Not Much of an Engineer

by Stanley Hooker

  ram compression The increase in pressure of the fluid entering an inlet duct to an aircraft engine brought about solely from the velocity of the aircraft relative to the air. The effect becomes overwhelmingly important at high supersonic speeds.

  RCVs Reaction control valves, needed for the control of a jet-lift aeroplane at forward speeds so low that conventional control surfaces are ineffective.

  reheat Injection and combustion of additional fuel downstream of the turbine(s) of a turbojet engine, giving greatly increased thrust (especially for supersonic flight) at the cost of extremely high fuel consumption. In North America called afterburning. It calls for a specially designed jetpipe and a nozzle of variable profile and area.

  spool One complete axial compressor rotor, usually comprising a number of stages.

  SST Supersonic transport.

  stage Any one complete row of blades, all in the same transverse plane, forming part of a compressor spool. A centrifugal rotor forms a stage by itself, so the Merlin 60-series, with two superchargers in series, became a two-stage engine.

  supercharger Any pump for increasing the density of mixture supplied to a piston engine, especially at high altitude where, without a supercharger, the density and thus the power would fall dramatically.

  two-spool Gas turbine having two compressor spools in series, an LP spool feeding an HP spool.

  turbofan A turbojet whose upstream axial compressor blades are much larger than necessary to supercharge the core engine, the outer parts of the blades behaving as propeller blades to discharge air past the core engine to give thrust. In HBPR (high bypass ratio) turbofans the fan generates much more thrust than the small hot core jet.

  turbojet The simplest gas turbine, comprising a compressor, combustion chamber, turbine (driving the compressor) and a jetpipe giving propulsive thrust.

  zero stage An extra axial stage added at the front of an existing compressor, not only increasing the pressure ratio but also increasing the mass flow and hence the thrust and efficiency of an engine.

  Appendix I

  The engines

  Merlin

  Vee- 12 liquid-cooled piston engine, cylinders 5.4in bore by 6in stroke, capacity 1,649 cu in (27 litres). Designed by A. G. Elliott and others at West Wittering immediately prior to the death in 1933 of Sir Henry Royce. Early marks weighed 1,375 lb and were rated at 1,030 hp at 16,250 ft. After introduction of the improved supercharger the rated power (Mk 46, for example) went up to 1,415 hp at 14,000 ft, the weight being 1,385 lb. The first two-stage engine, the Mk 60, was rated at 1,125 hp at 29,000 ft, and weighed 1,550 lb. Late in the war many marks gave 2,030 hp for take-off, and 1,890 hp at 13,750 ft.

  Welland

  First British production turbojet; this was an original Whittle design at Power Jets, designated W.2B/23 after it had been inherited via the Rover Car Company in January 1943. Single double-sided impeller, ten reverse-flow combustion chambers and single-stage turbine. Flown in tail of Wellington in November 1942, at 1,250 lb thrust, and in E.28/39 research aircraft at 1,400 lb, in March 1943. Still at 1,400 lb rating, flown in F.9/40 Meteor on 12 June 1943. Passed 100-hour type test in April 1943, and put into limited production as Welland I at Barnoldswick in October 1943. Delivered to RAF from May 1944 at 1,600 lb, with 180 hours between overhauls, for a dry weight of 850 lb.

  Derwent I to IV

  Similar to Welland but with straight-through combustion chambers and greatly refined engineering design. Derived from Rover (Power Jets) W.2B/26 with Rolls-Royce designation B.37. Derwent I design started 1 April 1943, first run 29 June 1943, 100-hour type test at 2,000 lb completed 18 October 1943. Put into full-scale production at Newcastle-under-Lyme mid-1944 for Meteor III, with 500 Derwent I engines delivered during the war. Derwent II introduced Whittle’s W.2/700 impeller casing and gave 2,200 lb; first run 28 June 1944 at 2,200 lb. Derwent III was research engine for boundary-layer suction trials with Meteor. Derwent IV was rated at 2,400 lb and first tested on 22 February 1945, being flown in Wellington on 6 June 1945.

