That airfoil sure looks thick!

Well, in the case of the Windwalker, it’s not any thicker than the wing of most other planes--about 7 inches at the root. However, the chord is considerably smaller than most (39 inches at the root), so the airfoil is proportionally much thicker. This is usually expressed as the ratio of thickness to chord. In our case, its 18%, whereas most other planes run between 12% and 15%.

The Thunderbug airfoil is 15% thick and the new Cosman LSA airfoil is again 18% thick.

It's interesting to follow the evolution of thinking about airfoil shape and thickness. The Wright flyer had a two-surface airfoil just thick enough to get some curved ribs inside. The top and bottom skins followed each other with only about an inch of rib between. The curvature was modest, and the airfoil was primitive by modem standards. However, their approach was a significant step forward from the flawed theories and tabulated data of their predecessors. They were the first to build a wind tunnel and attempt to extract meaningful, repeatable measurements of airfoil lift and drag values. As a result of their work, they realized they needed much more wing area than others theorized. They eventually were able to quite accurately predict the areas, weights, drag and required power for their first successful configuration.

During those early years, the underlying theories were quite crude. Wings lift up by pushing air down. Almost any shape does an OK job. But the devil, as they say, was in the details. How much drag is created for each unit of lift? What is the maximum amount of lift that can be generated before it stalls? How much tail effect is needed to counteract wing pitching moments, and how do these moments change with angle of attack? How does the airfoil "stall"--gently, or suddenly? And so while almost any airfoil shape will produce useful lift under cruise conditions, most were given to uncivilized behavior at other speeds and angles of attack.

Aviation advanced by leaps and bounds during the first World War. Suddenly the behavior of the airfoil at the extremes of the envelope became very important. You needed to out climb, out turn, out run, and out dive you opponent. Thin airfoils with sharp leading edges tended to stall suddenly at relatively shallow angles of attack and relatively low maximum lift coefficients. As speed increased, their pitching moments increased rapidly and had to be counteracted with increasing downloads provided by the tail. Stick forces increased rapidly with speed, and they were prone to sudden unannounced stalls with unpredictable snaps to one direction or the other when too much lift was demanded at high speed. At the same time designers began to realize the large amounts of drag produced by the spiderweb of flying wires, struts and braces trussing up these early biplanes.

Monoplanes, cantilever (internally braced) wings and thick airfoils all arrived at about the same time. Early adopters of the new paradigm discovered that thicker airfoils provided higher maximum lift, lower stall speed, gentler stall behavior, and (counter-intuitively) lower drag. It was now possible to design an airfoil that behaved in a consistent, predictable fashion over a wide range of speeds and lift coefficients. Theories were developed during this era that suggested families of airfoils whose behavior could be correlated to characteristics of their thickness profiles and camber lines. The National Advisory Committee on Aeronautics (NACA, forerunner of NASA) and its counterparts in Europe designed and tested large families of airfoils based on these composition rules. airfoils developed during these studies still comprise the majority found on small aircraft today. And generally, these airfoils have proven to be good, predictable "work-horse" sections without noticeable bad habits.

But human intuition is sometimes at odds with the laws of physics. Designers have always fought this constant urge to believe that thin is faster than thick. One almost universal mistake this caused was the use of a thinner section towards the tip of the wing, than at its root. The result was almost always a wing that stalled tip-first, with an inevitable loss of roll control and a departure into an incipient spin. This behavior was eventually tamed by copious amounts of wash-out (twisting the tip to fly at a lower angle of attack), or leading-edge slots, or cobbled-on droop cuffs, or a combination of these. The problem is right there in the numbers, screaming out from the theories and equations, and yet is still occasionally ignored even today. Thinner sections always have lower maximum lift coefficients, lower stalling angles of attack, and more sudden departures from attached to separated flow. The theory suggests that if the wing planform is tapered, the tip section should be thicker, not thinner than the root section, if tip stall is to be avoided without resorting to these crutches.

