Computational Studies on Tip Clearance Influence Regarding Flow and Heat Transfer Characteristics of Modern Aero Turbines (DTA)

This study addresses the study of the flow and heat transfer at the tip clearance region in aero turbines. Fluid dynamics at this region tends to be extremely complex. The flow is inherently unsteady and three dimensional due to interactions of the tip leakage flow/vortex with other flow structures, interactions between adjacent rows of blades and viscous effects including shock-boundary layer interaction. Large Eddy Simulation(LES) /RANS might be used in order to resolve the significant flow structures in time and space, to figure out turbulent flow parameters and apprehend the details of the unsteady flow. These are necessary steps towards investigating new methods for enhancing isentropic turbine/engine efficiency. A major source of losses in axial turbomachinery stems from the secondary flow imposed on top of the primary one, as the flow passes each blade row. According to well-known empirical correlations, such as Howell’s model, the secondary flow typically accounts for most losses when the flow conditions are close to the design point . Among the various forms of secondary flow encountered in turbomachinery, either compressor or turbine, the tip leakage flow accounts for a major chunk of the total losses, nearly 20% to 40% depending on the machine, while the corresponding loss in efficiency can be anywhere between 2% to 4% . The phenomenon appears more intense and performance-wise more important in turbines rather than in compressors, since the driving force of the tip leakage flow – the pressure difference between the blade’s suction and pressure side – is by far greater.
The area under consideration is extremely confined, the clearance gap between an unshrouded rotor blade tip and the outer-casing wall is ~ O(1%) of the blade span, corresponding typically to a physical length ~ O(1mm) or less. The direct loss associated with the flow passing through the clearance gap without change in angular momentum represents a fairly small fraction of the total tip leakage losses. The leakage, however, of the tip clearance region affects dramatically the mean flow, downstream and inwards to a considerable portion of the blade span. The leakage flow escaping from the pressure side separates (with possible reattachment) on top of the blade tip, and subsequently rolls up at the suction side forming the leakage vortex, which interacts with the other flow structures, such as the passive vortex at the central region of the row passage, the trailing edge vortices at the blade’s wake and the boundary layer at the annulus. Strong three-dimensionality is therefore introduced into the mean flow leading to losses associated with the kinetic energy of the velocity perturbations of the various flow structures, which diffuse downstream and eventually dissipate by viscous forces. The latter indirect loss mechanism (the mixing losses) produces the major fraction of the total tip leakage losses. More complexity is introduced when compressibility effects are present, for instance at high-subsonic or transonic tip speeds present at high pressure turbines. In such cases, shock-boundary layer interactions occur at the tip clearance region regarding both confining walls at the annulus and the tip. High tip speed also gives rise to the scrapping of the annulus boundary layer by the moving rotor blade, which subsequently results into the formation of the scrapping vortex at the pressure side. At low pressure turbines, on the other hand, periodical transition from laminar to turbulent flow affects the separation pattern on the blade tip surface. Additionally, the heat transfer at the tip is heavily affected by the tip leakage flow/vortex, with a subsequent impact on the cooling requirements of the region.
The above described perturbations in the velocity and the temperature fields are perceived by the succeeding stator blade row as an inlet distortion, making an impact to both the flow development and heat transfer and leading to a subsequent distortion to the nozzle row outflow, which affects the inlet of the next rotor blade row and so forth.
While it is neither possible nor intended to fully describe the flow field within the present proposal, the complexity of the aerophysics taking place in the vicinity of the tip clearance has already become apparent. The main flow features that build up the complexity can be summarised as follows:
1) The flow is inherently unsteady,
2) The geometries involved are complicated,
3) There is relative motion between the rotor and stator blade rows and between the rotor blade rows and the outer casing,
4) The flow periodically transitions between laminar and turbulent,
5) Shock-boundary layer interactions occur at high tip speeds.