  Clyde

  This extremely advanced two-shaft turboprop, the RB.39, was designed from December 1943. It had a single-stage HP turbine driving a centrifugal compressor and an independent LP turbine driving the nine-stage axial compressor and reduction gear for contra-rotating propellers. The first runs in 1945 were at 2,560 shp, and using 11 engines the power was increased by 1947 to no less than 4,543 ehp (4,200 shp + 830 lb jet thrust) with water/methanol injection. The Clyde was a superb engine, and the first turboprop in the world to pass full military and civil type tests, but Lord Hives refused a production order (for Wyvern aircraft) because he doubted the potential market would be worthwhile.

  Nene

  The first totally new Rolls-Royce turbojet, designed wholly at Barnoldswick. First studies, designated RB.40, were for 4,200 lb, superseded by the B.41 rated at 5,000 lb. Similar in general layout to Derwent but only nine chambers, and disproportionate increase in airflow. Design started 1 May 1944, first run 27 October 1944, 100-hour type test at 4,000 lb on 10 January 1945, first flown (Lockheed P-80) on 21 July 1945 and type tested at 5,000 lb on 20 November 1945.

  Derwent V

  Totally new Derwent, designed as a photographic scale of the Nene. Design started 1 January 1945, first run 7 June 1945 and 100-hour type test completed five days later on 12 June; second type test completed on 14 July, and first flown (Meteor) on 5 August 1945. Rated at 3,500 lb, for weight of 975 lb.

  Trent

  This experimental turboprop, the RB.50, was basically an early Derwent fitted with a spur reduction gear to an offset shaft driving a small five-blade propeller. Rated at 1,230 ehp, this engine was first run on 20 June 1944 and became the first turboprop in the world to fly (in a Meteor) on 20 September 1945.

  Avon

  After the paper studies AJ.25 and AJ.50 (axial jet, 2,500 lb and 5,000 lb thrust), the AJ.65 was launched in 1945 and type tested at 6,000 lb in mid-1947. This engine, with single-stage turbine, weighed 2,400 lb, but the production RA.3 with a two-stage turbine was 125 lb lighter and was rated at 6,500 lb. Subsequently the Avon was developed in very many versions rated at up to more than 17,000 lb with reheat.

  Proteus

  This complex turboprop was designed from 1944 as a highly economical engine to power the Brabazon 2 and Princess. The original form had a 12-stage axial compressor and single HP centrifugal stage, eight slim reverse-direction combustion chambers, a two-stage compressor turbine and an independent single-stage LP turbine driving the propeller gearbox. First run in February 1947, and the massive Coupled Proteus for the two aircraft applications followed on the test bed in November 1949. Details of the difference between the target and achieved powers and weights are given in the text. In the completely redesigned Proteus 3 (700 series) the size and weight were greatly reduced, and the compressor and power turbines both had two stages, all except the final stage having shrouded blades. The redesigned engine was eventually put into service at 4,445 ehp, for a weight of 2,900 lb and with cruise sfc of 0.48 lb/h/ehp.

  Orion

  The BE.25 Orion was a completely new turboprop, which would have been an outstanding engine had the prospects justified its production. It had a 7-stage LP compressor, 5-stage HP compressor, cannular combustor, single-stage HP turbine, and three-stage LP turbine driving the LP compressor and propeller. Weight was 3,240 lb and rating a constant 5,150 ehp from sea level up to 15,000 ft (thus, potential sea level power was over 9,000 shp).

  Olympus

  Few engines of any kind have been developed to the degree shown by this superb two-spool turbojet. Design began as the BE. 10 in 1946 and the first run took place on 13 June 1950. This original BOl.l version weighed 3,520 lb and was rated at 9,750 lb with sfc of 0.766 lb/h/lb. From this engine was developed the Mk 101 which went into production for the Vulcan B.l bomber in 1955 at a rating of 11,000 lb. The Mk 102 of 1956 had a zero-stage, giving seven LP stages upstream of an unchanged 8-stage HP spool, and was rated at 12,000 lb for a w
eight of 3,700 lb. The Mk 104 of 1957 was uprated to 13,000 lb without change in size or weight. The largely redesigned 200-series had greatly increased airflow (approximately from 200 to 300 lb/s) and achieved higher pressure ratio despite using a 5-stage LP spool and 7-stage HP. The Mk 201 was rated at 16,000 lb and went into production at 17,000 lb for the Vulcan B.2. The 300-series engines introduced a zero-stage giving further increases in airflow and pressure ratio, and for many years the standard Vulcan engine has been the Mk 301 rated at 20,000 lb. From this engine was derived the Mk 320 (BOl.22R) for the TSR.2. These engines had fully modulated reheat and were cleared to fly for over an hour at Mach 2 and also for over an hour at full throttle at supersonic speed at sea level, ratings being 19,600 lb dry and 30,160 with full reheat. Finally, from this engine the Olympus 593 for the Concorde was developed, with SNECMA jetpipe (incorporating limited reheat) and noise-suppressor/reverser nozzle. The complete power plant weighed 7,465 lb and could easily give well over 40,000 lb (and often did), but the certificated rating is 37,700 lb.

