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Research Papers

J. Turbomach. 2016;139(3):031001-031001-10. doi:10.1115/1.4034800.

High-fidelity simulations, e.g., large eddy simulation (LES), are often needed for accurately predicting pressure losses due to wake mixing and boundary layer development in turbomachinery applications. An unsteady adjoint of high-fidelity simulations is useful for design optimization in such aerodynamic applications. In this paper, we present unsteady adjoint solutions using a large eddy simulation model for an inlet guide vane from von Karman Institute (VKI) using aerothermal objectives. The unsteady adjoint method is effective in capturing the gradient for a short time interval aerothermal objective, whereas the method provides diverging gradients for long time-averaged thermal objectives. As the boundary layer on the suction side near the trailing edge of the vane is turbulent, it poses a challenge for the adjoint solver. The chaotic dynamics cause the adjoint solution to diverge exponentially from the trailing edge region when solved backward in time. This results in the corruption of the sensitivities obtained from the adjoint solutions. An energy analysis of the unsteady compressible Navier–Stokes adjoint equations indicates that adding artificial viscosity to the adjoint equations can dissipate the adjoint energy while potentially maintaining the accuracy of the adjoint sensitivities. Analyzing the growth term of the adjoint energy provides a metric for identifying the regions in the flow where the adjoint term is diverging. Results for the vane obtained from simulations performed on the Titan supercomputer are demonstrated.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031002-031002-9. doi:10.1115/1.4034816.

In this paper, the modifications induced by the presence of an inlet flow nonuniformity on the aerodynamic performance of a nozzle vane cascade are experimentally assessed. Tests were carried out in a six vane linear cascade whose profile is typical of a first stage nozzle guide vane of a modern heavy-duty gas turbine. An obstruction was located in the wind tunnel inlet section to produce a nonuniform flow upstream of the leading edge plane. The cascade was tested in an atmospheric wind tunnel at an inlet Mach number Ma1 = 0.12, with a high turbulence intensity (Tu1 = 9%) and variable obstruction tangential and axial positions, as well as tangential extension. The presented results show that an inlet flow nonuniformity influences the stagnation point position when it faces the vane leading edge from the suction side. A relevant increase of both 2D and secondary losses is observed when the nonuniformity is aligned to the vane leading edge. When it is instead located in between the passage, it does not affect the stagnation point location, in the meanwhile allowing a reduction in the secondary loss.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031003-031003-11. doi:10.1115/1.4034683.

The effect of coolant addition or “mixing loss” on aerodynamic performance is formulated for the turbine, where mixing takes place between gas streams of different compositions as well as static temperatures. To do this, a second-law efficiency measure is applied to a generalization of the one-dimensional mixing problem between a main gas stream and a single coolant feed, first introduced and studied by Hartsel (1972, “Prediction of Effects of Mass-Transfer Cooling on the Blade-Row Efficiency of Turbine Airfoils,” AIAA Paper No. 1972-11) for the turbine application. Hartsel's 1972 model for mass transfer cooling loss still remains the standard for estimating mixing loss in today's turbines. The present generalization includes losses due to the additional contributions of “compositional mixing” (mixing between unlike compositions of the main and coolant streams) as well as the effect of chemical reaction between the two streams. Scaling of the present dissipation function-based loss model to the mainstream Mach number and relative cooling massflow and static temperature is given. Limitations of the constant specific heats assumptions and the impact of fuel-to-air ratio are also quantified.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031004-031004-9. doi:10.1115/1.4034847.

While much is known about how macrogeometry of shaped holes affects their ability to successfully cool gas turbine components, little is known about the influence of surface roughness on cooling hole interior walls. For this study, a baseline-shaped hole was tested with various configurations of in-hole roughness. Adiabatic effectiveness measurements at blowing ratios up to 3 showed that the in-hole roughness caused decreased adiabatic effectiveness relative to smooth holes. Decreases in area-averaged effectiveness grew more severe with larger roughness size and with higher blowing ratios for a given roughness. Decreases of more than 60% were measured at a blowing ratio of 3 for the largest roughness values. Thermal field and flowfield measurements showed that in-hole roughness caused increased velocity of core flow through the hole, which increased the jet penetration height and turbulence intensity resulting in an increased mixing between the coolant and the mainstream. Effectiveness reductions due to roughness were also observed when roughness was isolated to only the diffused outlet of holes, and when the mainstream was highly turbulent.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031005-031005-10. doi:10.1115/1.4034799.

