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J. Turbomach. 2018;141(1):011001-011001-10. doi:10.1115/1.4041133.

Effects of axial casing grooves (ACGs) on the stall margin and efficiency of a one and a half stage low-speed axial compressor with a large rotor tip gap are investigated in detail. The primary focus of the current paper is to identify the flow mechanisms behind the changes in stall margin and on the efficiency of the compressor stage with a large rotor tip gap. Semicircular axial grooves installed in the rotor's leading edge area are investigated. A large eddy simulation (LES) is applied to calculate the unsteady flow field in a compressor stage with ACGs. The calculated flow fields are first validated with previously reported flow visualizations and stereo particle image velocimetry (SPIV) measurements. An in-depth examination of the calculated flow field indicates that the primary mechanism of the ACG is the prevention of full tip leakage vortex (TLV) formation when the rotor blade passes under the axial grooves periodically. The TLV is formed when the incoming main flow boundary layer collides with the tip clearance flow boundary layer coming from the opposite direction near the casing and rolls up around the rotor tip vortex. When the rotor passes directly under the axial groove, the tip clearance flow boundary layer on the casing moves into the ACGs and no roll-up of the incoming main flow boundary layer can occur. Consequently, the full TLV is not formed periodically as the rotor passes under the open casing of the axial grooves. Axial grooves prevent the formation of the full TLV. This periodic prevention of the full TLV generation is the main mechanism explaining how the ACGs extend the compressor stall margin by reducing the total blockage near the rotor tip area. Flows coming out from the front of the grooves affect the overall performance as it increases the flow incidence near the leading edge and the blade loading with the current ACGs. The primary flow mechanism of the ACGs is periodic prevention of the full TLV formation. Lower efficiency and reduced pressure rise at higher flow rates for the current casing groove configuration are due to additional mixing between the main passage flow and the flow from the grooves. At higher flow rates, blockage generation due to this additional mixing is larger than any removal of the flow blockage by the grooves. Furthermore, stronger double-leakage tip clearance flow is generated with this additional mixing with the ACGs at a higher flow rate than that of the smooth wall.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;141(1):011002-011002-10. doi:10.1115/1.4041379.

A double-wall cooling scheme combined with effusion cooling offers a practical approximation to transpiration cooling which in turn presents the potential for very high cooling effectiveness. The use of the conventional conjugate computational fluid dynamics (CFD) for the double-wall blade can be computationally expensive and this approach is therefore less than ideal in cases where only the preliminary results are required. This paper presents a computationally efficient numerical approach for analyzing a double-wall effusion cooled gas turbine blade. An existing correlation from the literature was modified and used to represent the two-dimensional distribution of film cooling effectiveness. The internal heat transfer coefficient was calculated from a validated conjugate analysis of a wall element representing an element of the aerofoil wall and the conduction through the blade solved using a finite element code in ANSYS. The numerical procedure developed has permitted a rapid evaluation of the critical parameters including film cooling effectiveness, blade temperature distribution (and hence metal effectiveness), as well as coolant mass flow consumption. Good agreement was found between the results from this study and that from literature. This paper shows that a straightforward numerical approach that combines an existing correlation for film cooling from the literature with a conjugate analysis of a small wall element can be used to quickly predict the blade temperature distribution and other crucial blade performance parameters.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;141(1):011003-011003-11. doi:10.1115/1.4041602.

