Winglets are plane surfaces with certain thickness and different shapes. Winglets are used in aircraft to reduce wing tip vortex which is created due to differential pressure in between pressure surface and suction surface. In transonic axial compressor, rotor tip leakage vortex interaction with shock layer and shroud boundary layer leads to total pressure loss and initiation of stall phenomenon. Effect of tip winglets are investigated in compressor rotor cascade. Cascade investigation shows that rotor tip winglet are able to reduce total pressure loss due to tip leakage flow and blade passage secondary flow. Cascade studies are performed with winglet on blade suction side, pressure side and combination of both. From cascade studies it is revealed that suction side winglet are aerodynamically better than pressure side and combined winglets. Owing to favorable results of tip winglet on compressor cascade performance, it was assumed that tip winglets would enhance overall performance of transonic compressor stage with rotating rotor. Results of present CFD simulations have predicted both positive and negative effects of winglets. Effect of different winglet configurations on pressure side and suction side of rotor blade tip are investigated to analyze the compressor stage overall performance. Rotor tip winglets are able to increase stage total pressure ratio compare to the baseline stage without winglet. Stage with winglets have shown better performance in choke region. Winglets are able to vary rotor blade loading from hub to tip region. Presence of winglet has shown ability to reduce to total pressure loss in trailing edge wake region. Stall margin is decreased in compressor stage with winglets due to more blockage towards trailing edge in tip region.

Numerical studies have been carried out on the effectiveness of trailing edge Gurney flap on a transonic axial compressor rotor. The baseline geometry of the rotor blade was modified at the trailing edge by introducing Gurney flaps of varying depth and span-wise length, viz. 1 mm, 2 mm and 3 mm depth with 20% span length of Gurney flap from tip (designated as GF1-20, GF2-20 and GF3-20 respectively), and 1 mm depth with 50% and 100% span length (designated as GF1-50 and GF1-100 respectively). Geometric models of the compressor rotor without and with Gurney flaps were generated using CATIA V5 software and CFD simulations at 100% design rotor speed were carried out using ANSYS CFX software. Results have shown that the compressor total pressure ratio increased with increase in both depth and spanwise length of Gurney flap. Peak pressure ratio increased from 1.51 for baseline case to 1.58 for rotor GF1-100. However, the peak isentropic efficiency remained almost constant for various Gurney flap configurations, except for GF1-100 which showed a tendency for improvement in efficiency. The stall margin reduced with the introduction of Gurney flap and was lowest for configuration GF1-100 which gave highest peak pressure ratio. Higher blade loading with Gurney flap was responsible for lowering the stall margin. Analysis of the flow through the blade passages has shown clear formation of trailing end vortex structure in the presence of Gurney flap that resulted in bending of the streamlines towards suction surface of the rotor blade, with consequent reduction in flow deviation and increased flow deflection, and hence increased total pressure ratio.

A single stage torque converter consists of three elements — pump, stator and turbine. Pump and turbine are coupled by transmission fluid. Unlike a fluid coupling, however, a torque converter is able to multiply torque when there is a substantial difference between input and output rotational speed, thus providing the equivalent of a reduction gear. During its operation all these elements are subjected to centrifugal load, fluid pressure load and heat generated in transmission fluid. Overloading a converter can result in several failure modes, some of them potentially dangerous in nature: ballooning, blade deformation and defragmentation, overheating. In the current work a single stage torque converter, was modelled and analysed numerically for evaluating stress distribution and deformation. The engine operating speed at 2000 rpm was considered for analysis. For static analysis of torque converter components centrifugal load and fluid pressure load were considered. Analysis was carried out for six different speed ratios varying from zero to one. Variation of principal stresses (hoop stress and radial stress) and von-Mises stress has been discussed. Maximum stresses are found to be in pump at speed ratio of one and in turbine at speed ratio of zero. Maximum stresses are at shell core that is near to hub. Blade deformation in pump is maximum at coupling phase and in turbine it is maximum at stall condition. From these results it helps to predict the failure of torque converter components under different operating conditions.

In a transonic compressor rotor, tip leakage flow interacts with passage shock, casing boundary layer and secondary flow. This leads to increase in total pressure loss and reduction of compressor stability margin. Casing treatment is one of the passive endwall geometry modification technique to control tip leakage flow interaction. In the present investigation effect of rotor tip casing treatment is investigated on performance and stability of a NASA 37 transonic compressor stage. Existing literature reveals, that endwall casing treatment slots i.e., porous casing treatment, axial slots axially skewed slots, circumferential grooves, recirculating casing treatment etc. are able to improve compressor stability margin with penalty on stage efficiency. Turbomachinery engineers and scientists are still focusing their research work to identify an endwall casing treatment configuration with improves both compressor stall margin as well as stage efficiency. Hence in the current work, as an innovative idea, effect of casing treatment slot along rotor tip mean camber line is investigated on NASA 37 compressor stage. Casing treatment slot with rectangular cross-section was created along the rotor tip mean camber line. Four different casing treatment configurations were created by changing number of slots on rotor casing surface. In all four configurations casing treatment slot width and height remains same. Flow simulation of NASA 37 compressor stage was performed with all these four casing treatment configurations. A maximum stall margin improvement of 3% was achieved with a particular slot configuration, but without any increase in compressor stage efficiency.

In the present research work, effect of airfoil vortex generator on performance and stability of transonic compressor stage is investigated through CFD simulations. In turbomachines vortex generators are used to energize boundary and generated vortex is made to interact with tip leakage flow and secondary flow vortices formed in rotor and stator blade passage. In the present numerical investigation symmetrical airfoil vortex generator is placed on rotor casing surface close to leading edge, anticipating that vortex generated will be able to disturb tip leakage flow and its interaction with rotor passage core flow. Six different vortex generator configuration are investigated by varying distance between vortex generator trailing edge and rotor leading edge. Particular vortex generator configuration shows maximum improvement of stall margin and operating range by 5.5% and 76.75% respectively. Presence of vortex generator alters flow blockage by modifying flow field in rotor tip region and hence contributes to enhancement of stall margin. As a negative effect, interaction of vortex generator vortices and casing causes surface friction and high entropy generation. As a result compressor stage pressure ratio and efficiency decreases.