Increasing turbine inlet temperature is one of the main strategies used to accomplish the demand for increased performance of modern gas turbines. Thus, optimization of the cooling system is becoming of paramount importance in gas turbine development. Leading edge (LE) represents a critical part of cooled nozzles and blades, given the presence of the hot gases stagnation point, and the unfavorable geometrical characteristics for cooling purposes. This paper reports the results of a numerical investigation, carried out to support a parallel experimental campaign, aimed at assessing the rotation effects on the internal heat transfer coefficient (HTC) distribution in a realistic LE cooling system of a high pressure blade. Experiments were performed in static and rotating conditions replicating a typical range of jet Reynolds number (10,000–40,000) and Rotation number (0–0.05). The experimental results consist of flowfield measurements on several internal planes and HTC distributions on the LE internal surface. Hybrid RANS–large eddy simulation (LES) models were exploited for the simulations, such as scale adaptive simulation and detached eddy simulation, given their ability to resolve the complex flowfield associated with jet impingement. Numerical flowfield results are reported in terms of both jet velocity profiles and 2D vector plots on two internal planes, while the HTC distributions are presented as detailed 2D maps together with averaged Nusselt number profiles. A fairly good agreement with experiments is observed, which represents a validation of the adopted modeling strategy, allowing an in-depth interpretation of the experimental results.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031006-031006-8. doi:10.1115/1.4034917.

A time-resolved particle image velocimetry (TR-PIV) system has been employed to investigate a laminar separation bubble which is induced by a strong adverse pressure gradient typical of ultrahigh-lift low-pressure turbine (LPT) blades. Proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) are described and applied within this paper. These techniques allow reducing the degrees-of-freedom of complex systems producing a low-order model ranked by the energy content (POD) or by the modal contribution to the dynamics of the system itself (DMD), useful to highlight the dominant dynamics. The time–space evolution of the laminar separation bubble is characterized by rollup vortices shed in the surrounding of the bubble maximum displacement as a consequence of the Kelvin–Helmholtz (KH) instability process as well as by a low-frequency motion of the separated shear layer. The decomposition techniques proposed allow the identification of these coherent structures and the characterization of their modal properties (e.g., temporal frequency, spatial wavelength, and growth rate). The POD separates the different dynamics that induce velocity fluctuations at different frequencies and wavelength looking at their contribution to the overall kinetic energy. The DMD provides complementary information: the unstable spatial frequencies are identified with their growth (or decay) rates. DMD modes associated with the Kelvin–Helmholtz instability and the corresponding vortex shedding phenomenon clearly dominate the unsteady behavior of the laminar separation bubble, being characterized by the highest growth rate. Modes with longer wavelength describe the low-frequency motion of the laminar separation bubble and are neutrally stable. Results reported in this paper prove the ability of the present methods in extracting the dominant dynamics from a large dataset, providing robust and rapid tools for the in depth analysis of transition and separation processes.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031007-031007-14. doi:10.1115/1.4034848.

In this paper, the detailed steady and unsteady numerical investigations of a 1.5-stage axial flow turbine are conducted to determine the specific influence of interplatform steps in the first stator—as caused by deviations in manufacturing or assembly. A basic first stator design and a design consisting of a bow and endwall contours are compared. Apart from step height, the position and geometry of the interplatform border are varied for the basic design. To create the steps, every third stator vane was elevated, together with its platforms at hub and shroud, such that the flow capacity is only little affected. The results show that the effects of steps on the platform borders in front and aft of the first stator can be decoupled from those occurring on the interplatform steps. For the latter, being the main contributor to the additional loss, the intensity of recirculation zones and losses increase substantially when the platform border is located close to the suction side. Using a relative step height of 1.82% span, the entropy production doubles when compared to a position close to the pressure side, which can be explained by differences in local flow velocity level. Regarding a circular-arc-shaped platform, the losses can be more than halved—mainly due to lower included angles between step and endwall flow streamlines. The findings can be explained by a nondimensional relation of the local entropy production using local values for step height and characteristic flow quantities. Furthermore, a reduction in step height leads to an attenuation of the otherwise linear relationship between step height and entropy production, which is mainly due to lower local ratio of step height and boundary layer thickness. In the case of laminar or transitional flow regions on the endwall, typical for turbine rigs with low inlet turbulence and low-pressure turbines under cruise conditions, the steps lead to immediate local flow transition and thus substantially different results.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031008-031008-13. doi:10.1115/1.4034939.

Internal cooling passages of turbine blades have long been at risk to blockage through the deposition of sand and dust during fleet service life. The ingestion of high volumes of volcanic ash (VA) therefore poses a real risk to engine operability. An additional difficulty is that the cooling system is frequently impossible to inspect in order to assess the level of deposition. This paper reports results from experiments carried out at typical high pressure (HP) turbine blade metal temperatures (1163 K to 1293 K) and coolant inlet temperatures (800 K to 900 K) in engine scale models of a turbine cooling passage with film-cooling offtakes. Volcanic ash samples from the 2010 Eyjafjallajökull eruption were used for the majority of the experiments conducted. A further ash sample from the Chaiten eruption allowed the effect of changing ash chemical composition to be investigated. The experimental rig allows the metered delivery of volcanic ash through the coolant system at the start of a test. The key metric indicating blockage is the flow parameter (FP), which can be determined over a range of pressure ratios (1.01–1.06) before and after each experiment, with visual inspection used to determine the deposition location. Results from the experiments have determined the threshold metal temperature at which blockage occurs for the ash samples available, and characterize the reduction of flow parameter with changing particle size distribution, blade metal temperature, ash sample composition, film-cooling hole configuration and pressure ratio across the holes. There is qualitative evidence that hole geometry can be manipulated to decrease the likelihood of blockage. A discrete phase computational fluid dynamics (CFD) model implemented in Fluent has allowed the trajectory of the ash particles within the coolant passages to be modeled, and these results are used to help explain the behavior observed.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031009-031009-8. doi:10.1115/1.4034976.

