, everything affects almost everything else! When you settle for simple answers you are likely to fall short of the mark.

Are you allowed to use a full flying stab? That is basic sailplane technique. It should minimize horizontal stab drag at any AOA.

For racing, it sounds like you need a foil with very low cd at low cl and low cd at high cl. What you can give up is the medium cl range. High cd in the mid cl range may be acceptable for racing. For most airplanes, however, this would be bad. Just wondering.

The main advantage of an all moving stab is that the decalage can be adjusted without having to change the wing or tail mounting to acomodate a range of CG locations. A well designed and constructed articulated tail does not necessarily have a drag disadvantage compared to an all moving tail.

Ollie,

Thanks for your comments. Now I realize there is some more details that I left out of my original post.

I've been building my racers with everything set to 0-0. Engine, wing and tail. These airplanes turn great! On the upwind you bank to near verticle and yank the elevator and hold until you're looking almost straight at the airplane, let off the elevator and roll slightly right and the airplane will do the entire downwind at near knife edge. Then bring the airplane around two and three. It seems that very little speed is lost in the turn. There is no tendency for the airplane to drop the nose in the turn or snap. In fact these turns are really fun! Especially if your passing someone at the same time! I was really wondering if setting the airplane at zero zero was the best thing to do from a drag perspective. As you may have guessed I could use some more speed. The mh33 foil is a 6.28% thickness. Maybe I could go a little thinner? Rgds,...... Jeff.....

I have wind tunnel measured polars for the MH 32 but not for the MH33. Also, Martin Heperle's web site is discontinued so I don't have access to the MH33 polars as generated by X-Foil. The lowest drag measured in Dr. Selig's series of tests at the University of Illinois was for the S6063 with a little reflex inaccuracy in the trailing edge of the test model. The S6063 has a thickness a hair over 7 percent so i don't think you need to go thinner. If Heperle recommends the MH33 for pylon racers, it should be one of the best. It sounds like you are flying an almost circular path over much of the course. I'm guessing that you are probably not exceeding a lift coefficient of about 0.5. This probably means that the induced AOA is only around 2.5 degrees and the total AOA excursion doesn't put the fuselage in a pitch attitude with hgh drag. My conclusion is that the zero-zero setup is close to optimum.

Ollie,

Too bad about the Heperle's site not being available anymore. He had a ton of good info on it. Thanks again for your comments... I was hoping we'd find another 10 mph in there somewhere! Rgds, Jeff....