Balamurugan Srinivasan

Numerical Studies on Effect of Channel Orientation in a Rotating Smooth Wedge-Shaped Cooling Channel

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.Copyright

Effect of Rotation on Heat Transfer and Flow Field Inside Leading Edge Cooling Passage Using Impinging Jets

In this paper, the effect of rotation on impingement cooling on the internal surface (profile) of the leading edge region in a turbine rotor blade is investigated using Computational Fluid Dynamics (CFD) simulations. The flow domain is obtained by stretching the middle cross section of the blade. The simulations are performed for 3 different leading edge profiles to increase the heat transfer rate in the cooling flow passage in stationary domain. In all the profiles, the nozzle position and Mach number of cooling fluid is kept constant at E/D = 4.32 and 0.4 respectively. In the profile 1, the suction side profile is modified to facilitate vortex shift so that it may reduce the crossflow effect which will enhance the Nuavg. But Nuavg reduced by 1.9% when compared to base case. In the profile 2, the coolant flow passage is smoothened at the apex to reduce dead zones and to enhance spreading of the jet. In this case, a 3.48% increase in Nuavg is obtained. Based on the analysis of velocity contours in the profile 2, the leading edge profile is further modified (profile 3). This resulted in 5.37% increase in the Nuavg. The effect of rotation on the Nuavg for different Mach numbers (M) in the profile 3 and base case are studied. When compared to stationary domain Nuavg is reduced for base case and profile 3. In rotating domain profile 3 shows improvement for different M when compared with base case. For M = 0.2, the Nuavg is increased by 2.7%, for M = 0.4, Nuavg is increased by 3.98% and for M = 0.6, it is negligible.Copyright

Cooling Efficiency Enhancement Using Impingement Cooling Technique for Turbine Blades

The effect of impingement cooling on the internal surface (cooling passage) of the leading edge region in a commercial turbine high pressure first stage rotor blade is investigated using Computational Fluid Dynamics (CFD) simulations. The flow domain is obtained by stretching the middle cross section (50% span) of the above mentioned blade. The simulations are performed for 3 different profiles in the cooling flow passage. In all the cases, the nozzle position and Mach number of cooling fluid is kept constant at E/D = 4.32 and 0.4 respectively. In the first case, the suction side profile is modified to facilitate shift in the vortex. This may reduce the crossflow effect, which will enhance the Nuavg. However, simulation results showed that Nuavg is reduced by 2% when compared to base case. In the second case, the coolant flow passage is smoothened at the apex to reduce dead zone and to enhance spreading of the jet. In this case, a 3% increase in Nuavg is obtained. Based on the analysis of velocity contours in the second case, the coolant flow passage in the third case is further modified to improve the spreading of flow. This resulted in 5% increase in the Nuavg when compared to base case.Copyright

Numerical Analysis on Axial Compressor Stage Performance With Vortex Generators

The performance of axial flow compressor stage can be improved by minimizing the effects of secondary flow and by avoiding flow separation. At higher blade loading, interaction of tip secondary flow and separated flow on blade surface is more near the tip of the rotor. This results in stall and hence decreases compressor performance. A previous study [1] was carried out to improve the performance of a rotating row of blades with the help of Vortex Generators (VGs) and considerable effects were observed. The current investigation is carried out to find out the effect of Vortex Generator (VG) on the performance of a compressor stage. NASA Rotor 37 with NASA Stator 37 (stage) is used as a test case for the current numerical investigation. VGs are placed at different chord wise as well as span wise locations. A mesh sensitivity study has been done so that simulation result is mesh independent. The results of numerical simulation with different geometrical forms and locations of VGs are presented in this paper. The investigation includes a description of the secondary flow effect and separation zone in compressor stage based on numerical as well as experimental results of NASA Rotor 37 with Stator 37 (without VG). It is also observed that the shape and location of the VG impacts the end wall cross flow and flow deflection [1], which result in enhanced stall range.Copyright

Numerical Investigation on the Effect of Vortex Generator on Axial Compressor Performance

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.Copyright