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Aviation History
1946
1946 - 1179.PDF
JUNE 13TH, 1946 FLIGHT 605 RECENT AERODYNAMIC DEVELOPMENTS OO7 •O O-4 encounter difficulties, and the greater isour ignorance of the precise nature of those difficulties.One reason for ignorance is that it is not possible effectively to use the windtunnel to acquire knowledge in this region. High-speed tunnelscan be used up to a Mach number of about 0.85, but ifwe attempt to go higher, they " choke," as we say.This means that the Mach number in the neighbourhoodof the model rapidly increases to unity and a shock waveextends right across the air stream. The reason for thisbehaviour is not far to seek. If the relation between theMach number and the cross- sectional area of a stream-tube is calculated, it is found that near a Mach number ofunity the area varies exceed- ingly slowly. Thus, for the10 per cent change of Mach number from 0.9 to 1.0, thechange of area, of the stream- tube is only 0.9 per cent. Theflow is thus in a very critical state when the speed is closeto that of sound. Even if an empty tunnelcould be persuaded to run at, say, M = o.g5, the introduc-tion of even a very small model would choke it, because of thechange of cross-sectional area of the stream brought about by the model. AsDr. Hilton graphically puts it, the size of the model that can be used shrinks tozero as the speed reaches the sonic speed and then grows again as the speed, in-creases above that of sound. Hence the wind tunnel is of little use in a range ofMach number from about 0.85 to 1.15. Yet we must fly through this range ifwe are to attain supersonic speeds, and so we must seek means of investigatingaerodynamic behaviour in this range. The obvious answer is to move themodel through the air, instead of blow- ing an air current over the model. Oneway of doing this is to drop heavy bodies from a great height and to makeobservations on them by means of in- ternal instruments communicating withthe ground by radio or salved after the fall, or by external observation such asis provided by radar tracking. This line of attack is being explored, and somefew observations have already been O-12 011 0-1O - O-O9 O-O8 0-07 O-O6; 0-O5 OO4 O-4 O-6 O-8 IO 1-2 14 1-6 1-3 Z'O MACH NUMBER Fig. 12. Slope of lift curve of super- sonic aerofoil. obtained. A second way is to projecta model through the air by means of a rocket and to use similar methods ofobservation. This also is being de- veloped. The last way is by flight itself, but it O-7 O-8 1-O 1-2 1-4 MACH NUMBER 1-6 1-8 Fig. II. Comparative drag curves against Mach for supersonic aerofoil and shell. / / 1A 11 \ J f / / y \ \, N > is evidently unwise to proceed far in thisdirection, on account of possible unex- pected dangers, until something hasbeen learned by the use of pilotless models. The second method is, ofcourse, precisely that used in ballistics, and it is significant that the drag ofshells is the only aerodynamic quantity which has hitherto been measuredaccurately over the whole range of speed from zero to more than three timesthat of sound. I have already said that we know agood deal about aerodynamic pheno- mena at speeds well above that ofsound. It is instructive to see how much we can guess of what will happenclose to that speed from our knowledge of what happens well below and wellabove it. I have prepared three diagrams to illustrate this point. In the first of these, Fig. 11, we havethe drag curve of a typical supersonic aerofoil over the whole range of speedsfor which it is known, there being a gap between the Mach numbers of approxi-mately 0.8 and 1.2. On the same figure is shown the established drag curve forshells. It is evident that we can make a very reasonable guess as to how theaerofoil drag curve will behave in the unknown region near the sonic speed;the dotted part of the curve is my own guess. But if we plot the slope of thelift curve instead of the drag, as is done in Fig. 12, we find that guessing isquite out of the question, nor have we in this case any rough guide fromballistic experiments. In Fig. 13 I have shown the elevatorangle to trim as a function of Mach number for two aircraft The subsonicvalues were measured in flight at the R.A.E.; the supersonic values were esti-mated by theoretical considerations and are certainly of the right order. Againit would be a bold man who thought he could join the two parts of the curve with any real degree of assurance.We are thus driven to the inexorable conclusion that only by experiments oithe kind I have suggested can we learn enough to make the design of aircraftfor flight in or through the transonic region a safe procedure; wecannot guess the answer from what we already know, ex-cept possibly in the case of the drag coellicient. It isthus necessary to develop a new and very difficult experi-mental technique if we wish to put our knowledge on afirm basis. It may not prove very difficult to measuresimple things like drag, but the detailed study of suchthings as control character- istics on a free projectedmodel moving at about 700 miles per hour may well taxour ingenuity to the utmost. There is also the still moredifficult problem of nutter in the transonic region. Evenin the range of high Mach number covered by the windtunnel nothing has yet been done to study this importantsubject, because of the utter inadequacy of the number ofhigh speed tunnels now avail- able. This matter of the tran-sonic region is urgent. It may be some time before aircraft are developed whichfly through the region and attain super- sonic speed, but our present-daymachines can get very nearly up to the speed of sound in a dive. It is now quifea Jong time since a Spitfire attained a » Mach number in the neighbourhood of0.9; and the development of jet propul- sion, together with continued aero-dynamic improvement of design, may be expected to increase this figure. With con-ventional designs of aircraft, however, the increase is slow, because of the extremelyrapid rise of drag near the sonic speed. But is this rapid rise of drag un-avoidable? We are beginning to think it* is not, or at least that it can be put of!till a still higher Mach number is reached, possibly a supersonic one. Very littleis known, basically, about drag in the region well above the critical Mach 2-O number 6 4° 2 £-2" 0 UJ o° j Dl-< w -6* B» 1 / A B / 1 V \ \ \ \ A B O-5 1-O MACH NUMBER 1-S Fig. 13. Elevator angle to trim for two different aircraft A and B. Subsonic values measured, supersonic estimated
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