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Research Papers

J. Turbomach. 2016;138(12):121001-121001-9. doi:10.1115/1.4033514.

In the 1950s, NACA conducted a series of low-speed cascade experiments investigating the performance of NACA 65-series compressor cascades with tests covering multiple airfoils of varying camber and with variations in solidity and air inlet angle. Most of the configurations show transition via laminar separation—both on suction and pressure side—characterized by a relatively flat region in pressure distribution, while turbulent reattachment is characterized by a rapid pressure recovery just downstream of the separated region. In the current study, wall-resolved large-eddy simulation (LES) has been used to predict transition via laminar separation in such compressor configurations as well as the resulting airfoil losses. Six different cascades with local diffusion factor varying from 0.14 to 0.56 (NACA 65-010, 65-410, 65-(12)10, 65-(15)10, 65-(18)10, and 65-(21)10 cascades) were analyzed at design conditions. In addition, the loss bucket for various angles of attack off-design conditions has been computed for the NACA 65-(18)10 cascade. Chord-based Reynolds number for all the experiments considered here was held at 250,000. This allows sufficient grid resolution in these LES analyses at an acceptable computational cost, i.e., up to 20,000 CPU hours per case. Detailed comparisons to test data are presented in the form of surface pressure coefficient, drag coefficient, losses, and momentum thickness ratio. The results show that LES is capable of capturing transition via laminar separation relatively well for most of the cases, and consequently, may constitute a predictive tool for assessing losses of different compressor airfoils.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121002-121002-10. doi:10.1115/1.4033481.

Boundary layer suction can be effective in delaying compressor surge, if the surge is triggered by flow separation on the shroud- or hub-casing. This work aims at positioning a suction slot in a radial vaned diffuser, which is thought to be the limiting component in a centrifugal compressor, such as the one considered here. The location of the slot is determined based on the results of both steady and unsteady flow simulations of a transonic centrifugal compressor of a turboshaft. Although the overall performance of the compressor is well-described by steady RANS, large discrepancies are observed between the steady and unsteady simulations of the diffuser flow, discrepancies imply different flow-separation scenarios. Steady results show more low-momentum fluid near the hub, whereas it is concentrated near the shroud in the unsteady simulations, hence no valid physical conclusions can be expected from the steady simulations. Analysis of the instantaneous skin-friction distribution from the unsteady simulations reveals that the separation is fixed and leads to a slot location on the shroud casing, near the diffuser main-vane suction side, so that it covers the range of separation saddle positions as the operating point is changed.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121003-121003-9. doi:10.1115/1.4033473.

In this paper, vane trailing-edge losses which occur in organic rankine cycle (ORC) turbines are investigated. Experiments are performed to study the influence of dense gas effects on trailing-edge loss in supersonic flows using a novel Ludwieg tube facility for the study of dense-gas flows. The data is also used to validate a computational fluid dynamics (CFD) flow solver. The computational simulations are then used to determine the contributions to loss from shocks and viscous effects which occur at the vane trailing edge. The results show that dense gas effects play a vital role in the structure of the trailing-edge flow, and control the extent of shock and viscous losses.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121004-121004-11. doi:10.1115/1.4033506.

While pressure sensitive paint (PSP) technique has been widely used to measure adiabatic film cooling effectiveness distributions on the surfaces of interest based on a mass transfer analog to traditional thermal-based measurements, very little can be found in literature to provide a comprehensive analysis on the uncertainty levels of the measured film cooling effectiveness distributions derived from PSP measurements. In the present study, a detailed analysis is performed to evaluate the effects of various associated uncertainties in the PSP measurements on the measured film cooling effectiveness distributions over the surfaces of interest. The experimental study is conducted in a low-speed wind tunnel under an isothermal condition. While airflow is used to represent the “hot” mainstream flow, an oxygen-free gas, i.e., carbon dioxide (CO2) gas with a density ratio of DR = 1.5 for the present study, is supplied to simulate the “coolant” stream for the PSP measurements to map the adiabatic film cooling effectiveness distribution over a flat test plate with an array of five cylindrical coolant holes at a span-wise spacing of three diameters center-to-center. A comprehensive analysis was carried out with focus on the measurement uncertainty and process uncertainty for the PSP measurements to determine the film cooling effectiveness distributions over the surface of interest. The final analysis indicates that the total uncertainty in the adiabatic film cooling effectiveness measurements by using the PSP technique depends strongly on the local behavior of the mixing process between the mainstream and coolant flows. The measurement uncertainty is estimated as high as 5% at the near field behind the coolant holes. In the far field away from the coolant holes, the total measurement uncertainty is found to be more uniform throughout the measurement domain and generally lower than those in the near field at about 3%.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121005-121005-9. doi:10.1115/1.4033704.

