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Aviation History
1934
1934 - 1477.PDF
DECEMBER 27, 1934 91 THE AIRCRAFT ENGINEER SUPPLEMENT TO FLIGHT 1382c the wing pitching moment has been plotted against in- cidence for various flap settings at constant kL (0.5). dkm/da from these points has a value of 0.009 per degree. An average tail plane designed to give adequate longitudinal stability may have a value of 0.007 f°r dkm/da, so that there is a net nose down pitching moment of 0.002 per degree change of incidence at constant kL. For full flap angles the change in incidence is about 12 deg., giving a net pitching moment coefficient of — 0.024 only. Some tests made by the Zap Corporation on models of the XOJ-i biplane and Parnall Parasol monoplane even show resulting nose up moments at a given kL and practically no change in moment at maximum, lift. In these tests it is hard to see where the additional nose up moment comes from. Failure to carry the lift right across the centre of the wing, either due to interruption of the flap or to bad body-wing interference will reduce the down-wash at the tail for a given lift coefficient and almost certainly make it impossible to attain stalling incidence if that incidence is only just attainable with no flap. In order to ensure that stalling incidence is just attainable with and without flap it may be necessary to gear together the flap and tail plane adjustment or alternatively to arrange that the gearing between elevator and stick is increased as the flap is pulled down. The former system was adopted in the Fieseler 97 in the Rundflug competition. This machine has a Fowler flap involving a very big C.P. movement and its tail plane is set down 16 deg. by full flap movement. In spite of this its minimum speed figure (engine on) only gives a kL of 1.57, so it is quite likely that there was insufficient longitudinal control to trim the aeroplane to the incidence for maximum lift, though engine power and lateral control may also have been inadequate. The sharp drop in lift at the stall that has been noticed in model and full-scale wind tunnel tests of split flaps has given rise to some anxiety on the score of a possible large loss of height in recovery from a stall, combined with acute lateral instability. Mr. Relf has suggested that on this account the longitudinal control should be limited to prevent the attainment of stalling incidence, but full-scale evidence on this point is needed. Operating Forces for Split Flaps Data on the hinge moment and centre of pressure of a split flap are a little scarce. Curves of hinge moment for flaps of varying chord are shown in Fig. 6. For such flaps the hinge moment is very nearly proportional to the angle of deflection. The rapid increase of hinge moment with chord shows that flaps should be kept as long and narrow as possible for a given area, a requirement in conflict with the avoidance of aileron troubles. A cal- culation based on these hinge moment figures shows that the moment required to hold down the flaps of a machine like the Douglas D.C.2 is about 700 lb. ft. at 75 m.p.h. This demands the use of auxiliary operating gear which may be electric or hydraulic but is bound to be fairly slow acting. A criticism of the flap installation on the Northrop " Sky Chief" was that 45 turns of a crank handle were required to get the flaps down ; this manual system has been replaced on later Northrop machines (with smaller flaps) by a hydraulic gear, and it appears that manual operation of a plain split flap of useful size is impracticable on any but light aeroplanes. It is in connection with operating force that the Zap arrangement is of principal interest ; by hingeing the flap to a link about one-third of its chord back and allowing the nose of the flap to slide back, the flap may be nearly in balance at high angles. The position of the flap C.P. is open to some doubt ; N.A.C.A. tests on plain split flaps show that, at the operating incidences, the C.P. moves back from 30 per cent, of the flap chord at 15 deg. deflection to about 40 per cent, at full deflection, whereas some tests by the Zap Corporation show the C.P. to move forward from 0.5 to 0.3 of the flap chord as the angle increases from 20 deg. to 40 deg. On the latter basis the flap is self- opening at high angles. For a simple hinged trailing edge flap there is a theoretical relation between hinge moment, flap angle and lift coefficient due to Ff. Glauert; this relation has been checked very closely by pressure plotting on a R.A.F. 30 flapped wing and by direct measurement on a R.A.F. 31 model. The results show a straight line law for kH against flap angle at a given kL and the value of kH at 45 deg. flap angle is identical with that for the split flap given in Fig. 6. The Slotted Flap The principle of operation of this flap has already been referred to, and it would be expected that the drag would be comparatively low. Wind tunnel tests on a Handley Page slotted flap of 20 per cent, chord ratio show an increase in kL max. of 0.445 at 40 deg. deflection and an increase of only 0.018 in kD at kL = 0.5. With the slot blocked up AhL max. = 0.32, Aku = 0.017 at *L = °-5- Some other tests on slotted flaps show them to be the type having the least increase in drag for a given increase in maximum lift coefficient. With regard to hinge moments, no figures are available, but it appears possible to locate the hinge in such a position that the flap is well balanced and has a suitable slot opening. There is no advantage in having the flap angle greater than 40 deg., so that the range of movement is less than that required from a split flap to give the same increase in maximum lift coefficient. Air Brakes All the devices discussed above have been regarded principally from their effect on maximum lift, and the drag has been left to look after itself. If KL max. is assumed to be increased by a flap from 0.6 to 1.0 then KL on the glide will be increased from 0.45 (at 15 per cent, above the old stalling speed) to 0.7 (at 20 per cent, above the new stalling speed). For an effective aspect ratio of 6 this means that the induced drag coefficient is increased from 0.021 to 0.051. The parasitic drag may be increased any amount from 0.016 for a slotted flap to 0.089 f°r a 9° deg. split flap, so that the total increase in gliding angle, given by tan-1 Akn/KL ranges from 2.8 deg. to 11.5 deg. This can be said to cover any normal requirement, ranging from the case of low drag increase where improvement in take-off is required to quite a high drag increase where short landings are required. There are other means of increasing drag alone without altering the wing lift characteristics, but they are not very effective. A retractable undercarriage which is badly faired when down, rotatable strut fairings, flaps projecting from the fuselage, and such devices are only capable of producing quite small drag increases. A variable pitch airscrew windmilling at very low pitch has been calculated to give quite a good braking effect, but the tail controls might be impaired by a " negative " slipstream. A device which acts as a pure air brake and has been tried successfully in flight is the D.V.L. gliding angle control. This consists, essentially, of two small surfaces, one on the upper surface of each wing, which are parallel to the plane of symmetry when out of action and can be rotated about a vertical axis. When turned normal to the direction of flight they break up the span distribution of loading and so increase the induced drag of the wing appreciably though their own parasitic drag is small. Flight tests on a German low-wing monoplane showed an increase of 5 deg. in minimum gliding angle with this device without appreciable increase in stalling speed. Weight of Flap Installation There is very little evidence in this respect. Experi- mental flap installations on the Parnall Parasol monoplane give figures of 250 lb. for the complete outfit, including controls, of a 28 per cent, chord Zap flap of full span, except for a small centre section cut away, and 651b. (estimated) for a 10 per cent, chord full span curved
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