An Experimental High-Temperature Turbine ( No . 1 : 26 ) • Part I . mThe Cooling Performance of a Set of Extruded Air-Cooled Turbine Blades
Cooling characteristics are given in terms of the variation of average blade temperature at mid-span with cooling-flow rate, Reynolds number, gas/cooling-air absolute temperature ratio, and, for the rotor, gas-stream incidence. Chordwise and spanwise variations of blade temperature are also presented. Measured pressure-drop characteristics indicate that in a gas-turbine engine, with the cooling-air inlet pressure equal to the turbine gas inlet total pressure, cooling flows roughly equal to 1.8 per cent and 3.0 per cent of the main gas flow could be passed through the nozzle and rotor blades respectively, giving average relative temperatures at mid-span of approximately 0.6 in each instance. Comparisons with theoretical estimates of blade temperature show that the external heat-transfer rates to the nozzle blades agreed well with those measured on similar blades in cascade tests. For the rotor blades the external heat-transfer rates appear to have been between those predicted from cascade and earlier turbine test data. From the measured effects of Reynolds number on blade temperature it seems that approximately equal areas of laminar and turbulent boundary layer existed on both nozzle and rotor blades, and that as the turbine inlet temperature was raised, the extent of the turbulent layer tended to increase.
An Examination of the Flow and Pressure Losses in Bfade Rows of Axial ' Flow ' Turbines )
~The design of axial-flo~ turbines has been hampered in the past by a lack of comprehensive data regarding pressure losses and gas deflections through rows of turbine blades. In the present report much of the available information relating to this subject is studied and analysed to determine magnitudes of gas pressure losses and deflections in a wide variety of blade rows and also to determine the separate influences of variables such as blade shape, blade spacing, gas Mach number, Reynolds number, incidence, etc. Of particular importance are the effects of secondary flows on the aerodynamic performance of a blade row and special attention is paid to ' secondary losses ', which form the difference between the total losses occurring in an actual turbine blade row and the smaller two-dimensional flow losses which are usually measured in a blade cascade tunnel. Eff&cts of blade tip clearance are also studied. Resulting from this analysis a number of empirical guiding rules and charts have been derived from which approximate values of the overall pressure losses and gas deflections in a range of blade rows can be deduced. A particularly significant feature brought to light is that the secondary losses can in many instances be large, the loss being generally found to be great when the blading has low reaction.
Reaction Tests of Turbine Nozzles for Supersonic Velocities

 A machine for testing turbine nozzles by the reaction method, which was described in a previous paper, was used to test a series of convergent-divergent turbine nozzles. The results of these tests, along with the test of a convergent turbine nozzle, are compared with each other and with analytical values. Two kinds of analytical values are employed, namely, the usual values obtained from an assumed isentropic expansion from inlet state to exhaust pressure, and the values obtained from the assumption that the processes in the nozzle are isentropic except for a normal shock which takes up a position in the nozzle such as to cause the stream to fill the exit area at the exhaust pressure whenever possible. This latter kind of analytical value involves no shock when the exit area can be filled at the exhaust pressure by means of isentropic processes only, or when the exhaust pressure is lowered so far that the shock has passed out of the passage. The agreement of the test results with the calculated results of this latter kind is good, and the disagreement which exists can be attributed largely to separation at the shock and to transmission of exhaust-pressure effects upstream through the boundary layer.