This paper collects the final results of a combined experimental and numerical investigation on pressure side (PS) film cooling in a high-pressure turbine vane, including two staggered rows of cylindrical holes and a trailing edge cutback, fed by one plenum. Having learned that the scale resolving simulation technique is essential to get reasonable predictions of adiabatic film cooling effectiveness, the stress-blended eddy simulation (SBES) model has been selected as the best among the available hybrid RANS–LES options. Mainstream conditions were limited to low speed and low turbulence intensity due to the need of high temporal and spatial resolution. The choice of one only coolant-to-mainstream mass flow ratio equal to MFR = 1.5% was dictated by the hole discharge: on the one side, mainstream injection into the cooling holes and, on the other side, jet liftoff were avoided to get an effective thermal coverage downstream of the holes. SBES potential was evaluated on the basis of qualitative and quantitative characteristics of the flow along the interface between coolant and mainstream because of their ultimate effect on vane surface temperature. The focus was set on shape and dynamics of coherent structures: SBES provided evidence of shear layer Kelvin–Helmholtz instability and hairpin vortices, downstream of cooling holes, with a Strouhal number (St) of 1.3 and 0.3–0.4, respectively. Simulated vortex shedding in the cutback region was characterized by St of 0.32 to be compared against the measured St value of 0.40.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;141(1):011004-011004-10. doi:10.1115/1.4041291.

Current turbulent heat flux models fail to predict accurate temperature distributions in film cooling flows. The present paper focuses on a machine learning (ML) approach to this problem, in which the gradient diffusion hypothesis (GDH) is used in conjunction with a data-driven prediction for the turbulent diffusivity field αt. An overview of the model is presented, followed by validation against two film cooling datasets. Despite insufficiencies, the model shows some improvement in the near-injection region. The present work also attempts to interpret the complex ML decision process, by analyzing the model features and determining their importance. These results show that the model is heavily reliant of distance to the wall d and eddy viscosity νt, while other features display localized prominence.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;141(1):011005-011005-11. doi:10.1115/1.4041467.

Recent advances in experimental methods have allowed researchers to study nozzle guide vane (NGV) film cooling in the presence of combustor dilution ports and endwall films. The dilution injection creates nonuniformities in temperature, velocity, and turbulence, and an understanding of the vane film cooling performance is complicated by competing influences. In this study, dilution port temperature profiles have been measured in the absence of vane film cooling and compared to film effectiveness measurements in the presence of both films and dilution, illustrating the effects of the dilution port turbulence on film cooling performance. It is found that dilution port injection can create significant effectiveness benefits at the difficult-to-cool vane stagnation region due to the more turbulent hot mainstream enhancing the mixing of film coolant jets that have left the airfoil surface. Also explored are the implications of endwall film cooling for infrared (IR) vane surface temperature measurements. The reduced endwall temperatures reduce the thermal emissions from this surface, so reducing the amount of extraneous radiation reflected from the vane surface where measurements are being made. The results of a detailed calibration show that the maximum local film effectiveness measurement error could be up to 0.05 if this effect were to go unaccounted for.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;141(1):011006-011006-13. doi:10.1115/1.4041466.

The leakage flows within the gap between the tips of unshrouded rotor blades and the stationary casing of high-speed turbines are the source of significant aerodynamic losses and thermal stresses. In the pursuit for higher component performance and reliability, shaping the tip geometry offers a considerable potential to modulate the rotor tip flows and to weaken the heat transfer onto the blade and casing. Nevertheless, a critical shortage of combined experimental and numerical studies addressing the flow and loss generation mechanisms of advanced tip profiles persists in the open literature. A comprehensive study is presented in this two-part paper that investigates the influence of blade tip geometry on the aerothermodynamics of a high-speed turbine. An experimental and numerical campaign has been performed on a high-pressure turbine stage adopting three different blade tip profiles. The aerothermal performance of two optimized tip geometries (one with a full three-dimensional contoured shape and the other featuring a multicavity squealer-like tip) is compared against that of a regular squealer geometry. In the second part of this paper, we report a detailed analysis on the aerodynamics of the turbine as a function of the blade tip geometry. Reynolds-averaged Navier-Stokes (RANS) simulations, adopting the Spalart–Allmaras turbulence model and experimental boundary conditions, were run on high-density unstructured meshes using the numecafine/open solver. The simulations were validated against time-averaged and time-resolved experimental data collected in an instrumented turbine stage specifically setup for the simultaneous testing of multiple blade tips at scaled engine-representative conditions. The tip flow physics is explored to explain variations in turbine performance as a function of the tip geometry. Denton's mixing loss model is applied to the predicted tip gap aerodynamic field to identify and quantify the loss reduction mechanisms of the alternative tip designs. An advanced method based on the local triple decomposition of relative motion is used to track the location, size and intensity of the vortical flow structures arising from the interaction between the tip leakage flow and the main gas path. Ultimately, the comparison between the unconventional tip profiles and the baseline squealer tip highlights distinct aerodynamic features in the associated gap flow field. The flow analysis provides guidelines for the designer to assess the impact of specific tip design strategies on the turbine aerodynamics and rotor heat transfer.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;141(1):011007-011007-14. doi:10.1115/1.4041293.