The present paper numerically investigates the stall inception mechanisms in a centrifugal compressor stage composed of a splittered unshrouded impeller and a vaned diffuser. Unsteady numerical simulations have been conducted on a calculation domain comprising all the blade passages over 360 deg for the impeller and the diffuser. Three stable operating points are simulated along a speed line, and the full path to instability is investigated. The paper focusses first on the effects of the mass flow reduction on the flow topology at the inlet of both components. Then, a detailed analysis of stall inception mechanisms is proposed. It is shown that at the inlet of both components, the mass flow reduction induces boundary layer separation on the blade suction side, which results in a vortex tube having its upper end at the casing and its lower end at the blade wall. Some similarities with flows in axial compressor operating at stall condition are outlined. The stall inception process starts with the growth of the amplitude of a modal wave rotating in the vaneless space. As the flow in the compressor is subsonic, the wave propagates upstream and interacts with the impeller flow structure. This interaction leads to the drop in the impeller pressure ratio.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121006-121006-10. doi:10.1115/1.4032928.

Early stage gas turbine blades feature complicated internal geometries in order to enhance internal heat transfer and to supply coolant for film cooling. Most film cooling experiments decouple the effect of internal coolant feed from external film cooling effectiveness, even though engine parts are commonly fed by cross-flow and feature internal rib turbulators which can affect film cooling. Experiments measuring adiabatic effectiveness were conducted to investigate the effects of turbulated perpendicular cross-flow on a row of 45 deg compound angle cylindrical film cooling holes for a total of eight internal rib configurations. The ribs were angled to the direction of prevailing internal cross-flow at two different angles: 45 deg or 135 deg. The ribs were also positioned at two different spanwise locations relative to the cooling holes: in the middle of the cooling hole pitch and slightly intersecting the holes. Experiments were conducted at a density ratio of DR = 1.5 for a range of blowing ratios including M = 0.5, 0.75, 1.0, 1.5, and 2.0. This study demonstrates that peak effectiveness can be attained through the optimization of cross-flow direction relative to the compound angle direction and rib configuration, verifying the importance of hole inlet conditions in film cooling experiments. It was found that ribs tend to reduce adiabatic effectiveness relative to a baseline, smooth-walled configuration. Rib configurations that directed the internal coolant forward in the direction of the mainstream resulted in higher peak adiabatic effectiveness. However, no other parameters could consistently be identified correlating to increased film cooling performance. It is likely that a combination of factors is responsible for influencing performance, including internal local pressure caused by the ribs, the internal channel flow field, in-hole vortices, and jet exit velocity profiles. This study also attempted to replicate the possibility that film cooling holes may intersect ribs and found that a hole which partially intersects a rib still maintains moderate levels of effectiveness.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121007-121007-11. doi:10.1115/1.4033472.

In this study, the impact of single grooves at different locations on compressor stability and tip clearance flow are numerically and experimentally investigated. Initially, the numerical stall margin improvement (SMI) curve is examined using experimental data. Then, the evolution of the interface between the tip leakage flow (TLF) and the incoming main flow (MF) in the prestall and stall inception processes for two typical grooves, i.e., the worst and the optimal grooves in terms of their SMI, are compared with the smooth casing. The results show two different interface behaviors throughout the throttling process. The compressor with the worst single groove casing first experiences a long-length-scale disturbance after the interface near the blade suction side spills in front of the rotor leading-edge plane, and then goes through spikes after the whole interface spills. With the smooth casing and the optimal single groove near midchord, the interface reaches the rotor leading edge at the last stable operating point and spikes appear once the whole interface spills over the rotor leading edge. A model that illustrates the spillage patterns of the interface for the two stall precursors is thus proposed accordingly and used to explain their effectiveness in terms of the SMI. At last, the relevance of these results to the preliminary selection of groove locations for multigroove casing treatments (CTs) is verified by test data and discussed.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121008-121008-16. doi:10.1115/1.4033420.

