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Aviation History
1950
1950 - 1704.PDF
284 FLIGHT, 7 September 1950 285 ATMOSPHERIC FACTORS 782 MACH NUMBER 07 0-8 0-9 10 11 1-2 1-3 1-4 1-5 1-6 17 1-8 1-9 20 40 SI - 40D 500 600 700 800 900 1000 M. P. h 100 200 400 500 600 700 900 1000 1100 1200 1300 1400 1500M. P H. 1600 1700 1800 1900 2000 2100 .347 434 521 608 695 782 868 955 1042 1129 12)6 1303 KNOTS TRUE AIR SPEED 87 261 347 434 521 606 ' 695 782 868 955 1042 1129 KNOTS TRUE AIR SPEED 1216 1303 1390 1476 1563 1650 1737 1824 A LTHOUGH the effect which the atmosphere has upon aircraft performance is /% well known, we believe that the graphic presentation of certain of the fundamental and derived data is of value not only as a means of emphasizing points which are apt to be taken for granted but also as a means of providing ready references. Graph No. 1 gives the true speed equivalents for a variety of Mach number values from sea-level up to the tropopause. Above this height the temperature remains constant and as Mach number is variable directly as a function of tempera- ture, above the tropopause a given Mach number is equivalent to a fixed speed. A Mach number of 1.0 is, of course, the speed of sound in air. Graph No. 2 shows comparative values of Indicated Air Speed and True Air Speed for a variety of altitude levels, and the T.A.S. may be read off from the I.A.S.—or vice versa—simply by finding the intersection point of the relevant speed with the altitude line concerned. The Mach number lines have been included to permit the direct association of Indicated Air Speed with Mach number. Graph No. 3 shows the progressive drop in temperature from sea-level up to the tropopause, and the reduction of relative density p/p0) and relative pressure (P/Po) from sea-level up to 8o,oooft. In the I.C.A.N. standard atmosphere, freezing level is reached at 7,500ft, and at 36,000ft the temperature stabilizes at 55 deg C below zero—i.e., 99 deg of frost as measured by the familiar domestic Fahrenheit scale. It is worth noting that, at 25,000ft, the atmospheric density has dropped to less than half its value at sea-level, whilst pressure has dropped to less than 40 per cent of its surface value. At 40,000ft, the reduction in density and pressure is such that, even with 100 per cent oxygen, human life can no longer be sustained. Even if it could, however, the water content in the body would vaporize and begin to boil at 55,000ft, whilst at 60,000ft the blood would boil : the importance of cabin pressurization for flight at high altitudes is, therefore, readily apparent. Graph No. 4 has been designed to illustrate the constriction of speed range with altitude for a typical modern fighter aircraft. The normal (high-incidence) stalling speed (without flaps) has been taken as no m.p.h. and the critical Mach number as 0.8. For these purposes the Men- is assumed to be the compressibility shock- stall threshold ; it is, therefore, the limiting Mach number at which the aircraft remains controllable. Thus it can be seen that the normal stalling speed mounts steadily—in terms of true air speed—until at 76,500ft it intersects the 0.8 M line ; at this height the aircraft could no longer fly, since if it travelled any slower it would stall, and if it travelled any faster it would become uncontrollable. The broken lines represent the stalling speeds when the wing loading is doubled and quadrupled. This is of great importance in fighter aircraft, where the pilot may well be forced in combat to impose loads on his aircraft in excess of that produced by an acceleration of 4 g. Accepting the 4 g stall limit as a reasonable boundary, however, it is clear that the typical aircraft we have selected—which is representative of some of the best fighters now current—has what may be described as an effective fighting ceiling of little more than 45,000ft. This can only be lifted by increasing the critical Mach number and/or by reducing the normal unflapped stalling speed. The former is purely a matter of aerodynamic refinement, but the latter can be achieved either by reducing the wing loading, or by increasing the lift coefficient. Of these alternatives, the easiest is to reduce wing loading (though this is by no means easy) and it does appear almost inevitable that the trend in future fighter types will be toward the use of moderate wing-loadings. 80 75 70 65 60 55 50 45 40 35 30 25 ZO 15 10 5 SI T R 0 \ P C \ . c i T J P \ A r A P J A3J s \ 0 A F B Jm \ R 0 A T • E_j / UR \ E •• • R E V \ \ s •I R 0 LA t b A \ •• E L E N Tl S U A \ m AT 0 1 n Kb t 1 V T 2mk -k 1 I*A11 m m• • •••• •1 7 1-0 0-9- 0-8 0-7 06 05 04 03 0-2 01 RELATIVE D E N S I T Y %, awaT RELATIVE PRESSURE^ (5 +10 +15 M. P. H. :4TJ 15 15 43 ~8T 3t) 174 216 TEMP 'C 261 304 347 391 KNOTS TRUE AIR SPEED "431 478 521 567
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