FlightGlobal.com
Home
Premium
Archive
Video
Images
Forum
Atlas
Blogs
Jobs
Shop
RSS
Email Newsletters
You are in:
Home
Aviation History
1951
1951 - 1587.PDF
FLIGHT, 17 August 1951 THE DESIGN-STUDY . . . I have evolved a brief method of performing a preliminary design- study, whereby certain weight, performance and stability criteria are tied down in reasonable approximations. The method makes it possible for a designer to carry out a design-study in a matter of hours, and has been used successfully by the author in lieu of the well-recognized formulae. The orthodox, unabridged methods have frequently been largely negatived by some change in the size and weight of equipment, or by some unaccounted-for drag. All too frequently when such an omission is discovered, everyone concerned loses patience, and the calculators themselves lose faith in their calculations and start approximating. By contrast, the use of approximate formulae from the start makes it possible to calculate more variants to the same specification, and to proceed with project sketching at the same time. The following method of preliminary design-study permits calculation of the conventional aircraft dimensions as a function of the desired structure weight, providing that the following specifi- cation demands are stated: (1) Performance requirements: 1.1 Required take-off distance to 50ft screen. 1.2 Required minimum rate of climb. 1.3 Required minimum cruising speed. 1.4 Required landing distance from 50ft screen. Sometimes, a few additional requirements are given : 1.5 Maximum range (for long-range aircraft irrespective of speed). 1.6 Stalling speed with appropriate load, flaps up or down. In most cases, the first four of the above requirements are the most severe and, consequently, the method presented here has been based on them. If necessary, however, the last two require- ments can be evaluated separately, and "ofFered-up" to the dimensions previously derived. (2) Determination of values independent of wing size and area: The following items have to be determined first, and indepen- dently of the final aircraft dimensions : 2.1 Aircraft all-up weight minus structure weight. This comprises disposable load (crew, passengers, luggage, fuel and oil, furnishing), plus engine installation complete with fuel and oil installation (tanks and pipes), power services, including electrics, hydraulics and pneumatics. A good approximation to the weights of all the items listed above can conveniently be taken from the weight breakdowns available for similar existing types of aircraft. Such a breakdown is shown in Table I (Ref. 2). In order to choose the correct engine power relative to engine weight, and so ensure prevention of the aircraft being under- powered, it is wise to provide for a power loading not greater than 8.5 lb/take-off b.h.p. (Ref. 3.) 2.2 Fuselage and nacelle sizes and cross-sectional areas. These can be determined from the following formulae : Fuselage: A$ = 0.15 X (payload)% (Ref. 3) Nacelle: 4 = Kx T-O b.h.p. where K = 0.0042/T-O b.h.p. for turboprops. K = 0.0050/T-O b.h.p. for piston engines. 2.3 Empennage size approximation. This can be obtained from the following formulae (Ref. 4) : LL= 4.0 for multi-engined aircraft \ for tailplane and = 3.0 for single-engined aircraft / elevator = 3.0 for multi-engined aircraft \ f fi d dd = 2.5 for single-engined aircraft/ m"luUiUUU>:' where 5T is empennage nett area excluding the fuselage. LT is empennage moment arm from the aircraft's e.g. LFF is front fuselage length ahead of the wing. Spp is fuselage area in plan form ahead of the wing. 5FS is fuselage area in side view ahead of the wing. Note: Fin and rudder area calculated to this formula do not include the fuselage side area immediately under the tailplane on single-fin aircraft. The latest tests indicate a very poor fin efficiency under the tailplane. (Ref. 4.) 2.4 The parasitic drag of the aircraft can now be expressed in terms of the wing area, provided that the items under paras. 2.2 and 2.3 have been determined. CDo=Cowing+ Cpfus X A¥ + Cpemp X (STp + SPR) 4- Cpnac X As^ where CD0 wing, Co fuselage, Co nacelles, Cp empennage, and A CB0 flaps have values as suggested in Refs. 5 and 6. For very rough calculations, the following formulae have been successfully used by the author : 2.4.1 Wing drag: CDo wing=C+o.ooi (log V m.p.h. + log E—4.075)2 for mean t/c=o.i5. where c=wing mean chord. C=o.oo38 for laminar sections with transition at 0.5c. =0.0068 for conventional sections with transition at the leading edge and a substantially linear variation between them for the appropriate transition. 