The results of a field test in a commercial turbine engine showed that we can remotely measure the temperature of engine components in operating engines using thermographic phosphors. The remote-measurement method exploits the temperature dependence of the characteristic decay time of the laser-induced fluorescence of thermographic phosphors. This paper summarizes recent work leading up to and including a successful test of the thermographic-phosphor method in an operating turbine engine.

The literature on the gas turbine already contains much information concerning the performance of the “constant-pressure” cycle both in its simple form and in its more complicated forms involving intercooled compression, reheated expansion, and heat exchange. Such information enables the design performance of a particular engine to be estimated, but there appears to be very little published work to help the designer to foretell what changes in performance are to be expected when particular engines operate at non-design conditions. What work there is seems to have been restricted to very simple engines, and is often further limited by the assumption of unalterable component performance characteristics. In this lecture we attempt to include in a general comparison of part-load performance characteristics not only the simpler designs but also some of the more complex turbine engines which will be needed for land and marine applications. At the same time, by considering, in appropriate cases, the influence of changes in the assumed component performance features on the part-load operation of an engine, a broadening of the basis of comparison is made possible. As a result of the work we shall describe we feel that, although as yet so little of the possible field of investigation has been surveyed, it is nevertheless possible to indicate the main guiding principles by which the comparative part-load performance of different gas turbine schemes may be assessed.

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SummaryThe current interest in supersonic transport aircraft makes it appropriate to consider the problems of propulsion nozzle systems for high speed flight. The discussion will be confined mainly to this field of activity.In the absence of external drag and weight factors, the ideal configuration to maintain optimum thrust at all flight speeds would employ a mechanically variable area ratio nozzle. By such means the jet flow could be fully and properly expanded at all times.A practical design to accommodate such variations is inherently complex, and so an alternative solution using aerodynamic techniques has been sought. It is necessary to remember that the success of whatever method is eventually chosen must rest on its ability to combine a high performance at cruise with very small loss during other critical phases of the flight plan.This paper describes some of the results of a nozzle research programme aimed primarily at solving the above problem, with an attempt to explain the philosophy behind the aerodynamic techniques tried. Some discussion is given of the experimental difficulties in establishing a sufficiently accurate standard of measurement for design point performance. In a few chosen configurations experimental results have been extensively backed by theoretical studies.The need to demonstrate off-design behaviour in the presence of an external flow field is stressed, with special reference to the misleading results often obtained from static tests.Finally, some attention is given to the question of base bleed, with its repercussions on the installation of the propulsion nozzle in an aircraft.

The achievement of high propelling nozzle performance in supersonic flight is vital to the success of any aircraft designed to cruise supersonically for prolonged periods. For such projects, the payload fraction is small compared with subsonic aircraft, and the ratio of gross to net thrust is high. As a result, the penalty in range/payload performance for a shortfall in nozzle gross thrust efficiency is inevitably severe.