But surely thick sections have more drag than thin ones, right? Well, there is a component of section drag that is related to profile thickness, but it is not the dominant factor, and in most cases is not even a significant factor. When the entire mission profile is taken into account, the thicker section wins because it has less drag in the high-lift configuration (take-of and climb), more lift under all conditions (hence the wing can have less area). and offers a lower landing speed than a similar-sized wing of thinner section.

One interesting specific example is the Nemesis formula racer, which holds all the records for this class of aircraft. In an arena where most racers had wings with sections as thin as 6%, and averaging 10%, the Nemesis has a 14% section! Another example is the Rutan Boomerang asymmetric twin, which has 16% sections. In fact, thick sections are increasing in popularity as designers realize that they can be made stronger and lighter, hold more fuel, perform better, and don't really suffer any measurable penalty in top speed. Only when you venture into the transonic region (indicated airspeeds above 300 kts) does the balance begin shifting back toward thin sections.

The particular choice of an airfoil depends on several other factors. The most important is the design cruise lift coefficient. The plane is intended to cruise at a certain speed. It has a design gross weight. Once the planform is nailed down, the wing area is known. This allows us to compute what the required lift coefficient (and corresponding angle of attack) will be in the cruise configuration. An airfoil should be selected that delivers this lift coefficient at its design angle of attack, where its drag is minimized. If the wrong airfoil is selected, it will be either flying at too low or too steep an angle of attack in the cruise configuration. In either case, drag is higher than with a more proper selection of airfoil.

If the wing can be made suitably smooth, it is worthwhile to consider a laminar-flow section. These airfoils decrease drag by delaying the onset of turbulent flow farther back along the section. In general, this is done by moving the point of maximum thickness aft. Typical turbulent sections from the WWII era have their thickest point at about 25% chord, while laminar sections push this back to sometimes as far as 40%. The downside is that the leading edge is somewhat sharper (as is the stall), and if the airfoil is flown at an angle of attack too far away from its design angle, you get premature transition to turbulent flow anyway. If you look at the drag-versus-angle-of-attack plot for a laminar flow section, you will see that the region of minimum drag is a shallow "bucket" in an otherwise conventional drag profile.

One additional tweak can further improve the drag characteristics of a laminar-flow section, and that is to contour the airfoil for a rather constant curvature along the top surface, and a concave "cusp" on the bottom aft portion of the wing. This allows you to curve the camber line suddenly downward towards the trailing edge where the pressure gradients are positive (i.e. air pressure is increasing as you approach the trailing edge). Doing the same thing with the top surface, where the pressure gradients are not as favorable, would result in flow separation, a particularly nasty form of airflow transition that results in sudden, large increases in drag and sudden, large changes in lift and pitching moment.

But this tweak has a dark side. The sudden change in camber line creates a large lift component acting on the aft portion of the airfoil. This shows up in two ways. Moveable controls (flaps and ailerons) get heavy quickly with increasing speed, and the overall pitching moment of the airfoil is significantly larger in the nose-down direction. Going to powered flaps (rather than manual ones) solves some of the problem, and most designers go back and fill in the cusp on the underside of the ailerons to tame stick forces. But the larger pitching moment can only be overcome by a larger stabilizing downforce generated by the tail, and this looks just like more weight to the wing. When you factor in this pitch-trim drag penalty, the benefits of the cusp are generally nulled out.

So where did we wind up? Our section is thick because it provides a high maximum lift coefficient. can be made very strong without getting too heavy, and has lots of room for controls and fuel. We chose a laminar section based on Harry Ribblet's modification of a NACA 64 series, with the thickest point at about 37%. He found that by fixing a minor compositional problem with the very leading edge, and flattening out the cusp, overall performance was not degraded but the section was easier to build and didn't have the pitch-trim and control force problems of the original 64 series section. Our selection of airfoil is one reason we expect to have a much lower, tamer, safer landing speed than might otherwise be possible.

Mike Cosman - Aircraft Designer