Superior rotor tip geometries possess the potential to simultaneously mitigate aerodynamic losses and severe thermal loads onto the rotor overtip region. However, classical design strategies are usually constrained to a specific type of geometry, narrowing the spread of shape topologies considered during the design phase. The current paper presents two novel multi-objective optimization methodologies that enable the exploration of a broad range of distinct tip configurations for unshrouded rotor blades. The first methodology is a shape optimization process that creates a fully carved blade tip shape defined through a Bezier surface controlled by 40 parameters. Combined with a differential evolution (DE) optimization strategy, this approach is applied to a rotor blade for two tip gap sizes: 0.85% (tight) and 1.38% (design) of the blade span. The second methodology is based on a topology optimization process that targets the creation of arbitrary tip shapes comprising one or multiple rims with a fixed height. The tip section of the blade has been divided into more than 200 separate zones, where each zone can be either part of an upstanding rim or part of the cavity floor. This methodology was tested with a level-set approach in combination with a DE optimizer and coupled to an optimization routine based on genetic algorithms (GAs). The current study was carried out on a modern high-pressure turbine operating at engine-like Reynolds and high subsonic outlet Mach numbers. A fully hexahedral unstructured mesh was used to discretize the fluid domain. The aerothermal performance of each tip profile was evaluated accurately through Reynolds-averaged Navier–Stokes (RANS) simulations adopting the shear-stress transport (SST) turbulence model. Multi-objective optimizations were set for both design strategies that target higher aerodynamic rotor efficiencies and simultaneous minimization of the heat load. This paper illustrates a wide variety of profiles obtained throughout the optimization and compares the performance of the different strategies. The research shows the potential of such novel methodologies to reach new unexplored types of blade tip designs with enhanced aerothermal performances.

In this two-part paper, a novel test-case for transonic low-pressure turbines (LPT) is presented. The current study is a comprehensive report on the design, commissioning and testing of a high-speed LPT cascade. Part II reports the characterization of the aerodynamics at on- and off-design flow conditions. A detailed analysis of the steady aerodynamics of the highspeed low-pressure turbine blade is presented for a range of engine representative outlet Mach numbers from 0.70 to 0.95 and Reynolds numbers from 65,000 to 120,000. The blade 3D aerodynamics are characterized using an innovative set of traversable blades, enabling high-resolution radial measurements. A novel method to estimate the location of separation-reattachment based on pneumatic tap measurements is presented. The separation on the blade suction side is strongly influenced by the Reynolds number at the lower Mach numbers, while open separations were observed at transonic exit conditions independently on the Reynolds number. Downstream measurements by means of a five-hole probe, and pressure taps located in the passage endwall are employed to study the secondary flow development and structures at the cascade outlet. The results show that the losses follow different trends at high and low Reynolds numbers. The profile losses at Re = 70k decrease with increasing Mach number, contrary to what is observed for Re = 120k. The minimum secondary flow losses are found for an off-design condition with lower Mach number with respect to the nominal. The off-design comparisons presented in this paper indicate that at low Reynolds, operating at transonic outlet Mach numbers leads to beneficial effects on the performance.

In this paper the aerodynamics of an innovative multisplitter low pressure (LP) stator downstream of a high pressure turbine stage is presented. The stator row, located inside a swan necked diffuser, is composed of 16 large structural vanes and 48 small airfoils. The experimental characterization of the steady and unsteady flow fields was carried out in a compression tube rig under engine representative conditions. The one-and-a-half turbine stage was tested at three operating regimes by varying the pressure ratio and the rotational speed. Time-averaged and time-accurate surface pressure measurements are used to investigate the aerodynamic performance of the stator and the complex interaction mechanisms with the high pressure (HP) turbine stage. Results show that the strut blade has a strong impact on the steady and unsteady flow fields of the small vanes depending on the vane circumferential position. The time-mean pressure distributions around the airfoils show that the strut influence is significant only in the leading edge region. At off-design condition (higher rotor speed) a wide separated region is present on the strut pressure side and it affects the flow field of the adjacent vanes. A complex behavior of the unsteady surface pressures was observed. Up to four pressure peaks are identified in the time-periodic signals. The frequency analysis also shows a complex structure. The spectrum distribution depends on the vane position. The contribution of the harmonics is often larger than the fundamental frequency. The forces acting on the LP stator vanes are calculated. The results show that higher forces act on the small vanes but largest fluctuations are experienced by the strut. The load on the whole stator decreases 30% as the turbine pressure ratio is reduced by approximately 35%.

Tip leakage flows in unshrouded high speed turbines cause large aerodynamic penalties, induce significant thermal loads and give rise to intense thermal stresses onto the blade tip and casing endwalls. In the pursuit of superior engine reliability and efficiency, the turbine blade tip design is of paramount importance and still poses an exceptional challenge to turbine designers. The ever-increasing rotational speeds and pressure loadings tend to accelerate the tip flow velocities beyond the transonic regime. Overtip supersonic flows are characterized by complex flow patterns, which determine the heat transfer signature. Hence, the physics of the overtip flow structures and the influence of the geometrical parameters require further understanding to develop innovative tip designs. Conventional blade tip shapes are not adequate for such high speed flows and hence, potential for enhanced performances lays in appropriate tip shaping. The present research aims to quantify the prospective gain offered by a fully contoured blade tip shape against conventional geometries such as a flat and squealer tip. A detailed numerical study was conducted on a modern rotor blade (Reynolds number of 5.5 × 105 and a relative exit Mach number of 0.9) by means of three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) calculations. Two novel contoured tip geometries were designed based on a two-dimensional (2D) tip shape optimization in which only the upper 2% of the blade span was modified. This study yields a deeper insight into the application of blade tip carving in high speed turbines and provides guidelines for future tip designs with enhanced aerothermal performances.

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