Effect Of Sweepback On wing Planform
Swept-Back Leading Edges
This type of planform is used on high speed aircraft and may take the form of a sweptback wing, or a delta, with a tailplane or without a tailplane. The reason for the use of these planforms is their low drag at the higher speeds. However, the high speed / low drag advantages are gained at the cost of a poorer performance at the lower end of the speed range.
Effect of Sweepback on Lift
If a straight wing is changed to a swept planform, with similar parameters of area, aspect ratio, taper, section and washout, the CL max is reduced. This is due to premature flow separation
from the upper surface at the wing tips. For a sweep angle of 45°, the approximate reduction in CL max is around 30%. Fig below shows typical CL curves for a straight wing, a simple swept-back wing, and a tailless delta wing of the same low aspect ratio.
Effect of Sweepback on CL
A swept wing presents less camber and a greater fineness ratio to the airflow. However, the reasons for the lowering of CL slope are more readily apparent from an examination of two Figs given below. From first Fig it can be seen that the velocity V of the RAF can be divided into two components, V1 parallel to the leading edge which has no effect on the lift, and V2 normal to the leading edge which does affect the lift and is equal to V cos Λ. Therefore, all other factors being equal, the CL of a swept wing is reduced in the ratio of the cosine of the sweep angle.
Flow Velocities on a Swept Wing
Effect of Change in Angle of Attack
Second Fig above (Effect of change in AOA) shows that an increase in fuselage angle of attack ∆α will only produce an increase in the angle of attack ∆α cos Λ in the plane perpendicular to the wing quarter chord line. Since we have already said that it is the airflow in the latter plane which effects CL, the full increment of lift expected from the ∆α change is reduced to that of a ∆α cos Λ change.
From the first graph above, it can be seen that the stall occurs on all three wings at angles of attack considerably greater than those of wings of medium and high aspect ratios.
On all aircraft it is desirable that the landing speed should be close to the lowest possible speed at which the aircraft can fly. To achieve this desirable minimum the wing must be at the angle of attack corresponding to the CL max.
On all wings of very low aspect ratio, and particularly on those with a swept-back planform, the angles of attack giving the highest lift coefficients cannot be used for landing. This is because, swept-back planforms have some undesirable characteristics near the stall and because the exaggerated nose-up attitude of the aircraft necessitates, among other things, excessively long and heavy undercarriages. The maximum angle at which an aircraft can touch down without recourse to such measures is about 15°, and the angle of attack at touchdown will therefore have to be something of this order. The graph above shows that the CL corresponding to this angle of attack is lower than the CL max for each wing.
to obtain a common minimum landing speed at a stated weight, an unswept wing needs a smaller area than either of the swept planforms, because Compared with the maximum usable lift coefficient available for landing aircraft with unswept wings, those of the swept and delta wings are much lower, necessitating higher landing speeds for a given wing loading. The simple swept wing needs a greater area, and so a lower wing loading, in order that the reduced CL can support the weight at the required speed. The tailless delta wing needs still more area, and so a still lower wing loading, to land at the required speed.
Effect of Sweepback on Drag
The main reason for employing sweepback as a wing planform is to improve the high speed characteristics of the wing. Unfortunately this has adverse effects on the amount of drag produced at the higher range of angles of attack. The induced drag increases approximately in proportion to 1 / Cos
Effect of Sweepback on Drag
The main reason for employing sweepback as a wing planform is to improve the high speed characteristics of the wing. Unfortunately this has adverse effects on the amount of drag produced at the higher range of angles of attack. The induced drag increases approximately in proportion to 1 / Cos Λ ( where Λ is the sweep angle) . This is because, by creating sweep, the wing CL is reduced, and therefore to maintain the same lift the angle of attack has to be increased. This increases the induced downwash and hence the induced drag.
The practical significance of this high increase in drag is the handling problems it imposes during an approach to landing. Because of the greater induced drag, the minimum drag speed is higher than for a comparable straight wing, and the approach speed is usually less than the minimum drag speed. Therefore, if a pilot makes a small adjustment to the aircraft’s attitude, for example, by raising the nose slightly, the lift will be increased slightly, but there will be a large increase in drag which will result in a rapid reduction in speed, and a large increase in power to restore equilibrium. In fact, the stage may be reached where the use of full power is insufficient to prevent the aircraft from descending rapidly.
On some aircraft this problem is overcome by employing high drag devices such as airbrakes or drag-chutes to increase the zero lift drag. This results in a flatter drag curve with the minimum drag speed closer to the approach speed (shown in graph below). A further advantage is that more power is required on the approach, which on turbojet aircraft, means better engine response
Improvement In Approach Speed Stability
Effect Of SweepBack on Stalling
When a wing is swept back, the boundary layer tends to change direction and flow towards the tips. This outward drift is caused by the boundary layer encountering an adverse pressure gradient and flowing obliquely to it over the rear of the wing.
