Abstract:
Part I of this report describes an experimental, cooled turbine, designed and built for research into problems associated with the operation of turbines at very high gas temperatures, in particular, the problem of turbine blade cooling. The turbine structure and the associated test rig have been designed to permit testing of turbine blades over a wide range of turbine entry gas temperature from 900°K to 1600°K and entry gas pressure from two-thirds to 5 atmospheres giving a wide range of gas-flow Reynolds numbers to the blades. A further aim of the design has been to achieve rapid access to blades under test and to allow the testing of a moderately wide variety of both blades and cooling systems with relatively minor alteration to the turbine or test rig. Initial proving tests have been made to observe the mechanical behaviour of the turbine and test rig and to establish optimum coolant quantities required to cool the turbine structure. Following minor alterations and adjustments satisfactory performance was achieved over the major portion of the design speed range while operating at moderate gas temperatures, from 900°K to 1100°K (atmospheric turbine exhaust). Resuks suggest that the turbine structure should be sufficiently well cooled to permit the testing of blades over the full range of turbine inlet gas temperature envisaged. Part II of this report describes the results of an endurance test at a gas inlet total temperature of 1500°K. Eighty hours of running were completed at a rotational speed of 12,000 rev/min with a rotor-blade cooling-air flow, equal to 1.8 per cent of the main gas flow, supplied to the rotor assembly at an inlet temperature of 310°K. During this period three blades failed in a manner similar to that predicted, and other blades exhibited spanwise cracks. A method for estimating the redistribution of combined thermal and centrifugal stress in a non-uniformly cooled rotor blade due to creep effects is described and its predictions compared with observations.