  Orpheus

  One of the simplest of all modern turbojets, the BE.26 Orpheus was produced in large numbers with 7-stage compressor, cannular combustor and single-stage turbine at ratings from 4,230 to 5,000 lb, and in advanced forms with ratings of 5,760 lb or 6,810 lb dry and 7,900 lb with simplified reheat. Standard production engines typically weighed 810 lb.

  Pegasus

  After passing through metamorphoses the BE.53/2 Pegasus ran in August 1959 with a two-stage overhung fan (Olympus), 7-stage HP compressor (Orpheus), cannular combustor (Orpheus), single-stage HP turbine (Orpheus) and two-stage LP turbine. Thrust was 9,000 lb, obtained through two ‘cold’ front nozzles and two ‘hot’ rear nozzles, all vectored in unison to give lift or thrust. The Pegasus 2 of 1960 obtained over 11,000 lb using the higher-airflow Orpheus 6 HP spool. The Pegasus 3 of 1961 was rated at 13,500 lb using an 8-stage HP spool and HP turbine with two stages of improved blading. In 1962 the Pegasus 5 gave 15,000 lb using a new 3-stage fan, annular combustor and aircooled HP turbine blades. In 1965 the Pegasus 6 was rated at 19,000 lb using a titanium fan, new combustor with water injection, revised fuel system, two-vane nozzles and aircooled stage-2 HP turbine blades. Increased temperature in 1969 enabled the Pegasus 10 to give 20,500 lb, and the production engine for the Harrier in 1969 was the Pegasus 11 rated at 21,500 lb with increased airflow and various minor improvements. Since then many small improvements have combined thrusts up to 23,200 lb with extended life with reduced costs. PCB (plenumchamber burning) will boost thrust for later supersonic V/STOL aircraft.

  BS.100

  This larger vectored-thrust turbofan was developed for the supersonic P.1154 for the RAF and Royal Navy, until cancelled in 1965. It was an excellent engine, the final standard being the BS. 100/8 rated at 19,200 lb dry maximum cruise (sfc 0.615) and with PCB at 35,600 lb (sfc 1.16).

  RB211

  As finally certificated the original RB211-22B is flat-rated at 42,000 lb up to ambient temperature of 28.9°C, for an engine weight of 9,195 lb. The RB211-524 series typically weigh 9,850 lb and have ratings from 50,000 to 55,000 lb, with outstanding sfc and performance retention (according to airline customers no competitor engine can equal the -524 family in these respects).

  Appendix II

  Merlin Power and Jet Thrust

  When, in August 1940, I persuaded Hs (E. W. Hives, Chief of Rolls-Royce, Derby) to come to Lutterworth to see the Whittle engine running, it was producing between 800 and 1,000 lb of thrust.

  Neither Hs nor myself were very impressed by this for we were unable to visualise the comparison with ‘the driving force of 1,000 horse’ of the Merlin piston engine, which sounded so much more. I resolved to do the comparison, which involved only a simple calculation, but which was destined to have a dramatic effect upon the development of the jet engine.

  Thrust is a force which can be exerted on either a stationary or moving aircraft. In the latter case thrust horse power is produced. Thus 1 lb of thrust exerted on an aircraft moving at 33,000 ft/min, which is the same thing as 375 mph, is equivalent to 1 thrust horse power.

  Hence, when Hs first saw the jet engine, its 800 lb of thrust would have become 800 thrust horse power if flying at 375 mph in, say, the Spitfire.

  Now, the Merlin developed 1,000 brake horse power with which it rotated the propeller. Before this can be converted into thrust hp, it must be multiplied by the propeller efficiency, which in those days was a little less than 80%

  Thus 1,000 brake horse power gave, at the most, 800 thrust horse power when flying.