The pressure distribution on a swept wing is shown by isobars in Fig below. The velocity of the flow has been shown by two components, one at right angles and one parallel to the isobars. Initially, when the boundary layer flows rearwards from the leading edge it moves towards a favorable pressure gradient, i.e. towards an area of lower pressure. Once past the lowest pressure however, the component at right angles to the isobars encounters an adverse pressure gradient and is reduced. The component parallel to the isobars is unaffected, thus the result is that the actual velocity is reduced (as compared to an unswept wing) and also directed outwards towards the tips.
Outflow of Boundary Layer
The direction of the flow continues to be changed until the component at right angles to the isobars is reduced to zero, whilst the parallel component, because of friction, is also slightly reduced. This results in a “pool” of slow moving air collecting at the tips. The spanwise drift sets up a tendency towards tip stalling, since it thickens the boundary layer over the outer parts of the wing and makes it more susceptible to separation, bringing with it a sudden reduction in CL max over the wing tips.
At the same time as the boundary layer is flowing towards the tips, at high angles of attack, the airflow is separating along the leading edge. Over the inboard section it re-attaches behind a short “separation bubble”, but on the outboard section it re-attaches only at the trailing edge or fails to attach at all. The separated flow at the tips combines with the normal wing tip vortices to form a large vortex (the ram’s horn vortex). The factors which combine to form this vortex are:
- Leading edge separation.
- The flow around the wing tips.
- The spanwise flow of the boundary layer.
From the next Fig it can be seen that the ram’s horn vortex has its origin on the leading edge, possibly as far inboard as the wing root. The effect of the vortex on the air above it (the external flow) is to draw the latter down and behind the wing, deflecting it towards the Fuselage
Ram’s Horn Vortex
The spanwise flow of the boundary layer increases as angle of attack is increased. This causes the vortex to become detached from the leading edge closer inboard . As a result, outboard ailerons suffer a marked decrease in response with increasing angle of attack. This, in turn, means that comparatively large aileron movements are necessary to manoeuvre the aircraft at low speeds and the aircraft response may be correspondingly sluggish. This effect may be countered by limiting the inboard encroachment of the vortex as described below, or by moving the ailerons inboard. Another possible solution is the use of an all-moving wing tip.
Alleviating (Avoiding) the Tip Stall
Most of the methods used to alleviate the tip stall aim either at maintaining a thin and therefore strong boundary layer, or re-energizing the weakened boundary layer.
- Vortex Generators. These re-energize the boundary layer by making it more turbulent. The increased turbulence results in high-energy air in layers immediately above the retarded layer being mixed in and so re-energizing the layer as a whole. Vortex generators are most commonly fitted ahead of control surfaces to increase their effect by speeding up and strengthening the boundary layer. Vortex generators also markedly reduce shock-induced boundary layer separation, and reduce the effects of the upper surface shockwave (explained in chapters of lift augmentation and high speed aerodynamics).
- Boundary Layer Blowing. High velocity air is injected into the boundary layer to increase its energy.
- Boundary Layer Suction. Suitably placed suction points draw off the weakened layer and a new high-energy layer is then drawn down to take its place.
- Leading Edge Slots. These have the effect of re-energizing the boundary layer.
- Boundary Layer Fences. Used originally to restrict the boundary layer out-flow, fences also check the spanwise growth of the separation bubble along the leading edge.
- Leading Edge Extension. Also known as a “sawtooth” leading edge, the extended leading edge is a common method used to avoid the worst effects of tip stalling. The effect of the extension is to cut down the growth of the main vortex. Further effects of the leading edge extension are:
- The t/c ratio of the tip area is reduced, with consequent benefits to the critical Mach number.
- The Centre of Pressure (CP) of the extended portion of the wing lies ahead of what would be the CP position if no extension were fitted. The mean CP position for the whole wing is therefore further forward and, when the tip eventually stalls, the forward shift in CP is less marked, thus reducing the magnitude of the nose-up movement.
- Leading Edge Notch. The notched leading edge has the same effect as the extended leading edge insofar as it causes a similar vortex formation thereby reducing the magnitude of the vortex over the tip area and with it the tip stall. Pitch-up tendencies are therefore reduced. The leading edge notch can be used in conjunction with extended leading edge, the effect being to intensify the inboard vortex behind the devices to create a stronger restraining effect on boundary layer out-flow. The choice whether to use either or both of these devices lies with the designer and depends on the flight characteristics of the aircraft.