Some Aerofoil Characteristics

The NACA series of aerofoils was introduced in Chapter 4. In this appendix, we examine three of these aerofoils in more detail and look at the ways in which changes in cross-sectional shape, particularly camber and thickness distribution, influence their performance. In each case, the aerofoil section is shown, together with a typical distribution of pressure around the lifting sec­tion, the variation of lift with angle of attack and the variation of section drag with lift. The lift and drag are plotted in coefficient form (Chapters 1 and 3). For the pressure distribution, a coefficient form is also used. The pressure coefficient is defined as the local pressure on the aerofoil surface minus the ambient pressure divided by the dynamic pressure (p. 12). Negative pressure coefficients are plotted upwards, so that the upper surface of the aerofoil appears as the upper line on the graph.

The first aerofoil, the NACA 0012 (Fig. A.1), is a 12 per cent thick symmet­rical ‘4 digit’ series aerofoil. It is commonly used for tail surfaces and for wind – tunnel test models. It is also used as the wing section on a number of aircraft including the Cessna 152. This is a popular light general aviation aircraft and the NACA 0012 is used for the outboard wing section. From the graph of lift coefficient against angle of attack for this aerofoil, it can be seen that there is a sharp stall at about 15° angle of attack. The pressure distribution also shows quite a sharp suction peak on the upper surface.

The second aerofoil, the NACA 2214 (Fig. A.2), is used on the centre wing section of the Cessna 152. With a 14 per cent thickness/chord ratio, it is slightly thicker than the NACA 0012 and has some camber. The effect of the camber is evident in the positive lift coefficient that is seen at zero angle of attack. Minimum drag is obtained at a lift coefficient of approximately 0.2, rather than 0.0 for the NACA 0012. The drag is, however, higher for this thicker cambered section and the stall is somewhat more gentle.

The final aerofoil, the NACA 6618 (Fig. A.3), is one of the ‘low drag’ 6 series and is used on the Phantom supersonic fighter. Only the low speed char­acteristics are given here. This aerofoil was designed using a so-called ‘inverse method’. The pressure distribution on the upper surface was chosen to be as flat as possible at a particular ‘design’ lift coefficient and the resulting cross­section was then determined. The flat top surface pressure distribution allows a laminar boundary layer to be maintained over much of the surface, leading to a reduced drag. The laminar layer can be maintained over a small range of angle of attack, either side of the angle of attack at the design lift coefficient, resulting in the typical ‘laminar bucket’ drag variation which is seen in the graph of drag coefficient plotted against lift coefficient. The position of max­imum thickness on this aerofoil is further aft than on either the NACA 0012 or the NACA 2214. This leads to a much gentler acceleration of the air near the front of the aerofoil and the absence of the associated suction peak that pro­motes the transition to a turbulent boundary layer. The data are for a Reynolds Number of 6 x 106.

Angle of attack (degrees)

c) Variation of lift with angle of d) Variation of drag with lift attack

Fig. A.1 NACA 0012

c) Variation of lift with angle of attack

Fig. A.2 NACA 2214

a) Aerofoil section

Angle of attack (degrees)

c) Variation of lift with angle of attack

Fig. A.3 NACA 6618

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