  The two engines were, therefore, broadly equal when flying at 375 mph, a fact which staggered both Hs and myself, and caused him to resolve to acquire Whittle’s jet engine at the appropriate opportunity, which in fact came about, as related earlier, in the autumn of 1942 when Rolls-Royce took over the engine’s future from the Rover Company.

  It is interesting to note that at the speed of the Concorde, which is 3.6 times 375 mph (1,350 mph), 1lb of thrust is 3.6 thrust horse power, and, thus, at 50,000 ft when the Concorde is cruising at Mach 2.05 and requiring a thrust of 10,000 lb from each engine, the power produced per engine is 36,000 hp, giving a total aircraft power of 144,000 thrust horse power from the four Olympus 593 engines, which is more power than that required for the aircraft carrier HMS Illustrious, which has four Olympus marine engines giving nearly 30,000 brake horse power apiece.

  In a span of 30 years, from 1940 to 1970, the power of jets made the gigantic steps from 1,000 to 36,000 thrust horse power, and still the end is not in sight.

  Not quite as dramatic, but of greater significance is the reduction in specific fuel consumption in the same period.

  The Merlin used 500 lb of fuel per hour to produce 800 thrust horse power, giving a specific consumption of 500 divided by 800 or 0.625 lb of fuel per hour per thrust horse power. This is equivalent to an overall thermal efficiency of 22% (including the propeller).

  The Olympus 593 at Mach 2.0 uses 1.20 lb of fuel per hour for every 1 lb of thrust or 3.60 thrust horse power. This gives a specific consumption of 1.20 divided by 3.60 or 0.333 lb of fuel per hour per thrust horse power. This is equivalent to an overall thermal efficiency of 42%, double that of the Merlin, and one of the highest figures ever achieved. For example, few, if any, of our most up to date and complex electricity generating stations achieve 40%

  Appendix III

  The Propulsive Efficiency of a Jet Engine

  Propulsive efficiency is defined as that proportion of the power produced by the engine which appears as thrust horse power on the aircraft.

  In the case of a variable pitch propeller, whose pitch can be continuously adjusted during flight to the optimum angle, the efficiency can be maintained at about 80% for all forward speeds of the aircraft until at high speeds in excess of 400 mph, when the Mach number at the propeller tip approaches and exceeds unity, the efficiency begins to fall dramatically.

  The thrust produced by a jet engine is given according to Newton’s law:

  where M is the mass flow of air through the engine, and Vj and Va are the jet and aircraft velocities respectively.

  The thrust horse power (thp) is obtained by multiplying the thrust by the aircraft speed, giving:

  The horse power that the engine produces is the rate of increase of the kinetic energy of the air as it passes through the engine.

  Thus, jet horse power (jhp) is:

  By definition, therefore, the propulsive efficiency ηp is:

  This formula shows that the lower the ratio of Vj to Va, the higher the propulsive efficiency, which is highly desirable. On the other hand, if we reduce the jet velocity to achieve this, the lower the thrust per unit of mass flow, and, consequently, the mass flow must be increased to compensate and restore the thrust.

  This is the principle of the High By Pass Ratio or modern fan engine, where the jet velocity is reduced and the mass flow increased by a factor of four or five to restore the thrust.

  T
ypical figures:

  (a) For a straight jet engine such as the Olympus in a subsonic aircraft such as the Boeing 707:

  (b) In a fan engine, such as the RB211 in the Boeing 747, the jet velocity can be reduced to 1,500 ft/sec.

  and this would result in the specific fuel consumption being reduced in the ratio 60:72 i.e. by 20%

  (c) For the Concorde, cruising at Mach 2.0 with Olympus 593:

  which shows the alternative way of increasing ηp, not by reducing the jet velocity but by increasing the aircraft speed.

  Appendix IV

  Supercharging the Merlin Engine

  The supercharger has two main functions, the first to increase the power at sea level by ‘boosting’ the induction pipe pressure, and the second to maintain this power as the aircraft climbs to higher altitudes where the air is less dense. For the latter, the throttle must be gradually opened during the climb (done automatically on the Merlin, where a constant induction pipe or boost pressure was maintained) until an altitude is reached where the throttle is fully opened (the full-throttle altitude), and thereafter the power falls roughly proportional to the falling air density.