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Aviation History
1964
1964 - 0022.PDF
FLIGHT International, 2 January 1964 TURBM.ENT WAKf VORTEXCORE 17 Fig II This sketch by the Lockheed California Co suggests the unusual pitch-up bthaviour of the F-104; at extreme angles of attack the problem is aggravated by a download on the tailplane. Nevertheless, say Lockheed, the wing lift does not fall away Fig 12 Leading-edge separation from a crank wing: the vortex has moved inboard to the inner crank, leaving a region of separated flow (streamlines shown dotted) at the tip tailplanes (tailerons). Wave-drag penalty would appear to prohibit the use of vortex generators for suppressing trailing-edge separation on supersonic aircraft; moreover they have not proved very effective for preventing leading-edge separation, even when mounted right at the leading edge. Once the wing vortex has formed it is necessary to delay its progress along the leading edge with the aid of a kink or fence and/or to counteract its effects with a second, artificially generated vortex. The latter is achieved by means of the fence if fitted (as on most current Russian supersonic aircraft), the saw-tooth junction of an extended leading edge, or a notch cut in the leading edge. The extra lift induced on the outboard panels by the secondary vortex balances the lift lost due to the onset of separation, the main vortex having moved inboard towards the root. It may be necessary to prevent the main vortex from moving right into the root, otherwise it may cause pitch-up by overriding the effects of the secondary vortex with high lift in the root region. If this is the case, downward camber applied to the leading edge in the root region (Lightning) will hold the vortex out, as will a crank in the leading edge near the root. Once the main vortex has moved inboard as far as it can, a stable flow pattern is reached with the lift from the two vortices giving a fore- and-aft balance. However, pitch-up is not entirely eliminated, because the secondary vortex will lift off the wing surface at a much lower incidence than the main vortex, thus ceasing to give the required induced lift. The fact that the main vortex, now at the root, remains stable and attached to the surface to quite high incidences is of great importance, in that it forms the basis of slender-wing aerodynamics. Something must now be said on the subject of shock-induced separation on swept-wings. This is essentially a transonic-flow phenomenon, with the occurrence and severity strongly dependent on wing planform, section geometry (particularly thickness and trailing-edge angle), Mach number and incidence. As such it is of relatively little importance on most supersonic aircraft having very thin wings and operating only transiently in this flow regime (and then only at low incidences). But it must be suitably accounted for in the design of high-subsonic and transonic aircraft—which look like being around for a long time, at least in the form of transports. Airflow becomes transonic when the local velocity somewhere over the aircraft surface reaches Mach 1.0. However, since the wings form the most important aerodynamic component, penetra- tion into the transonic regime can be said to begin effectively when a sonic velocity-component is reached normal to the isobars at any point on the wing surface. Beyond Merit, so defined, the appear- ance of shock-waves causes a rapid increase in drag, which appears as pressure drag and is due to the energy absorbed from the air- stream by the shock-waves. The drag increase with Mach number is so large* that for economy/range reasons, a subsonic aircraft will not usually operate for any length of time at a speed substanti- ally beyond its drag-rise Mach number, even if it has the power available to do so. However, a high-speed dash or an abrupt increasem CL while manoeuvring at speeds close to Mcrit can take the aircraft well into the transonic regime, and it is therefore necessary to ensure that any adverse stability effects are delayed and minimizedas far as possible. Delay of pitch-up thus becomes an extension to Typically, the drag of a Hunter rises by a factor of five through the"'""ionic speed range. the overall problem of raising the CL-Mcrjt boundary of the wing, the present tendency being to design for a suitable safety margin over and above the drag-rise Mach number. Since the criterion is Mach number normal to the isobars, rather than free-stream Mach number, an improvement in Mcrit can be obtained by reducing the t/c or, more effectively, by increasing the sweep. Both measures tend to be undesirable, the first from a structure/weight standpoint, the second from lift and stability considerations, particularly at low speed. In practice then, for a reasonable all-round performance without the use of BLC or variable geometry, a compromise wing of moderate sweep will be chosen; all present knowledge of transonic flow about swept wings will then be applied to the detail design of the wing such that the drag rise and flow separation are delayed as far as possible. Usually stability itself is not much affected by the presence of shock-waves until the latter become strong enough to induce flow separation by interaction with the boundary layer. This can occur because the boundary layer will suffer a substantial loss in energy (and rise in pressure) when it passes through a shock; also the flow may be diverted in a more spanwise direction as a result of the shock reducing the chordwise velocity component while leaving the spanwise component unchanged. When separation does take place behind a shock it tends to develop a vortex form; thus the flow pattern at transonic speeds is further complicated by the possibility of interaction of shock-waves both with this type of vortex and, at high incidence, with the normal leading-edge vortex. As has been shown earlier, the boundary layer on a swept wing tends to be thickest in the tip region; unfortunately, when the Mach number and/or incidence is raised, shocks of significant strength first appear in this region also. Accordingly, when separation does occur it will probably take place at the tip, giving a pronounced pitch-up tendency. Because the flow and shock patterns change rapidly with even slight changes in incidence or Mach number, buffeting will probably occur when flow separation takes place; in addition there is a possibility of rapid control-surface oscillations (aileron buzz) occurring. Since the condition of the boundary layer is a dominant factor causing shock-induced separation, BLC can be very effective in suppressing, or at least delaying, it. This has been proved many times in wind-tunnel tests; but, as yet, the incorporation of BLC has found no practical application because it would be a heavy price to pay for the suppression of a phenomenon which is encoun- tered only transiently, the wave-drag alone being a sufficient deterrent to flight in the transonic regime. The penalty associated with vortex generators is of course much less, and these devices have been used with varying success on many aircraft. In transonic flow, the effectiveness of a row of vortex generators is very much dependent on its position with respect to the shock- waves which form. The vortex generators need to be placed a short distance ahead of the shock in order to re-energize the boundary layer before it passes through the shock; hence the optimum positioning and spacing have to be found more or less by trial and error, because of the considerable variation of shock pattern with incidence and Mach number. A beneficial effect of a different kind may result if the shock attaches itself to a row of vortex generators; when this happens, as it does on the Javelin, the normal forward progress of the shock (with incidence) is inhibited and the spread of flow-separation is curtailed.
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