FlightGlobal.com
Home
Premium
Archive
Video
Images
Forum
Atlas
Blogs
Jobs
Shop
RSS
Email Newsletters
You are in:
Home
Aviation History
1919
1919 - 0276.PDF
SOME POINTS IN AEROPLANE DESIGN* BY F. S. BARNWELI I STARTED with a. very ambitious programme, but soon found that it was quite impossible in the scope of a single paper to deal with more than a few points in the vast subject of aeroplane design. What I have attempted, therefore, in a much chastened spirit, is to work out and analyse a few points which seem worthy of particular attention, and thence to suggest what would appear to be the best practice. It is probable that what I have done has been done already by most people who have been engaged on designing work for the last few years, and probably done more fully and more accurately. But it has entailed quite a fair amount of work, and the results have been of some interest and value to myself, so I hope that they may prove the same to others. I shall try to make quite clear the data from which I start, the assumptions I make and the limitations of the conclusions arrived at; so I trust that nothing in this paper should be misleading even if it be more or less incorrect. To those of you who have not done very much the same work for your selves I hope that it may be of some value, whilst to those of you who have, I hope that it may be of interest as corro borating or contradicting your own conclusions. It is perhaps unnecessary to state that I have not consciously plagiarised, so that anything; here given may at any rate be regarded as an independent check. The first subject touched upon is the aerofoil (Fig. i, p. 274.) It is a debatable question whether any one form of aerofoil section ca.n be taken as the best for all types of aeroplane. Generally speaking, the thicker the section the lighter will be the main spars for the same strength, but the worse the optimum lift over drag. Moreover, in a thick section the maximum lift value is not much higher than in a thinner one, and the critical angle occurs earlier. Taking it all round, I believe that little is to be gained by using different forms of section, and that it will nearly always prove advantageous to use a so-called " high-speed " form of section and to vary the loading and points of support so as to suit best any particular case. I show an aerofoil section here which is quite a good stand ard. I do not wish to infer that it is the best form, but it is not a bad one, and will serve as a starting point. It is, actually, a mean between two sections for which I have experimental data. I have assumed, therefore, that its aerodynamic properties are the means ; this is probably rather inaccurate, but I escape the accusation of giving away actual data. I shall allude to it henceforth as " standard section." The next point to consider is the plan form, which includes the question of aspect ratio and of shape of ends. In any aerofoil structure other than a cantilever, I do not think there is any noticeable advantage in tapering from root to tip. There i3 a distinct aerodynamic gain in rounding off and fining down the tips, but there is little more to be gained by rounding them off more than as shown in the figure. The form of end shown here is what I think, at present, to be the right type. A tip of this form gives, I think, more efficient aileron control, by which I mean better rolling moment for the same force applied by the pilot, than does the " raking " tip which we have nearly all used for some years past ; it is probably as aerodynamically efficient a form as any other of the same length of taper, and it gives front and rear spars of approxi mately the same length, which is an advantage from the weight for strength point of view. I shall call this the " standard " wing tip. For purposes of getting out areas, the end of the rectangle of same chord and of same area as the aerofoil is shown in dotted lines ; area grrp3 T 1 chord gives ns the necessary extreme span for an aerofoil with these standard tips. The span of this rect angle is the " mean " span for the aerofoil, and the aspect ratio of the rectangle the " mean aspect ratio " of the aerofoil. Now to consider the aerodynamic properties of this aerofoil of " standard " section and with " standard " tips. The family of curves shown here are all for values of lift/drag on a base of " absolute " lift coefficient. Curve A is for an aerofoil of 5/1 " mean aspect ratio " and of 3-in. chord, with a relative wind speed of 40 ft. per second. It is actually run from the mean values of wind tunnel figures for two models of rectangular plan form of 6/1 aspect ratio, and probably represents facts fairly accurately. Curve 2 however is a result of rather wild exterpolation and guessing ; it purposes to be that for a full-size aerofoil (say, over 3 ft. chord), of same form as the model, at full-size * Paper read before the Royal Aeronautical Society, February »6, 1919. 