Effect of Sweepback on Flap Performance

Reference 1 contains information of the effect of sweepback on the lift, drag, and pitching moment changes caused by flap deflection. Admittedly, the scatter in the available data is rather large, but trends are established. Because of the thickening of the boundary layer at the tips of sweptback wings and the increased loading at the tips, CLmax would be expected to decrease with increasing sweepback. The increment in CLmax appears to decrease approximately as the cube of cos A, although there is no analytical substantiation for it. ACD seems to decrease approximately with cos A. The increment in the pitching moment is complicated by the fact that, because of the sweep, changes in the spanwise loading distribution can affect the pitching moment as much as changes in the section moment coefficients. It is recommended that an analysis of the spanwise loading distribution be performed to determine CM for the sweptback wing relating CM to CL according to Eq. (6-23).

Example

For illustrative purposes consider an aircraft equipped with 30% c Fowler flaps extending 65% out along the span. The horizontal tail is located three mean chord lengths aft of the wing aerodynamic center. The rectangular wing has an aspect ratio of 5.0 and is 15% thick. What is the maximum trim lift coefficient and what flap deflection is required to achieve it?

Based on the extended chord, the 30% Fowler flap becomes a 23% c flap; hence Cl6 is 3.9 C,/radian from Fig. 6-3. This must be corrected for nonlinearities (see Fig. 6-8). Using the curve marked “slotted,” we find that a maximum value of jj<5 occurs for a d of 60° and is equal to 22.2°. Hence AC, equals 3.9 (22.2)/57.3, or 1.51; AC, max is only two thirds of this amount, or 1.01. Assuming a sufficiently high Reynolds number, the unflapped C, max for an 11.5% (30% more chord) thickness ratio would be approximately 1.6, depending on its camber. Hence C, max for the two-dimensional section with flaps would be approximately 2.6. However, AC, max must be corrected for partial-span effects (see Fig. 6-20). Compared with the full-span case,

ACLmax is only 69% as great for a bf/b of 0.65. Also, it should be reduced approximately 27% to account for the fuselage. Hence ACLmax = 1.01 (0.69) (0.73) = 0.51, so that CLmax equals only 2.11. This also must be corrected for tail download.

Referring to Eq. (6-13) for the extended chord c/lT = 1/2.3, we find that fi1 and fi2 equal 0.28 and 0.65, respectively (see Figs. 6-23 and 6-24). Hence ACM is calculated from (6-24) as

Hence AC, corrected for trim becomes

= 0.47,

so that the final predicted CLmax, based on the actual chord and remember­ing that only 65% of the wing is flapped, would be equal to 2.48.

(b)

A discussion of high-lift devices can also be found in Chapter 8 of Ref. 7. In particular, the reader is referred to that reference for a discussion of the geometry of slotted flaps. The performance of a slotted flap can depend quite critically on its geometry. Just “any old” slot will not do the job. For example, Ref. 7 shows that the optimum slot width (for a particular case) is about 1.5% of the chord. Increasing this to 3% of the chord decreases C, from 2.8 to 2.5.

•max

Some mention should also be made of slots and slats before closing this chapter. The leading-edge retractable slat and leading-edge slot shown in Fig. 6-27 are very similar in appearance and action except that the slat may extend the chord and can be positioned for maximum benefit. When used alone a well-designed leading-edge slot can increase C, max by as much as

0. 80. When used in combination with a trailing-edge flap, the increment due to the slot is only about half this value.