Experiments in a refractive index-matched axial turbomachine facility show that semicircular skewed axial casing grooves (ACGs) reduce the stall flowrate by 40% but cause a 2.4% decrease in the maximum efficiency. Aiming to elucidate mechanism that might cause the reduced efficiency, stereo-PIV measurements examine the impact of the ACGs on the flow structure and turbulence in the tip region near the best efficiency point (BEP), and compare them to those occurring without grooves and at low flowrates. Results show that the periodic inflow into the groove peaks when the rotor blade pressure side (PS) overlaps with the downstream end of the groove, but diminishes when this end faces the suction side (SS). Entrainment of the PS boundary layer and its vorticity generates a vortical loop at the entrance to the groove, and a “discontinuity” in the tip leakage vortex (TLV) trajectory. During exposure to the SS, the backward tip leakage flow separates at the entrance to the groove, generating a counter-rotating circumferential “corner vortex,” which the TLV entrains into the passage at high flowrates. Interactions among these structures enlarge the TLV and create a broad area with secondary flows and elevated turbulence near the groove's downstream corner. A growing shear layer with weaker turbulence also originates from the upstream corner. The groove also increases the flow angle upstream of the blade tip and varies it periodically. Accordingly, the circulation shed from the blade tip and strength of leakage flow increase near the blade leading edge (LE).

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;141(1):011008-011008-13. doi:10.1115/1.4041559.

In modern lean-burn aero-engine combustors, highly swirling flow structures are adopted to control the fuel-air mixing and to provide the correct flame stabilization mechanisms. Aggressive swirl fields and high turbulence intensities are hence expected in the combustor-turbine interface. Moreover, to maximize the engine cycle efficiency, an accurate design of the high-pressure nozzle cooling system must be pursued: in a film-cooled nozzle, the air taken from last compressor stages is ejected through discrete holes drilled on vane surfaces to provide a cold layer between hot gases and turbine components. In this context, the interactions between the swirling combustor outflow and the vane film cooling flows play a major role in the definition of a well-performing cooling scheme, demanding for experimental campaigns at representative flow conditions. An annular three-sector combustor simulator with fully cooled high-pressure vanes has been designed and installed at THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion-cooled liners, and six film-cooled high-pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central airfoil aligned with the central swirler. In this experimental work, adiabatic film effectiveness measurements have been carried out in the central sector vanes, in order to characterize the film-cooling performance under swirling inflow conditions. The pressure-sensitive paint (PSP) technique, based on heat and mass transfer analogy, has been exploited to catch highly detailed 2D distributions. Carbon dioxide has been used as coolant in order to reach a coolant-to-mainstream density ratio of 1.5. Turbulence and five-hole probe measurements at inlet/outlet of the cascade have been carried out as well, in order to highlight the characteristics of the flow field passing through the cascade and to provide precise boundary conditions. Results have shown a relevant effect of the swirling mainflow on the film cooling behavior. Differences have been found between the central airfoil and the adjacent ones, both in terms of leading edge stagnation point position and of pressure and suction side film coverage characteristics.

Commentary by Dr. Valentin Fuster

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