FlightGlobal.com
Home
Premium
Archive
Video
Images
Forum
Atlas
Blogs
Jobs
Shop
RSS
Email Newsletters
You are in:
Home
Aviation History
1955
1955 - 1576.PDF
FLIGHT, 28 October 1955 693 PROPULSION PROGRESS Notes on the Past, Present and Future, by a Rolls-Royce Engineer AS part of a symposium held by the University of Delft,t\ Holland, a paper entitled "A General Survey of the History J- A- and Development of Aircraft Propulsion" has been pre-pared by Mr. A. G. Newton, of the Aero Engine Division of Rolls-Royce, Ltd. He read his paper in Delft last Friday,October 21st. Roughly the first half of Mr. Newton's paper concerned itself with thegeneral requirements of aircraft propulsion, outlining the processes whereby latent heat energy in fuel is converted to propulsive thrust.The lecturer particularly commented upon the far-reaching con- sequences of the classic paper by Sir B. Melvill Jones in 1927, inwhich the proposition of the truly streamlined aircraft was first fully discussed. In Fig. 1 a number of aircraft are shown on a graph ofbrake horse-power required per 1,000 lb weight for given flight speed; and the latest types—such as the Meteor (1942)—lie close to theJones' ideal curve. During the first decade of development, aero-engines assumed agreat variety of forms, of which all but the air-cooled radial and the air- or liquid-cooled in-line and yee were gradually abandoned. Therefollowed 30 years of refinement, in which such things as supercharging, constant-speed airscrews and anti-detonant fuels brought progressivelygreater efficiency and higher output. Mr. Newton devoted much of this dissertation to an analysis of the air-cooled engine versus theliquid-cooled, particularly for civil applications. A special section of the paper recorded the development of therenowned Rolls-Royce Merlin liquid-cooled 12-cyl vee engine. Con- ceived in 1932, the Merlin passed a 100-hr test at 790 h.p. in 1934,which power was steadily increased to 1,030 h.p. (for a weight of 1,335 lb) by 1939. During World War 2, the Merlin was producedin no less than 52 distinct Marks, to a total of 150,000 engines. By the end of the conflict a full type-test had been passed at a rating of2,200 h.p. and a specific weight of 0.8 lb/h.p. This was done prin- cipally by improving the supercharging to increase the rated altitude,and by developing two-stage supercharging coupled with intermediate charge-cooling; by various mechanical improvements; and by steadyimprovement in the anti-knock rating and other properties of the fuel. On 100-octane fuel the maximum boost pressure was 20 lb/sq in,which in the two-stage engine made possible a maximum output of just over 1,800 h.p. The addition of a further 6J c.c./gal (to a total ofHi c.c./gal) of tetra-ethyl lead increased the safe boost to 25 lb/sq in, and the power to just over 2,000 b.h.p. This was equivalent to some30 m.p.h. increase in speed for the Spitfire at sea level. It was finally found that a further considerable increase in knock rating could beachieved by adding 2J per cent of mono-methyl aniline, resulting in an octane rating of 150. By 1945 this fuel was universally used bythe R.A.F. and U.S.A.A.F. on Merlins, and with it a short test was run at 2,400 h.p. Development work was also carried out on turbo-super-chargers, with which the maximum Merlin output could have reached 3,000 h.p. This work was terminated by the cessation of hostilities.Turning to jet propulsion, Mr. Newton observed that the almost complete swing away from the centrifugal compressor in favour ofthe axial in no way invalidated the correctness of the design-thinking of Sir Frank Whittle. At the time that the early British engines weredesigned the aim was to get engines running and into the air, and the centrifugal offered far more prospect for the accomplishment of thisgoal. Nevertheless, even the early turbojet centrifugal was quite a big step forward, in view of the required pressure ratio of 4 : 1 at anefficiency of 75 per cent (existing piston-engine compressors were then working at about 2 : 1, with rather lower efficiency). In spite of thelack of rig-testing facilities, Whittle achieved this target, and also suc- ceeded in driving the compressor with a turbine of higher efficiencythan any other known at that time. In discussing the axial engine versus the centrifugal, Mr. Newtonnoted that, whereas the thrust per square foot of frontal area had improved from 153 lb/sqft for the early Welland to 460 Ib/sq ft for theTay, current axial engines were giving better than1,400 lb/sq ft, albeit at slightly higher flame tem-perature. Furthermore it was generally acceptedthat, at higher pressure ratios, the efficiency ofthe axial was some six to seven per cent better thanthat of the centrifugal. An exception to thegeneral trend in favour of axials was the case ofthe small, low-pressure Fig. 7. A plot of power required (per unit weight) for varying tor- ward speeds, by aircraft of differing vintage. Sir B. Melvill Jones' line is **• ultimate. turboprop, in which the diameter of the compressor did not deter-mine the diameter of the engine, and in which the efficiency of the small axial compressor was not high. At this point, Mr. Newton produced four curves of general progressin gas turbine development, all plotted with time as the x-axis. The first, of turbojet take-off thrust, continued to increase steeply, andshowed no sign of flattening-off. The second, of s.f.c. for a typical flight case for turbojets, was a continuously lowering line, although largereductions