FlightGlobal.com
Home
Premium
Archive
Video
Images
Forum
Atlas
Blogs
Jobs
Shop
RSS
Email Newsletters
You are in:
Home
Aviation History
1956
1956 - 0558.PDF
558 FLIGHT THE SUPERSONIC TURBOJET . . . It is logical that most of the next generation of military aircraftshould be powered by developments of engines already in service. As far as work-capacity is concerned, good engines can be"stretched" at least as much as can airliners, although the process may be less evident externally. Several methods for increasingthrust in a turbojet come readily to mind. Most obvious, perhaps, is to increase the rate at which airpasses through the engine; for, other things being equal, the thrust varies directly with the rate at which the air is handled.Another self-evident solution is to give each particle of air greater acceleration. This can be done by increasing the maximumpressures and temperatures in the thermodynamic cycle of the engine, or by reheating the air in an afterburning tailpipe. It is to be noted, however, that too much increase in pressureratio can actually lead to a decrease in thrust. This results from the reduction in jet-pipe pressure and hence in thrust per poundof airflow. Such an apparent paradox is particularly important at high supersonic Mach numbers, as is revealed by the set ofcurves on the opposite page. An increase in the mass flow in an existing design of engineis normally obtained by one of two methods: by increasing the annulus area of the compressor and so improving the swallowingcapacity of the engine, or by running the engine at higher r.p.m. The latter choice, although sometimes practical without intro- 1-4 13 •312 IIIb §•9 UJ £* I'7 •6 800^h 4O Sb<XyA DESIGN / \PRESSURE RATIO SO 6O 70 8O 9O Id SPECIFIC THRUST (ib per Ib/sec oirtlow) Generalized turbojet curves of performance at maximum r.p.m., sea-level static. Assumptions: compressor and turbine 90 per cent efficient (potytropic); total duct losses, five per cent. ducing a major design modification, increases the operatingstresses throughout the engine and particularly in the blading. Nevertheless, substantial gains in thrust, together with improvedpressure ratios, can result from an increase of only two or three per cent in the governed speed. For really worth-while increases in mass flow there is noalternative but to re-design the engine to improve its swallowing capacity. Most turbojets at present in service were designed toconservative hub/tip ratios (radius of the intake hub divided by overall intake radius). It is possible to increase the annulus areaeither by opening up the outer diameter or by reducing the diameter of the hub. In either case the effect can be remarkable;examples are the Sapphire, in which the thrust was increased from just over 7,000 lb to not far short of 10,000 (Sa.4) solely by thismeans, and the J47, which, during re-design to J73 standard (of similar basic dimensions), was persuaded to increase its outputfrom 5,800 to 9,500 1b largely by improving the swallowing capacity. There are, of course, penalties to be paid. The greater lengthof the compressor blades has been known to intensify mechanical problems, although this is by no means a sine qua nan. The in-creased pumping horse-power necessitates a larger turbine, which, unless an additional stage is added, can tend to become of em-barrassing diameter. It should also be noted that increases in mass flow are possible only with the co-operation of combustionengineers, who must persuade correspondingly more fuel to burn efficiently in roughly the same size of chamber. Increased pressure ratio at unchanged r.p.m. is rarely resortedto except where a gain in efficiency is required. The easiest way to increase pressure ratio is to run the engine at higher speed,and the N.A.C.A. have recorded a 15-per-cent rise in pressure ratio as a result of increasing the compressor tip-speed from 1,000 to 1,400 ft/sec. The penalty in this case is that, not only arethe blade stresses markedly increased, but excessive velocity will result in part of the flow through the compressor being transonic.In future engines, however, this may be unavoidable, as will presently be discussed. Were other factors to remain constant (which in practice theycannot), increased pressure ratio would normally result in marked improvement in specific fuel consumption, and a high pressureratio is sometimes adopted solely on mis account, in spite of the cost of greater weight and complexity, and possibly poorerhandling characteristics. In certain cases a slight improvement in specific power will be gained as well. As the flight Mach num-ber is increased the ram compression (compression of the intake air resulting from the reduction of kinetic energy) becomes ofincreasing importance, and it completely dominates the thermo- dynamic cycle at high supersonic speeds. This hypothetical process of increasing the output of existingengines is unlikely to require any major change in the design of the combustion system. Certainly this part of the engine isdifficult to make any shorter, for enough axial length of chamber must be provided to make sure that all the fuel is burned beforethe turbine is reached. In spite of steady improvements in com- bustion intensity this problem is accentuated by the probableincrease in flow velocity which would accompany any of the usual forms of increasing the thrust. A great deal can be doneby detail re-design; in fact, certain American turbojets have been run fitted with combustion chambers capable of passing doublethe original design mass flow while occupying the same bulk and introducing only a quite-acceptable increase in pressure drop.There are, in fact, many who believe that great improvements in all factors, particularly chamber length, will be realized duringthe next few years. Immediately behind the combustion system is mounted theturbine itself, and, more than any other component, the turbine is a fundamental stumbling block in the uprating of an engine.It can generally be assumed that the turbine of a basic engine will already be worked fairly close to its limits. An increase inmass flow can be accommodated by increasing the blade length; this alone may provide sufficient extra shaft horse-power, but it ispossible that an additional stage may have to be added. Con- siderably more unyielding and restrictive are the limits set onmaximum gas temperature. Hotter flow through the compressor, combustion system or jet pipe can generally be accommodatedwithout too much trouble, but even a slight increase in turbine- inlet temperature results in progressive reduction in the life ofthe turbine. Even in engines with an expendable "short-life" application gas temperatures cannot be increased above verydefinite limits. In practice, changes in turbine design can be made to permitthe unit to be driven by a hotter gas flow. Careful mechanical design can do much to ease the maximum stress levels andsmooth out points of peak temperature. Furthermore, all kinds of specialist firms and agencies are relentlessly searching formaterials with better mechanical properties (particularly creep strength) at elevated temperatures. When an improved bladematerial is introduced it can permit either greater mechanical stress (resulting from increased flow or r.p.m.) or a higher turbine-inlet temperature; invariably it is the latter course which is taken. Generalized curves for a turbojet of 90 per cent efficiency cruising at 500 kt in the stratosphere. Assumptions: top tem- perature and compressor total energy rise, 90 per cent of design values; ram efficiency, 90 per cent; duct losses, five per cent. DESIGN PRESSURE RATIO 2O JO 4O SO 6O SPECIFIC THRUST (I b per Ib/sec airflow) 7O
Sign up to
Flight Digital Magazine
Flight Print Magazine
Airline Business Magazine
E-newsletters
RSS
Events
