S. Lavagnoli

Blade Tip Carving Effects on the Aerothermal Performance of a Transonic Turbine

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.

Three-dimensional (3D) inverse heat flux evaluation based on infrared thermography

Inverse methods in heat transfer are generally associated with the estimation of unknown heat fluxes on inaccessible surfaces, based on temperature measurements performed on the accessible walls. In the present application, an inverse heat conduction method was developed that uses temperature measurements performed by infrared thermography. The three-dimensional (3D) transient inverse heat transfer problem was solved by using an iterative regularisation with a conjugate gradient method. The inverse solver was coupled with a finite element commercial code. The efficiency of the method was demonstrated through numerical simulations and experimentally validated in a simplified set-up.

Heterogeneous Optimization Strategies for Carved and Squealer-Like Turbine Blade Tips

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 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 differential evolution optimizer, and coupled to an optimization routine based on genetic algorithms.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 RANS simulations adopting the 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.Copyright

Impact of a multi-splitter vane configuration on the losses in a 1.5 turbine stage

This study presents the influence of a multi-body architecture on the aerodynamic performance of a low-pressure stator. The novel design has been studied numerically in a one-and-a-half stage turbine by means of a three-dimensional Reynolds-averaged Navier–Stokes simulation. This numerical research compares the behaviour of low-pressure stator composed of two different vanes against a conventional axisymmetric single airfoil row. The computational fluid dynamics predictions were calibrated using experimental aerodynamic measurements. Loss generation mechanisms were evaluated for the conventional and multi-splitter cascades at nominal and off-design conditions. At design conditions, the novel stator and conventional designs show comparable performances. However, the performance is drastically reduced at off-design conditions due to the sensitivity of the structural vanes to flow incidence. This article addresses the performance limitation for the multi-splitter vane configuration and presents a new tool to analyse the non-uniform flow conditions associated with such novel design. This procedure should help researchers in addressing any non-axisymmetric design.

Aero‐design of transonic LH2 and LOX contra rotating turbopumps in an expander rocket engine

Purpose – Contra‐rotating turbines offer reduced size, weight, and cooling requirements, compared to conventional co‐rotating machinery. In spite of the associated mechanical complexity, their aero‐thermal performance is superior to conventional turbines, not only due to the elimination of stator blade rows, but also because lower turning airfoils can be implemented as a result. The purpose of this paper is to present a methodology to determine the optimum velocity triangles of the turbine, together with a two‐dimensional design and optimization tool to minimize the blade unsteady force using radial basis function network, coupled to a genetic algorithm. The proposed design methodology is illustrated with the aerodynamic design of a contra‐rotating two‐axis turbine, which is able to deliver the power necessary to drive the LOX and LH2 pumps of an improved expander rocket engine.Design/methodology/approach – This paper presents a methodology to determine the optimum velocity triangles of the turbine, toget...

Aerodynamic Analysis of an Innovative Low Pressure Vane Placed in a S-Shape Duct

In this paper the aerodynamics of an innovative multisplitter 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 field 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 HP turbine stage. Results show that the strut blade has a strong impact on the steady and unsteady flow field 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 approx. 35%.Copyright

Performance Robustness of Turbine Squealer Tip Designs due to Manufacturing and Engine Operation

This paper quantifies the changes in turbine performance due to manufacturing tolerances and profile degradation of the blade-tip region during engine operation. An extensive numerical study was co...

Analysis of the Heat Transfer Driving Parameters in Tight Rotor Blade Tip Clearances

Turbine rotor tips and casings are vulnerable to mechanical failures due to the extreme thermal loads they undergo during engine operation. In addition to the heat flux variations during the transient phase, high-frequency unsteadiness occurs at every rotor passage, with amplitude dependent on the tip gap. The development of appropriate predictive tools and cooling schemes requires the precise understanding of the heat transfer mechanisms. The present paper analyzes the nature of the overtip flow in transonic turbine rotors running at tight clearances, and explores a methodology to determine the relevant flow parameters that model the heat transfer.Steady-state three-dimensional Reynolds-Averaged Navier-Stokes calculations were performed to simulate engine-like conditions considering two rotor tip gaps, 0.1% and 1% of the blade span. At tight tip clearance, the adiabatic wall temperature is not anymore independent of the solid thermal boundary conditions. The adiabatic wall temperature predicted with the linear Newton’s cooling law was observed to rise to non-physical levels in certain regions within the rotor tip gap, resulting in unreliable convective heat transfer coefficients. This paper investigates different approaches to estimate the relevant flow parameters that drive the heat transfer. The present study allows experimentalists to retrieve information on the gap flow temperature and convective heat transfer coefficient based on the use of wall heat flux measurements. Such approach is required to improve the accuracy in the evaluation of the heat transfer data while enhancing the understanding of tight-clearance overtip flows.Copyright

Blade Tip Carving Effects on the Aero-Thermal Performance of a Transonic Turbine

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 on the overtip flow 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 transonic turbine rotor blade (Reynolds number is 5.5 × 105, relative exit Mach number is 0.9) by means of three-dimensional Reynolds-Averaged Navier-Stokes calculations. The novel contoured tip geometry was designed based on a 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.Copyright

Experimental Heat Transfer Investigation in a Multisplitter LP Vane Architecture

This paper reports the external convective heat transfer in an innovative low pressure vane with multisplitter configuration. Three small aerodynamic blades are positioned between each structural vane, providing a novel architecture for ultra-high by-pass ratio aero-engines, with increased LP vane radius and swan-neck diffuser to link the HP turbine. The measurements have been performed in the compression tube test rig of the von Karman Institute, using single layered thin film gauges. Time-averaged and time-resolved heat transfer distributions are presented for the three aerovanes and for the structural blade, at three pressure ratios tested at representative conditions of modern aeroengines, with M2,is ranging from 0.87 to 1.07 and a Reynolds number of about 106 . This facility is specially suited to control the gas-to-wall temperature ratio. Accurate time-averaged heat transfer distributions around the aerovanes are assessed, that allow characterizing the boundary layer status for each position and pressure ratio. The heat transfer distribution around the structural blade is also obtained, depicting clear transition to turbulence, as well as particular flow features on the pressure side, like separation bubbles. Unsteady data analysis reveals the destabilizing effect of the rotor left-running shock on the aerovanes boundary layer, as well as the shift of transition onset for different blade passing events.Copyright