Tab Effectiveness, b3
The data and method that follows is taken from the USAF Datcom. It provides estimates of b3 [see (2.5,1)] for two-dimensional subsonic attached flow over airfoils with a control surface and tab. Corrections for part-span tabs can be made by multiplying the result for two dimensions by the ratio of the control surface area spanned by the tab to the total control surface area (both areas being measured aft of the hinge line).
Figure B.4,2 Effect of section lift coefficient on flap section hinge moments.
The method and data that follow are taken from the USAF Datcom. The average low- speed downwash gradient at the horizontal tail is given by
— = 4.44 [KAKXKH (cos Лс/4),/2]119 (B.5,1)
Figure B.4,3 Rate of change of angle of attack due to a change in tab deflection.
where KA, Kk, and KH are wing-aspect-ratio, wing-taper-ratio, and horizontal-tail-location factors obtained from Figs. B.5,1, B.5,2, and B.5,3, respectively. Ac/4 is the sweepback angle of the wing J chord line.
At higher subsonic speeds the effect of compressibility is approximated by
/Эе =/Э£ (ClJm
)M /low speed speed
( —— ] is obtained using (B.5,1)
у OQ? /]ow Speeti
(COlow speed ar|d (ClJm are the wing lift-curve slopes at the appropriate Mach
numbers, obtained by using the straight-tapered-wing method of Sec. B. l
B.6 Effect of Bodies on Neutral Point and Cmo
c local wing chord at center line of fuselage or nacelle
c mean aerodynamic chord
w maximum width of fuselage or nacelle
5 gross wing area
Дhn shift of neutral point due to fuselage or nacelle as a fraction of c,
SB area of planform of body
SIIF area of planform of body, forward of 0.25c
c’ root chord of wing without fillets
(Cmo)B increment to Cmo due to a body at zero lift
6 reflex angle of fillet, i. e., angle between wing root chord and lower surface of fillet for upswept fillets, or the upper surface for down – swept fillets, positive as indicated in Fig. B.8,2
A fillet lift-increment ratio, i. e., CJCla, considering the fillet to be a
flap of chord lf
Figure B.6,1: Дh„
The data for estimating Дhn presented in this graph were derived from wind-tunnel tests. The forward shift in neutral point is mainly dependent on the length and width of the body forward of the wing. The values of Д/г„ given by the curves are accurate to within ±0.01c, and are about 5% higher for low-wing, and the same amount lower for high-wing configurations. The data are inapplicable if the wing is clear of the body. Separate values should be computed for fuselages and nacelles, and the results added to obtain the total neutral-point shift.
Figure B.6,2: (Cmo)B
The curves given in this figure apply to stream-line bodies of circular or near circular cross section with midwing configurations. For high – or low-wing configurations a positive or negative Д(Стц)в = 0.004 is added, respectively, to the value derived from the curves. The curves apply only for angles of attack up to about 15° for stream-line bodies where the pitching moment of the body varies linearly with angle of attack.
In the wind-tunnel tests from which the data were derived, the wings had straight trailing edges at the wing-body junction. Fillets have a large effect on Cmo, however, especially if в is large. The following equation may be used to estimate the fillet effect if 0.12 < lf/c < 0.5 and 0.03 < Sf/b < 0.075:
Cmo due to fillets = [0.046 + 0m(dCJdCL)we – 0.2(c + lf)/c](w + Sf)/b
The value obtained from the curves, the fillet effect, and the effect due to wing position are added to determine (Cmn)B.