6 ' , CAPTAIN, R.A.F. speeds (say, over 40 m.p.h.). I do not guarantee that it is even approximately accurate, but I think it is so. Curve (3) is for monoplane form of mean aspect ratio 7. Curves (4), (5) and (6). are for biplane forms of mean aspect ratios 3, 7 and 9 respectively, whilst Curves (7), (8), (9) and (10) are for triplane forms of mean aspect ratios, 5, 7, 9 and n respectively. All of Curves (2) to (10), inclusive, are for " full-size " aerofoils of " standard " section, with " standard " tips, at " full-size " speeds. For the biplane and triplane forms the " mean " aspect ratio is that of each aerofoil, of course. In all cases the gap — .8 chord, and there is no stagger. To obtain these curves I have used hgures from model experiments, assuming that relative values are the same for full-size monoplane, biplane and triplane as they are for model. Similarly for increasing aspect ratio, I have taken that the effect is in the same proportion for full-size monoplane, biplane and triplane as it is for model monoplane. So it is quite probable that these curves are rather far from the truth. I propose, however, now that I have explained their manufacture, to use them later on for comparing different tvpes of full-size aeroplanes. I have attempted next an investigation into the optimum position for the main spars of an aerofoil. I have assumed the usual practice of two spars only, though there is an interesting field for further investigation as to whether it might not be of advantage to use a greater number. For a very highly loaded aerofoil it might prove economical to employ more than two spars, instead of increasing the number of points of support (Plate II.) I assume that the spars are solid spruce of one section, and that the section is of the proportions here given. It is doubtful, but again a subject for further investigation, whether there be any other form of cross section more economi cal than the I. and I do not think that the spindling of an I-section can be made appreciably greater than that shown here without fear of weakness in shear along the neutral lamination. The maximum depth of spar is governed by the aerofoil section, it is taken hereafter as equal to thickness of section on vertical centre line of spar, minus one-tenth of maximum thickness of section—to allow for thickness of flanges of ribs. I have worked out tables of " spar values," each table being for one position of front spar and for six different positions of rear spar. One sample table is shown here, that in which front spar centre is at .1 of chord length from leading edge, and rear spar centre is taken at .5, .54, .58, .62, .66 and .7 of chord length from leading edge, in turn. The methods used in getting out these tables are as follows : Centre of pressure taken as at . 28 for load on front spar, and as at .6 for load on rear spar; " factor of safety " for rear spar taken as . 75 that for front. Calling the depth of spar at . 14 chord from L.E. unity, the section value, or " figure of merit," for a spar at any other point is taken as = (2d' - d) •+• 3, where d is the depth of spar expressed as a fraction of depth of spar at .14. The reason for this value, which is certainly open to criticism, is that for a constant form of spar section the moment of inertia varies as the fourth power of the depth and the sectional area as the square, and I have assumed that normally about two-thirds of the stress applied to the spar is due to bending and one-third to direct compression. In the left-hand space in the table, then, we have the " figure of merit " for the front spar, at its fixed position for this particular table. Column 1 gives position of centre of rear spar, column 2 gives depth of rear spar (as fraction of depth of spar at .14 of course), column ^ gives values for square of rear spar depth, column 4 gives values for fourth power of rear spar depth, and column 5 gives " figure of merit " for rear spar. Column 6 gives value for load on front spar when C.P. is at .28, expressed as a fraction; of total load on both spars. Column 7 gives load value divided by " figure of merit" value for front spar, and this is called " required moment of inertia value." We now turn to the curves of value for standard I cross section, and find where this " required moment of inertia value," set up vertically to scale, touches the curve of " values for IV." Dropping a vertical line on the base line from this point, we get firstly a reading on the base line for bjd value required for cross section, which value is set down in column 9 ; secondly, we measure to scale the value of cross sectional area on the " values for A " curve, which value is put down in column 8 and called " correspond ing value on A curve." Column 10 gives a value for front 76
Sign up to
Flight Digital Magazine
Flight Print Magazine
Airline Business Magazine
E-newsletters
RSS
Events
