Plots are given of both the experimental and theoretical characteristic coefficients versus flap angle, in order to provide a comparison with the theory. Plots of normal force coefficient versus angle of attack for different flap displacements are given to show the effect of a displaced flap. Isometric diagrams of pressure distribution are given to show the effect of change in incidence, flap displacement, and scale upon the distribution. In order to study the effect of scale measurements were made with air densities of approximately 1 and 20 atmospheres. 30 (symmetrical) airfoil with trailing edge flaps. Only above 20.000 ft will the envelope expend enough to make flight at Mach 1.8 possible.įor a typical wing loading of $\frac$ = 4.Pressure Distribution Over a Symmetrical Airfoil Section with Trailing Edge Flap Measurements were made to determine the distribution of pressure over one section of an R. Normally, at low level the maximum dynamic pressure limit and engine thrust will only allow very low supersonic speeds around Mach 1.1 to 1.3. The angle of attack in straight flight depends on altitude, wing loading and aspect ratio and grows with decreasing air density.
In other words, if the plane flies at high angle of attack, it decelerates to subsonic speed quickly. The angle of attack in supersonic flight depends on air density and load factor, but generally is only a few degrees because drag increases with the square of angle of attack. How the idealised flow around a supersonic airfoil looks is explained in this answer. This produces nose suction, and flow around the wingtips disturbes the pure supersonic flow characteristics even when the leading edge is supersonic. Those airfoils are chosen because flow around them at supersonic flight speed is still heavily influenced by subsonic flow characteristics as long as wing sweep allows for a subsonic leading edge. As you can see, the F-16 uses camber throughout, indicated by the 0.2 design lift coefficient of the airfoil, while the F-15 uses an uncambered root airfoil. This information is from The Incomplete Guide to Airfoil Usage by Dave Lednicer. General Dynamics F-16 NACA 64A204 NACA 64A204 McDonnell Douglas F-15 NACA 64A006.6 NACA 64A203 Normally, supersonic fighter wings for which airfoil information is published use a very thin NACA 6-digit section with very little camber, such as Aircraft root airfoil tip airfoil Wondering about that made me think that the reason might be that there is high-pressure in front of the wing´s top-point like there is at the bottom in front of the low-point, and low pressure behind the low-point of the wing like there is behind the top-point.Īssuming a symmetrical biconvex wing profile, that would put the centre of lift the 50% of chord.Ī symmetrical biconvex wing going straight forward would produce no lift, so it has to be put at an angle to compensate for the weight of the airplane and the Mach-tuck, but still the lift would manifest itself at 50% of chord "seen my way" - and that gets more and more the faster the plane goes as the AoA gets closer to 0.Ī couple of extra questions: at what AoA is a fighter (F-16/18/22/35/Rafaele/Typhoon/Gripen) flying with in low (M 1.3) and high (M 1.8) supersonic flight?ĭo fighters have symmetrical biconvex wings or do their wings have a camber? In subsonic flight the lift manifests itself at 25% of chord, but shifts to 50% of chord when the airplane goes supersonic.
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