2.4.2 Fuselage drag: CD fuselage=0.078 on transport types, good multi-panel wind- screen. =0.105 on transport types, blunt Vee windscreen. = 0.07/0.08 on fighter types, excluding windscreen. 2.4.3 Windscreen drag: CD windscreen=0.03 for best streamlined canopies, no flat panels, no sharp corners. =0.10 for average type, no flat panels, but with retaining strips. =0.34/0.40 for blunt panels with sharp corners, or conical-type screens. All the coefficients in paras. 2.4.2 and 2.4.3 above refer to fuselage or windscreen cross-sectional areas. (Ref. 7.) 2.4.4 Empennage drag: ' CD=O.OO8 for good, thin (0.09/0.12 t/c) sections, sealed gaps. =0.01 for average surfaces, including control gaps. 2.4.5 Nacelles (cross-sectional area): CD=O.IO for good nacelles with minimum leaks. = 0.13 for poor nacelles with leaks and wheel doors badly sealed. 2.4.6 Undercarriage drag (u/c frontal area in "down" position) : CD=O.I8 for good, fully spatted leg with faired wheels. = 0.7 for spatted leg with unshielded wheel. = 1.0 for totally exposed single-strut leg, unshielded wheel. = 1.3 for exposed girder-type leg, unshielded wheel. Note: By arranging that the fairing doors close off the larger part of the retraction well whilst the undercarriage is still extended (during take-off and climb), the undercarriage drag on two well- known civil aircraft has been reduced to less than 60 per cent of its original value, and has enabled these aircraft to fly comfortably on one engine with the undercarriage down. For twin-engined aircraft, this configuration is, of course, one of the most stringent performance requirements. 2.4.7 Dead-engine airscrew drag: The drag of a windmilling airscrew is one of the most difficult quantities to estimate, for it varies with throttle and ignition setting as well as airscrew blade pitch setting. Average values can be derived from the following formulae : Drag, pounds, at iooft/sec Dioo=ND2IK where K=6j for airscrew stopped, blades feathered. = 18 for airscrew windmilling, coarse pitch. = 5 for airscrew windmilling, fine pitch. D = airscrew diameter, feet. N— number of blades. Note: It is apposite to point out that the drag of the dead airscrew causes some additional drag owing to rudder and aileron deflection required, and this consequently increases the induced and profile drags. This asymmetric power drag is difficult to estimate quickly, as it depends to a great extent on the rudder and aileron size, empennage and wing configuration and aspect ratio. Curiously enough, however, the additional drag represents almost a constant value expressed in terms of horsepower throughout the speed range, and can be taken to equal the drag of the dead engine (also expressed in horsepower) calculated in accordance with the above formula, at the minimum control speed. In other words, the horsepower required to fly with one engine dead equals the standard drag horsepower plus the double drag horsepower of the windmilling airscrew. (3) Determination of aircraft structure weight and dimen- sions : 3.1 Aircraft structure weight: This quantity depends primarily on the spanwise load per inch chord, and secondly, on stiffness requirements. These, in turn, depend on (a) wing area, loading, span, root t/c ratio and ultimate factor; (b) fuselage length and surface area; (c) empennage area; (d) undercarriage factors and size; and (c) nacelle size. It should here again be remembered that every one per cent reduction in structure weight for a conventional aircraft means 4 to 5 per cent increase in payload for the same performance. 3.1.1 Wing structure weight: There are several wing structure formulae; that used by the Lockheed Aircraft Corporation for their large civil aircraft is as follows : per cent=7.6 + (4g^) Alternatively, the author has used the following formula (derived from Ref. 8 and simplified) for light aircraft: (WW!W) per cent=Arxfrx(o.ooo27_x- + o.oo8) + — + 5.0 where N is the ultimate load factor and A, b, t, w are the wing aspect ratio, span, root t/c ratio, and loading respectively.
Sign up to
Flight Digital Magazine
Flight Print Magazine
Airline Business Magazine
E-newsletters
RSS
Events