This paper presents the development of a novel casing treatment to reduce compressor performance and stall margin sensitivities to tip clearance increase. A linked research project on blade design strategies for desensitization had discovered two flow features that reduce sensitivity to tip clearance, namely increased incoming meridional momentum in the rotor tip region and reduction/elimination of double tip leakage flow. Double tip leakage flow is the flow that exits one tip clearance and enters the tip clearance of the circumferentially adjacent blade instead of convecting downstream out of the blade passage. A new and practical casing treatment was developed and analyzed through Reynolds-averaged Navier–Stokes (RANS) computational fluid dynamics (CFD) simulations to decrease double tip leakage and reduce or even eliminate performance and stall margin sensitivity to tip clearance size. The casing treatment design consists of sawtooth-shaped circumferential indentations placed on the shroud over the rotor with a depth on the order of the tip clearance size. A detailed analysis of the flow field allowed for the elucidation of the flow mechanism associated with this casing treatment. A computational parametric study gave preliminary design rules for minimizing both performance/stall margin sensitivity to tip clearance and nominal performance loss. An improved casing indentation design was produced for which CFD simulations showed a complete desensitization of pressure ratio and stall margin while reducing efficiency sensitivity significantly for the tip clearance range studied with only a very small penalty in nominal pressure ratio. Further simulations showed that this casing treatment can be combined with desensitizing blade design strategies to further reduce tip sensitivity and reduce/eliminate/reverse nominal performance penalty. Lastly, preliminary CFD simulations on an axial compressor stage indicate that this shallow indentations' casing treatment strategy remains effective in a stage environment.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121009-121009-12. doi:10.1115/1.4033507.

In the present paper, direct numerical simulation (DNS) data of a low-pressure turbine (LPT) are investigated in light of turbulence modeling. Many compressible turbulence models use Favre-averaged transport equations of the conservative variables and turbulent kinetic energy (TKE) along with other modeling equations. First, a general discussion on the turbulence modeling error propagation prescribed by transport equations is presented, leading to the terms that are considered to be of interest for turbulence model improvement. In order to give turbulence modelers means of validating their models, the terms appearing in the Favre-averaged momentum equations are presented along pitchwise profiles at three axial positions. These three positions have been chosen such that they represent regions with different flow characteristics. General trends indicate that terms related with thermodynamic fluctuations and Favre fluctuations are small and can be neglected for most of the flow field. The largest errors arise close to the trailing edge (TE) region where vortex shedding occurs. Finally, linear models and the scope for their improvement are discussed in terms of a priori testing. Using locally optimized turbulence viscosities, the improvement potential of widely used models is shown. On the other hand, this study also highlights the danger of pure local optimization.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121010-121010-9. doi:10.1115/1.4033973.

Organic Rankine cycle (ORC) turbogenerators are the most viable option to convert sustainable energy sources in the low-to-medium power output range (from tens of kWe to several MWe). The design of efficient ORC turbines is particularly challenging due to their inherent unsteady nature (high expansion ratios and low speed of sound of organic compounds) and to the fact that the expansion encompasses thermodynamic states in the dense vapor region, where the ideal gas assumption does not hold. This work investigates the unsteady nonideal fluid dynamics and performance of a high expansion ratio ORC turbine by means of detailed Reynolds-averaged Navier–Stokes (RANS) simulations. The complex shock interactions resulting from the supersonic flow (M ≈ 2.8 at the vanes exit) are related to the blade loading, which can fluctuate up to 60% of the time-averaged value. A detailed loss analysis shows that shock-induced boundary layer separation on the suction side of the rotor blades is responsible for most of the losses in the rotor, and that further significant contributions are given by the boundary layer in the diverging part of the stator and by trailing edge losses. Efficiency loss due to unsteady interactions is quantified in 1.4% in absolute percentage points at design rotational speed. Thermophysical properties are found to feature large variations due to temperature even after the strong expansion in the nozzle vanes, thus supporting the use of accurate fluid models in the whole turbine stage.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;138(12):121011-121011-12. doi:10.1115/1.4033975.

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.

Commentary by Dr. Valentin Fuster

Technical Brief

J. Turbomach. 2016;138(12):124501-124501-4. doi:10.1115/1.4033672.

Nonintrusive measurement techniques such as particle image velocimetry (PIV) are growing in both capability and utility for turbomachinery applications. However, the restrictive optical access afforded by multistage research compressors typically requires the use of a periscope probe to introduce the laser sheet for measurements in a rotor passage. This paper demonstrates the capability to perform three-dimensional PIV in a multistage compressor without the need for intrusive optical probes and requiring only line-of-sight optical access. The results collected from the embedded second stage of a three-stage axial compressor highlight the rotor tip leakage flow, and PIV measurements are qualitatively compared with high-frequency response piezoresistive pressure measurements to assess the tip leakage flow identification.

Commentary by Dr. Valentin Fuster

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