The performance of the compressor blade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as corner flow separation between the wall and the blade. The present work is focused on the studying the effects of Vortex Generator (VG) on NASA Rotor 37 test case using Computational Fluid Dynamics (CFD). VG helps in controlling the inception of the stall by generating vortices and energizes the low momentum boundary layer flow which enhances the rotor performance. Three design configuration namely, Counter-rotating, Co-rotating and Plow configuration VG are selected based on the improved aerodynamic performance discussed in reference [1]. These VG are located at 90% span and 42% chord on suction side surface of the blade. Among the three configurations, the first configuration has greater impact on the end wall cross flow and flow deflection which resulted in enhanced numerical stall margin of 5.4% from baseline. The reasons for this numerical stall margin improvement are discussed in detail.

The increasing demands of better efficiency of modern advanced gas turbine require higher turbine inlet temperatures, which gives great challenges to turbine blade designers. However, the temperature limits of turbine blade material are not high enough to ensure its survival in such incredible operating temperature. Hence, both internal and external cooling approaches have been developed and widely used in today’s turbine blade. To internal cooling problems, a variety of cooling enhancement approaches, such as impingement and turbulators, are employed in order to meet the different needs in leading, middle and trailing region. One of the most critical parts in turbine blade is trailing edge where it is hard to cool due to its narrow shape. Pin-fins are widely used to cool the trailing edge of rotor and stator blades of gas turbine engine. Pin-fins offer significant heat transfer enhancement, they are relatively easy to fabricate and offer structural support to the hollow trailing edge region. The flow physics in a pin-fin roughened channel is very complicated and three-dimensional. In this work, we have studied the effect of channel orientation on heat transfer in a rotating wedge-shaped cooling channel using numerical methods. Qiu [1] studied experimentally heat transfer effects of 5 different angles of wedge shaped channel orientation for the inlet Reynolds number (5100 to 21000) and rotational speed (zero to 1000 rpm), which results in the inlet Rotation number variation from 0 to 0.68. They observed that compared to the non-rotating condition, there is about 35% overall heat transfer enhancement under highest rotation number. The above said results are validated using current studies with Computational Fluid Dynamics (CFD) revealed that rotation increases significantly the heat transfer coefficient on the trailing surface and reduces the heat transfer coefficient on the leading surface. This is due to the higher velocities associated with the converging geometry near trailing surface.

The operating range of any compressor is controlled by Surge and Choke. Surge occurs at lower mass flow rates with large pressure fluctuations and flow reversals, while choke occurs at higher mass flow rates when the flow rate reaches the limit which compressor can discharge. Ported shroud is a cost effective casing treatment that can greatly improve operating range of centrifugal compressors. By removing the stagnant and reverse flow from shroud wall boundary-layer region and recirculating it to impeller inlet, it has been demonstrated that larger range of operability can be achieved without much loss on compressor efficiency. This paper demonstrates the improvement of a centrifugal compressor operational range with ported shroud configuration.A series of CFD simulations were carried out with open source centrifugal compressor geometry (NASA HPCC 4:1) to create performance characteristics/speed-lines. The CFD methodology and practices were validated by comparing the results with the experimental data. Performance evaluation of ported shroud configuration is done with respect to solid shroud.Ported shroud compressor is proven to give higher choke mass flow and also a better surge margin compared to the Solid shroud model. The phenomena of in-flowing and outflowing port have also been demonstrated. Emphasis was given to understand how ported shroud helps to achieve a better performance. A design optimization study has also been carried out in order to establish the optimum ported shroud configuration. Design parameter such as port location has been selected and the effect of this parameter on the performance of the compressor is studied using CFD. Optimum port geometry was proposed.

Computational analyses are carried out to predict the effect of the hub leakage flow from the inner banded stator cavities on the overall performance of an axial high pressure compressor. The results of a full fidelity simulation model, which includes cavities, sealing and main flow, is compared with models without the hub leakage flow. The CFD prediction confirms a significant effect from the hub leakage on the performance of the multistage axial compressor. A simpler and faster CFD modeling technique is explored and discussed for the modeling of the effect of the hub cavity leakage. In this approach, costly details of the stator cavities are ignored. Transfer functions or correlations are derived from the 1.5 stage (IGV-Rotor1-Stator1) of a multistage compressor having inner banded stator cavity. These correlations are used as boundary conditions to the primary flow path of the compressor while carrying out the simplified CFD models for the complete 4 stage compressor model. A MINITAB study is conducted to understand the most influencing parameters and their interactions in deriving the correlations. Analyses are then carried out using the simplified model with the hub leakage boundary conditions derived from these correlations. The CFD results of the full fidelity simulation and the simplified models are compared in this paper.