Rising thermal dissipation from modern electronics has increased the challenge of cooling using conventional heat sinks. In addition to fans and blowers, focus is turning to active cooling devices for augmenting performance. A piezoelectrically-actuated synthetic jet array is one under consideration. Synthetic jets are zero-net–mass-flow jets realized by a cavity with an oscillating diaphragm on one side and an orifice or multiple orifices on the other side. They generate highly unsteady jetting flows that can impinge upon heated surfaces and enhance cooling. However, the synthetic jet actuation components might interfere with other components of the electronics module, such as the fan, requiring a displacement of the cavity center from the jet array center. Herein, heat transfer enhancement by an inclined piezoelectrically-actuated synthetic jet arrangement in a heat sink for electronics cooling has been experimentally and numerically studied. A wedge-shaped platform is designed to introduce the jets with an inclined configuration into the finned channels of the heat sink. The unit is inclined to avoid interference with other components of the module. The penalty is described in terms of velocities of jets emerging from this wedge-shaped platform, compared to those from an aligned cavity-orifice design. Effects on heat transfer performance for the heat sink are documented. The jets are arranged as wall jets passing over heat sink fins. The experimental study is complemented with a numerical analysis of flow within the synthetic jet cavity. Optimization is done on the number of jets against the penalty on jet velocity for obtaining maximum cooling performance. The jets are driven by piezoelectric actuators operating at resonance frequencies of 700–800 Hz resulting in peak jet velocities of approximately 35m/s from 92, 0.9 mm × 0.9 mm orifices. The results give guidance to those who face a similar interference problem and are considering displacement of the synthetic jet assembly.

The effects of an engine-representative combustor exit temperature profile and different disc cavity leakage flow rates on endwall adiabatic effectiveness distributions and passage temperature fields in a high pressure turbine rotor stage of a gas turbine are experimentally documented. The measurements are made on a stationary linear blade row cascade with an axisymmetrically-contoured endwall of modern engine geometry and with engine-representative approach flow thermal and fluid mechanics characteristics. The measurements give insight into mixing of coolant emerging as leakage flow and combustor liner coolant mix with hot core gases ahead of the airfoil row. Reported results are thermal fields in the passage, adiabatic wall temperatures and adiabatic effectiveness values in using an engine-representative approach flow temperature profile and with approach flow temperature profiles with 1) no coolant in the approach flow (flat profile) and 2) coolant only within 10% of the span (approach flow profile with a thin thermal boundary layer).The results give insight into mixing between the leakage flow and the mainstream passage flow and its effects on endwall cooling. The results demonstrate that for the conditions studied; much of the endwall cooling is contributed by the coolant in the approach flow. This is an important result that has previously not been well documented.

The flow field in the passage of a high pressure gas turbine is quite complex, involving strong secondary flows, transverse pressure gradients and strong streamwise acceleration. This complexity may have an adverse effect on cooling of the hub endwall, which is subjected to high thermal loading due to the flat combustor exit temperature profile of modern low-NOx systems. Therefore, given material limitations, better cooling management techniques that can be included with certainty in new gas turbine designs are needed. In the present study, film cooling has been investigated experimentally in a stationary linear cascade. The flow is representative of a high pressure gas turbine rotor with combustor liner coolant introduced to the approach flow. Focus is on the endwall axisymmetric contouring and the cooling effect of leakage flow bled from the compressor through the stator-rotor disc cavity. Two endwall contours, ‘shark nose’ (gradual slope over a larger distance) and ‘dolphin nose’ (steep slope over a shorter distance), are considered and comparison is made under conditions of three mass flow rates (MFR) of leakage, 0.5%, 1.0% and 1.5% of the approach flow rate. The performance of both endwall contours is compared at different streamwise locations in terms of adiabatic effectiveness values over the endwall. This study gives enhanced insight into the physics of coolant flow mixing, migration and subsequent coverage over the endwall. The results show the cooling effects of the contoured shapes over a range of leakage flow rates in the strong secondary flow environment. It is found that the leakage flow plays a crucial role in enhancing coolant coverage over the endwall. To add to our knowledge of mixing effects, detailed thermal field data are taken in the leakage flow discharge region. Doing so helps explain the behavior of the flow as it is ejected into the passage and interacts with the mainstream flow.