FlightGlobal.com
Home
Premium
Archive
Video
Images
Forum
Atlas
Blogs
Jobs
Shop
RSS
Email Newsletters
You are in:
Home
Aviation History
1964
1964 - 0021.PDF
16 THE HIGH-SPEED SHAPE . . . Fig 9 Subjected to aerodynamic development as extensive as that of its partner the Vixen, the Scimitar has a wing with considerable sweep and low aspect ratio, with dog tooth and low fence miniature aerofoils they generate a series of small vortices similar to the trailing vortices from a wing, and these decrease the likelihood of boundary-layer separation by injecting higher-energy air from the free stream. A row of vortex generators placed just forward of the ailerons will improve control effectiveness; but maximum CL may not be increased because, at high incidence, separation will probably take place forward of the generators. When this happens they will be enveloped in separated flow and will cease to function; the remedy is a second row of generators further forward (Fig 10). Under flight conditions in which the vortex generators do their required job, the improved flow over the wing surface will give reduced drag. Under other conditions, where they become super- fluous, a drag penalty must be expected; in practice this will gener- ally be negligible for a row of generators mounted well aft on the chord, but a row further forward will raise the aircraft profile drag by perhaps five per cent. Weight and complexity considerations have so far ruled out both retractable and featherable types of vortex generator; the same can be said of arrangements in which the vortices are generated by air jets issuing obliquely from a row of small holes in the wing surface. As aircraft cruise performance penetrates further into the transonic range, wing sweep must increase and thickness/chord ratio decrease in order to raise the drag-rise Mach number. As the thickness/chord ratio decreases, a stage is reached when the trailing- edge separation typical of thick wings at subsonic M gives way to laminar separation from the leading edge. For supersonic speeds wing sections of 5 % t/c or less are inevitable, because wing wave- drag is proportional to the square of the thickness/chord ratio; so leading-edge separation will occur on these wings also. Such thin profiles, often uncambered (i.e., symmetrical) and with fairly sharp leading edges, are basically unsuited to subsonic flow except at very low incidences; the airflow cannot negotiate the sharp leading edge and it therefore separates (Fig 11). In the case of a wing with an outboard crank this will probably occur at the kink, since a high suction peak is induced at this point. For a straight-leading-edge swept wing the highest suction peak and surface airflow velocity appear near the tip, so separation normally occurs there first. Initially, there is a tendency for the separated flow to re-attach to the wing surface further back, at perhaps 15 per cent chord, the laminar flow having become turbulent. When this happens there is a bubble of separated flow over the upper surface near the leading edge; the external flow is not greatly affected, and lift and pitching-moment effects are small. At slightly higher incidences, however, the bubble bursts; a strong vortex is then formed by the rolling-up of the vortex sheet between the separated-flow region and the free stream directly above it. This vortex lies back across the upper surface of the wing, and has an effect similar to that generated by a fence or notch. Because of the increased lift produced in the region of such a vortex, one close to the tip actually increases the lift in the tip region. Thus, the onset of separation on a thin wing causes an increase in the overall lift-curve slope, and may give an initial pitch-down moment. As the incidence is increased the vortex moves inboard along the leading edge, at a rate which depends on the latter's sharpness and sweep. A substantial area of low-lift separated flow is left at the tip; the lift slope decreases again, and the slight pitch-down moment changes rapidly to a strong pitch-up moment (Fig 12). This form of separation can be countered by two methods which are essentially complementary. One is to attempt to suppress the separation; the other is to ensure that it is controlled in such a way that overall stability is not lost. Suppression involves assisting the airflow to negotiate the sharp leading edge, by (for example) reducing the effective incidence at the tip with 2° to 3° of downward twist. A second method is to make the leading edge less sharp, and therefore less sensitive to incidence; tunnel tests show that, FLIGHT International, 2 January 1964 provided the blunting is slight and confined to the tip region, there may be little or no drag penalty. A more effective method is to increase the camber of the forward part of the wing. The cambered, or "drooped," leading edge is seen on many contemporary aircraft, though its primary purpose may be different from that suggested here. Thus the pronounced conical camber fitted to the wings of General Dynamics deltas is designed to improve the leading-edge suction and hence the lift: drag ratio for high-subsonic cruise; pitch-up improvements are a useful by- product. A limit to the amount of camber is imposed by a rise in supersonic drag, because at very low incidence the flow tends to separate from the underside of the drooped region. Variable cam- ber, in the shape of a nose flap which can be lowered when required is an obvious remedy; but this increases weight and complexity. The same can be said of the leading-edge slat, which gives similar benefits. Unlike the trailing-edge flap, the nose flap and slat have little effect on the lift attainable at a given incidence; rather they increase the range of incidence, and therefore lift, which can be used. In this connection it is necessary to consider the effects of boundary- layer control (for high lift rather than for laminar flow), since this is fitted to a number of current military types. When used at the leading edge, or at the knee of a nose flap, BLC by either blowing or suction delays leading-edge separation by helping the flow to adhere to the surface. Similarly BLC in front of the ailerons prevents trailing-edge separation and improves control effectiveness; in neither case is lift, as such, greatly affected. On the other hand BLC blowing applied to deflected trailing-edge flaps greatly increases the lift obtainable at low speed. BLC is at present used not so much to boost the overall lift obtainable but rather to reduce the incidence required for a given lift.* This not only alleviates the wing pitch-up problem but also reduces the pitch-up contribution due to the fuselage, which in the case of a long-nosed aircraft is very destabilizing. The same effect is obtained with the variable-incidence wing on the F-8; when the pilot operates the jacks on the approach, the fuselage rotates nose- down about the wing. The effects on aircraft pitching characteristics of the high local lift induced by blown flaps are similar to those of unblown flaps but more pronounced. If the flapped region is close to the aircraft e.g., the pitching moment can be trimmed out by a modest-sized tailplane, particularly one of the all-moving variety; thus inboard flaps are acceptable on most swept planforms. On the other hand, the outboard parts of a full-span flap give lift well behind the e.g. and impart pitching moments which may be intolerably high for a normal tailplane. Similarly, for inboard flaps, the lower the trailing- edge sweep and aspect ratio, the farther behind the e.g. will be the region of high flap lift and again the greater will be the tailplane power required. Because of these effects, the tailplane will have to be either very large (TSR.2 and A-5 Vigilante) or fitted with BLC (Buccaneer). In general, if no tailplane is fitted, flaps (blown or unblown) cannot be used. If leading-edge separation is suppressed by section changes, or leading-edge devices as described above, it may happen that separation over the rear part of the chord occurs first. If this is the case a fence will help considerably, while blunt trailing-edge ailerons can be used for greater control effectiveness (Lightning, Javelin). However, the problem of aeroelastic distortion on very thin wings is causing greater use of inboard ailerons, spoiler controls and rolling * Reducing the incidence in this way makes landing easier, particularly on a carrier; pilot view is improved and the "hook to eye" distance is reduced, as is the distance the nosewheel has to move through at touchdown (the nosewheel slam problem). Fig 10 Possibly the ultimate example of "palliatives" is the wing of the Hawker Siddeley Javelin. This, FAW Mk 9 has a leading-edge kink (extreme right), droop, and three rows of vortex generators
Sign up to
Flight Digital Magazine
Flight Print Magazine
Airline Business Magazine
E-newsletters
RSS
Events
