Table Of Content
The differences delineate the developments of the shock waves and other compressible flow patterns; the shock waves are the darker zones produced by the schlieren effects. This latter behavior directly results from the relatively thicker boundary layers on the airfoil surfaces at these lower Reynolds numbers. Furthermore, the development of a laminar separation bubble on the airfoil is often a source of these non-linear characteristics. In these cases, the airfoil’s lift behavior becomes much more challenging to generalize as a function of the angle of attack, i.e., the relationship in Eq. Where means conditions at “infinity” or just far away from the wing where there is an undisturbed “free-stream” flow.
Zenith Aircraft Co. CH-750
The center and pressure (as well as the aerodynamic center) can be determined if the lift and moment coefficients are known about any other point, the 1/4-chord often being used as a reference point, i.e., the values of and are available. The best way for the student to understand the process is to work through an example using actual airfoil measurements, which are given in the table below for a NACA 0012 (symmetric) airfoil. Nevertheless, a close inspection of the image shows a small dark line near the 25% chord, suggesting a weak shock wave and that the critical Mach number, , is being reached, i.e., the flow about the airfoil becomes locally supersonic. By , a series of small shocklets (which are mild shock waves) can be seen between 10% and 30% of the chord, confirming the existence of a supersonic pocket, and by , a weak shock wave has formed. The lifting properties of the F-16 wing are poor relative to the other two aircraft.
Representative Force & Moment Coefficients
Schemes have been devised to define airfoils – an example is the NACA system. An example of a general purpose airfoil that finds wide application, and pre–dates the NACA system, is the Clark-Y. Today, airfoils can be designed for specific functions by the use of computer programs.
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Their profiles are such that they are conducive to maintaining an extensive run of laminar flow over the leading-edge region, thereby lowering skin friction drag, at least over a range of angle of attack limited to low lift coefficients. Also shown in the figure above are results for a NACA 63-series laminar flow airfoil, which is a design that maximizes the extent of the laminar boundary layer over the leading edge of the airfoil, hence reducing skin friction drag. These airfoils (with very smooth surfaces) tend to produce low drag “buckets” where the drag is relatively lower, but typically only over a limited range when the airfoil operates at low angles of attack and low lift coefficients.
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Integration of Pressures & Shear Stresses
Again, notice the use of free-stream dynamic pressure as a reference pressure. The concepts of center of pressure and aerodynamic center are used routinely in aerodynamic analysis, and it is essential to understand their differences. They are frequently confused in practice, though they are quite different.
My airfoils
The number denotes the thickness-to-chord ratio in percent of the chord; e.g., a NACA 0015 has a 15% thickness-to-chord ratio, which means . The nose curvature or radius, , must also be formally located on the profile and is obtained with an inscribed circle. The resulting point then becomes the origin location for the leading-edge nose circle. Notice that the center of the nose circle will not lie on the mean camberline.
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NACA Method of Drawing Airfoil Shapes
The geometric shapes of these airfoils are different from those used on most airplanes and are designed to have a point of maximum thickness close to mid-chord. This shape produces a favorable pressure gradient over the leading edge, encouraging the boundary layer to be smooth and laminar for longer. The laminar flow produces less skin friction and so less drag on the airfoil. The downside is that such airfoils typically produce lower values of maximum lift coefficient, i.e., a stall occurs at lower angles of attack.
Principles of Flight
The resulting flow solution can then be used to calculate various properties around the airfoil, including local pressure, local Mach number, etc. CFD methods can also be used to design the shape of an airfoil to obtain a specified level of performance. However, this tends to be lengthy because of its iterative nature and slow numerical convergence.
When the airfoil is in action, the flow of air moving upwards over the airfoil is an upwash, and air moving downwards is downwash. The airfoil generates lift from the pressure distribution between the upwash and downwash by turning the incoming air downwards. The subtle differences in the shape of airfoils are really not that important for most model airplanes. Sure, if you’re going for an endurance record, you might want to think about it more, but in most cases fine margins are imperceivable.
We can use this information to determine the maximum lift coefficient that the airfoil + slat + flap combination produces. Not long after Phillips, Gustave Eiffel conducted more experiments using a wind tunnel of the open-return (single passage) type. The figure below shows a relatively rapid evolution of airfoil shapes tailored to aircraft applications between 1908 and 1944, with the thin and highly cambered airfoil sections used on early airplanes being relegated to history. Understanding the aerodynamic characteristics of two-dimensional airfoil sections is a prerequisite to understanding the characteristics of finite wings. While airfoil characteristics can be predicted, the most reliable results still come from measurements made in the wind tunnel, especially near the maximum lift or into the stall.
For example, airfoils for use on the wings of low-speed airplanes are generally thicker (in terms of their thickness-to-chord ratio) and have more surface curvature or camber. Airfoils for high-speed aircraft, especially for supersonic flight, are much thinner with more pointed leading edges and much less camber. The well-rounded, cambered airfoil sections that are well-suited to subsonic flight speed are generally inappropriate for high-speed and supersonic flight. Supersonic airfoils are distinctive in their geometric shapes in that they are thin (i.e., have a low thickness-to-chord ratio) with sharp leading edges. Supersonic airfoils generally have thinner sections formed of either angled planes called double-wedge airfoils or opposed circular arcs called biconvex airfoils, as shown below. The sharp leading edges on supersonic airfoils prevent the formation of a detached bow shock in front of the airfoil, which is a high source of drag called wave drag.
Supersonic airfoils are much more angular in shape and can have a very sharp leading edge, which is very sensitive to angle of attack. A supercritical airfoil has its maximum thickness close to the leading edge to have a lot of length to slowly shock the supersonic flow back to subsonic speeds. Generally such transonic airfoils and also the supersonic airfoils have a low camber to reduce drag divergence. Modern aircraft wings may have different airfoil sections along the wing span, each one optimized for the conditions in each section of the wing. Another set of NACA airfoils that have seen some use on various aircraft is the six-digit series. These airfoils were designed to achieve lower drag, higher drag divergence Mach numbers, and higher maximum lift coefficients.
By , both shocks have reached the trailing edge and become significantly bifurcated from their interaction with the relatively thick boundary layer. A summary of these preceding observations is shown below as a schematic for better clarity. Further increases in maximum lift coefficient may be possible using different types of flaps, such as slotted flaps, Fowler flaps, and double-slotted or triple-slotted flaps. In conjunction with leading edge slats, significant stall speed reductions can be achieved with larger airliner types of airplanes.
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