in consumption had only followed the introduction of a new engine or component. A similar plot for turboprops highlighted thedifference in efficiency between centrifugal and axial engines, due partly to the higher pressure ratio and operating efficiency of the latter,which was in turn partly due to the increase in size of engine. The fourth curve was of specific weight of turbojets at a typical flightcondition; it was demonstrated that the rate of progress here was diminishing, partly on account of efforts to improve specific fuelconsumption. The lecturer then briefly outlined the development of his company'sDart turboprop, and of its successor the Tyne (pages 638-9 in our last week's issue), and then went on to comment on the fact that gas-tur-bine performance—particularly turboprop performance—was limited severely by restriction on maximum flame-temperature. He went on, "The problem of blade cooling is receiving considerableattention at the moment. The problem is, of course, to obtain adequate cooling with the minimum amount of air and the minimum pressureloss. Obviously the maximum cooling surface area is required for a given cross-sectional area of air flow passages up the blade, and thisis best achieved by having a large number of small-diameter holes. This in turn introduces manufacturing problems, but the present indicationsare that a satisfactory solution is possible. Further progress will neces- sitate either the injection of water into the cooling air (which isuneconomical for long periods) or the use of sodium in the blade in order to transfer heat to the root where it may be more easily dealt with. "The other method of raising flame temperature is by improvedmaterials. The advance in turbine-blade materials was relatively rapid during the early days of the turbine engine but it was quite evident, twoor three years ago, that the rate of progress in producing better materials was dropping. This is evident if one considers the permissible flametemperature which can be achieved at a particular time with the materials available, assuming a constant stress level. Thus, althoughthere has been a rise of 130 deg C since 1946, 60 per cent of this occurred in the first three years. This effect was not unexpected, andhas given impetus to research on blade cooling. , -•'., • . Compressor blading "The mechanical reliability of a compressor rotor blade is dependentto a far less extent than turbine blades on the creep strength of the materials used, or at least this is the case at the moment. The practicehas been to use aluminium blades for the early stages of a compressor (say up to 250 deg C under the most adverse flight conditions) thecriterion for these blades being fatigue. Initially aluminium bronze was used for the hot end, partly because it had adequate strength andalso because it had good forging properties enabling thin trailing edges to be maintained. As techniques improved stainless steels were intro-duced and, more recently, titanium blades have been incorporated in production engines. With the increase in flight Mach number withthe consequent increase in intake temperature, creep strength will eventually supersede fatigue as the deciding factor, and aluminiumblades may be completely replaced by titanium, and some of the stainless steels by nickel alloys. It is interesting to note that projectedengines may have high-pressure compressor blades working at tem- peratures higher than some present-day turbine blades. "From the military point of view, the Air Forces of the world willalways be demanding higher speeds and higher altitudes. Although we are led to believe that, once the transonic problems are solved,there is no real bar to very much higher Mach numbers, the effect of increasing Mach number on an aero-engine is still a continuous processand thus one would expect continuous progress. Fig. 2 shows the way in which the World's Air Speed record has increased over theyears. It will be seen that the increase in speed is more or less con- tinuous, the effect of the advent of the jet engine being to restore theline of progress when it was beginning to turn over, the average speed increases being about 15 m.p.h. per year, and perhaps a little more inrecent years. Admittedly these speeds are attained at sea-level, and the lower temperature at altitude should enable higher speeds to beattained there. The conclusion to be drawn from this curve is that speeds of about M = 1.2-1.3 should be attained at sea-level in about10 years' time, which for the same temperature are equivalent to about M = 2 in the stratosphere (this conclusion being, of course, applied totrue aircraft, i.e., not guided missiles). "Probably the main problem of high-speed flight is that of the intake,both from the point of view of efficiency and matching. At high Mach numbers the intake pressure ratio is much higher than that of the engineand thus any inefficiency in the intake will have a correspondingly large effect on the engine performance. Fig. 3 shows this effect fora typical high-supersonic engine and it will be seen that for a Mach number higher than about 1.5, the development of an intake moreefficient than the pitot intake is essential. It should be remembered that, at high supersonic speeds, even a low pressure recovery mayrepresent quite a high recovery of energy. For example, at M = 2.5, a pressure recovery of 75 per cent is equivalent to an energy recovery
Sign up to
Flight Digital Magazine
Flight Print Magazine
Airline Business Magazine
E-newsletters
RSS
Events
