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Aviation History
1947
1947 - 0959.PDF
JUNE 12TH, 1947 FLIGHT 557 CAS TURBINE COMBUSTION io the combustion zone is that corre-sponding to an overall mixture of cor- rect proportions for complete combus-tion with a minimum of excess oxygen. It is fairly certain that a substantial gaincould be obtained were it possible to use for the flame tube a refractory whichcould operate at a temperature of 1,600 deg C in place of the present workingtemperature (approximately 750 deg C). TABLE III Fuel A fi C. D E F G D cri tonesc ip ion Aviation petrol ... Wide boiling range similar to Ameri- can J.P.2type ... Paraffinic. Of the kerosene boiling range A kerosene with the major'ty of its aromatics and sul- phur removed... Aviation kerosene Fuel oil " Tetranap " An aromatic fuel Specific gravity 0.720 0.774 0.777 0.S06 0.810 0.933 0.966 Net calorific value C.H.U.,' Ib 10,510 10,450 10,530 10,420 10,310 9,840 9,690 C.H.U. gallon 75,670 80,880 81,820 83,990 83.510 91,810 93,600 Variation of net calorific value with specific gravity. Unfortunately, lack of mechanicalstrength and the disastrous consequences of pieces of flame tube passing into theturbine have so far precluded any de- velopment of this kind for aircraft. The factor which can be varied overthe widest range is turbulence, but here again a limit is set by the maximumL.T.P. permissible. For a fixed value of the pressure drop the best design will be that which gives the minimum lamelength lor an agreed temperature distri- bution. When the flame length is fixedthe best design will be that giving the minimum pressure drop for the samequality of temperature traverse. Flame-tube Temperature Although from the standpoint of main-taining good combustion efficiency and a short flame length a high flame-tube tem-perature is desirable, it is found neces- sary to limit the working temperature toapproximately 750-800 deg C, under full load conditions. This is well remavedfrom the softening and oxidation point of the material (Nimonic 75), but themargin is necessary to cope with condi- tions of rapid acceleration when theengine may be appreciably over-fuelled. On a number of recent engines flame-tube cooling has been greatly assisted by the provision of shoulder holes, twotypes of which are shown in Fig. 1. A ring of small holes is provided to intro-duce a proportion of the dilution air as a sheath on the inside surface of theflame tube. The exit from each- combustion cham-ber is circular while the inlet to the turbine is of a segmental shape. It isnecessary to have detailed knowledge of both temperature and velocfjEy distri-butions at the turbine entry in order to assess both the combustion efficiency andthe L.T.P. in the system. Figs. 3 and 4 are typical contour diagrams. Whilethe aim is to obtain as uniform a tem- perature distribution as possible, somelatitude can be permitted if the highest temperature does not occur near theturbine blade roots where the metal is most highly stressed. With the highvelocities and high rates of heat trans- c1 U V A 8 C U. Aviation kerosene Petrol ... Tetranap Pool gas- oil. Aviation kerosene. Pool fuel- oil. 701 cetane 1. 2. 3. 4. 5. 6. (Aviation kerosene 7. 8. Properties « 2 .??U Visco i 0 deg . 3.42 0.74 2.3 8.5 3.42 180 c.s. pre- heated to 100 deg. C. 8.5 2.53 3.34 4.04 4.75 1.61 3.42 5.00 11.62 at—30deg. C. sity . c V Q .810 .729 .966 .847 .810 .942 .812 .825 .810 .834 .862 .774 .810 .777 .887 of Fuel • cen t led . 50 pe r disti l 219.5104 205 266 219.5 — 273 195.5 212.5 222.0 228.2 164.5 219.5 199.0 215.5 conten t nt b y •ht . Aromati c pe r c e wei | 15 10.5 100 19 15 — Nil. 23 1520 28 13 15 Nil 34 FABLE En £•3 V V Pert full S F 33 33 30 30 30 30 30 33 33 33 33 33 33 33 33 IV. gine Cc liver y Ib/in * s. Ai r d e pres s ab 17.5 » 16.4 ,. .. 18.5 ,, nditions. I» «-2 Ai r d temp . ( 38 •• 30 .. 41 t t" t. ,, s tlo w ec . Ai r ma s Ib/ s 1.18 •• 0.92 •• .. 1.81 fl ,, flo w gal/h i Fue l 6.1 • • 5.5 •• „ 9.45 ,. f ,, of burne i Typ e Simplex 1.35 FN » (Spill \Simplex Spill (Spill -. Simplex 11.35 FN Spill .. Simplex t ,, < i. c 20 •• 63 17 63 63 17 63 63 26 2626 26 26 26 26 26 3. efficien c er cent . B °-0 u 89 94 97 40 95 97.5 75 66 94 87 8676 84 85 87 86 70 fer from* gas to metal and the low ther-mal conductivity of the heat-resisting alloys, the heat flow along the blade doeslittle in the way of smoothing out any irregularity in temperature distribution.The range of an aircraft depends primarily on the amount of " availableheat " in the form of fuel that it can Fuels and performances in three contemporary chambers. Fig. 4. Velocity contour diagram atturbine entry (ft/sec.) carry. For a small aircraft such as afighter, where space is the limiting factor, a high calorific value per unit volume isdesirable, while if overall weight imposes the limit, then a high calorific value perunit weight will be advantageous. (Table III.) Fuel storage capacity isclosely linked with fire risk. In an air- craft carrier, if the flash point of a iuelis below 150 deg F, special tanks must be used. If the fuel flashes above 150deg F, special precautions are unneces- sary and about 50 per cent more fuel canbe carried. More important is the danger from firein the air or on crashing. The rapidity with which the inflammation spreads tothe main bulk of the fuel is closely linked witli its volatility. For fires inan enclosed space, such as a fuel tank, temperature conditions have a majoreffect. Under very cold conditions, .1 volatile fuel such as petrol will producean inflammable fuel-air mixture, while the mixture from a fuel such as kerosenewill be below the weak limit of inflam- mability. On the other hand, in a hotclimate the petrol vapour-air mixture will probably be too rich, while thatfrom kerosene will be in the inflam- matory range. A further property that needs atten-tion is the " freezing point " of the fuel. At a temperature some 5 deg to 10 degC lower than the " cloud point," when solid crystals begin to separate, the fuelwill no longer pour. Provisional specifi- cations for the " freezing point " of air-craft gas turbine fuels have been set at -60 deg C and -40 deg C, in U.S.A.and Britain respectively. The majority of existing engines havea compression ratio of approximately 4:1. Over a wide range of engine r.p.m.conditions combustion efficiency is nearly 100 per cent. At low engine speeds,such as idling, there is a falling off in efficiency, which is influenced by thetype of fuel used. The results of com- bustion efficiency determinations forthree chambers from contemporary jet engines at ground-level idling conditions,using various fuels, are given in Table IV. It is desirable to emphasize the factthat even at 40,000ft difficulties only arise when the engine speed is reducedbelow the cruising range. At speeds corresponding to cruising, climb orcombat conditions, there is no difficulty at such altitudes in maintaining a highcombustion efficiency and a short flame.
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