Unsteady boundary layer transition on axial compressor blades

The unsteady laminar to turbulent transition of boundary layers on the blading of axial turbomachinery has been investigated experimentally. Measurements were made on and around the outlet stator row of a 1.5-stage axial compressor using thermal anemometry. Pitch-wise hot-wire probe traverses were made upstream and downstream of the blade row to define the freestream disturbance field; these clearly show the interaction of the inlet guide vane and rotor wakes. Boundary layer traverses were taken with a single hot-wire on the stator suction surface. A blade instrumented with 61 surface hot-film sensors distributed around the surface at mid-blade height was used to obtain high frequency response skin friction measurements. High speed data acquisition triggered by a rotor angle position reference signal allowed ensembles of data to be acquired corresponding to a specific set of rotor wakes. This data was processed in a variety of ways to reveal the nature of the unsteady transition process. This work corroborates and complements that of Halstead et al. (1995) and provides a useful resource for validation of unsteady flow calculations. A new intermittency detection algorithm has been developed to aid in the interpretation of the surface hot-film data. Intermittency is a more reliable measure of the transitional boundary layer state (and is easier to interpret) than other statistics such as skew, especially as the boundary layer nears separation. The intermittency detection algorithm has also been adapted to detect the region of relaxing flow commonly observed following turbulent spots where the skin friction relaxes from a turbulent level to the laminar level. This region is more stable and more resistant to separation than the undisturbed laminar layer and has a lower skin friction than the turbulent layer (on average). It may be possible to exploit the properties of this relaxing flow to improve blade profile designs. The experimental data show the transition process to be periodically unsteady and dominated by the presence of turbulent spots. Not only do the transition onset and completion points oscillate markedly in stream-wise position with the passage of wakes from upstream, but small regions of boundary layer separation may appear in the transition region during certain parts of the cycle. In flows with large separations through the whole cycle, the re-attachment point has also been observed to oscillate in stream-wise position. Using experimental data recently compiled by Gostelow et al. (1995) a new method of predicting the transition length in rapidly changing pressure gradients has been proposed and tested in typical turbomachine pressure distributions. The new method, which extends the commonly used Chen-Thyson technique, allows the spot propagation and growth parameters to respond to changes in the local pressure gradient. This can be especially important in turbine blade boundary layers where transition may commence in a region of favourable pressure gradient and then move into a region of strong adverse pressure gradient. Only the new method is able to predict the shortening of the transition zone due to this change in pressure gradient.