The effect of the unsteady aerodynamic loading of oscillating airfoils in the low-reduced frequency regime on the work per cycle curves is investigated. The theoretical analysis is based on a perturbation analysis of the linearized Navier–Stokes equations for real modes at low-reduced frequency. It was discovered that a new parameter, the unsteady loading, plays an essential role in the trends of the phase and modulus of the unsteady pressure caused by the airfoil oscillation. Here, the theory is extended in order to quantify this new parameter. It is shown that this parameter depends solely on the steady flow-field on the airfoil surface and the vibration mode-shape. As a consequence, the effect of changing the design operating conditions or the vibration mode onto the work-per-cycle curves (and therefore in the stability) can be easily predicted and, what is more important, quantified without conducting additional flutter analysis. The relevance of the parameter has been numerically confirmed in the Part II of the paper.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031010-031010-8. doi:10.1115/1.4034978.

This paper numerically investigates the correlation between the so-called unsteady loading parameter (ULP), derived in Part I of the corresponding paper, and the unsteady aerodynamics of oscillating airfoils at low reduced frequency with special emphasis on the work-per-cycle curves. Simulations using a frequency-domain linearized Navier–Stokes solver have been carried out on rows of a low-pressure turbine airfoil section, the NACA65 section, and a flat plate, to show the correlation between the actual value of the ULP and the flutter characteristics, for different airfoils, operating conditions, and mode shapes. Both the traveling wave and influence coefficient formulations of the problem are used in combination to increase the understanding of the ULP influence in different aspects of the unsteady flow field. It is concluded that, for a blade vibrating in a prescribed motion at design conditions, the ULP can quantitatively predict the effect of unsteady loading variations due to changes in both the incidence and the mode shape on the work-per-cycle curves. It is also proved that the unsteady loading parameter can be used to qualitatively compare the flutter characteristics of different airfoils.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031011-031011-10. doi:10.1115/1.4034984.

Unshrouded centrifugal compressor impellers typically operate at high rotational speeds and volume flow rates. The resulting high mean stress levels leave little margin for dynamic excitations that can cause high-cycle fatigue. In addition to the well-established high-frequency impeller blade excitations of centrifugal compressors caused by the stationary parts, such as vaned diffusers or inlet guide vanes (IGVs), the presented study addresses an unsteady rotating flow feature (rotating stall) which should be taken into account when addressing the high-cycle fatigue during the design phase. The unsteady fluid–structure interaction between rotating stall and unshrouded impellers was experimentally described and quantified during two different measurement campaigns with two full-size compression units operating under real conditions. In both campaigns, dynamic strain gauges and pressure transducers were mounted at various locations on the impeller of the first compression stage. The casing was also equipped with a set of dynamic pressure transducers to complement the study. Rotating pressure fluctuations were found to form an additional impeller excitation at a frequency that is not a multiple of the shaft speed. The measurements show that the excitation amplitude and frequency caused by the rotating pressure fluctuations depend on the operating conditions and are therefore challenging to predict and consider during the design phase. Furthermore, the excitation mechanism presented was found to cause resonant impeller blade response under specific operating conditions. For the experimentally investigated impeller geometries, a rotating pressure fluctuation caused approximately 1.5 MPa of additional dynamic stress in the structure per 1 mbar of dynamic pressure amplitude when exciting the first bending mode of the impeller. The induced dynamic mechanical stresses due to rotating stall are in the order of 10% of the endurance limit of the material for the tested impeller geometries; therefore, they are not critical and confirm a robust and reliable design.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031012-031012-10. doi:10.1115/1.4034985.

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure-sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5% and 0.75% to 1% for the purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2016;139(3):031013-031013-9. doi:10.1115/1.4034982.

Aircraft engines ingest airborne particulate matter, such as sand, dirt, and volcanic ash, into their core. The ingested particulate is transported by the secondary flow circuits via compressor bleeds to the high pressure turbine and may deposit resulting in turbine fouling and loss of cooling effectiveness. Prior publications focused on particulate deposition and sand erosion patterns in a single stage of a compressor or turbine. This work addresses the migration of ingested particulate through the high pressure compressor (HPC) and bleed systems. This paper describes a 3D CFD methodology for tracking particles along a multistage axial compressor and presents particulate ingestion analysis for a high pressure compressor section. The commercial CFD multiphase solver ANSYS CFX® has been used for flow and particulate simulations. Particle diameters of 20, 40, and 60 μm are analyzed. Particle trajectories and radial particulate profiles are compared for these particle diameters. The analysis demonstrates how the compressor centrifuges the particles radially toward the compressor case as they travel through the compressor; the larger diameter particles being more significantly affected. Nonspherical particles experience more drag as compared to spherical particles, and hence a qualitative comparison between spherical and nonspherical particles is shown.

Commentary by Dr. Valentin Fuster

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