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IN THIS ISSUE

### Research Papers

J. Turbomach. 2017;139(9):091001-091001-10. doi:10.1115/1.4036104.

In turbomachinery design, the accurate prediction of the life cycle is one of the most challenging issues. Traditionally, life cycle calculations for radial turbine wheels of turbochargers focus on mechanical loads such as centrifugal and vibration forces. Due to the increase of exhaust gas temperatures in the last years, thermomechanical fatigue in the turbine wheel came more into focus. In order to account for the thermally induced stresses in the turbine wheel as a part of the standard design process, a fast method is required for predicting metal temperatures. In order to develop a suitable method, the mechanisms that cause the thermal stresses have to be understood. Thus, in a first step, a detailed analysis of the temperature fields is conducted in the present paper. Extensive numerical simulations of a thermal shock process are carried out and validated by experimental data from a test rig. Based on the results, the main heat transfer mechanisms are identified, which are crucial for the critical thermal stresses in transient operation. It is shown that these critical stresses mainly depend on local 3D flow structures. With this understanding, two fast methods to calculate the transient temperatures in a radial turbine were developed. The first method is based on a standard method for transient fluid/solid heat transfer. In this standard method, heat transfer coefficients are derived from steady-state computational fluid dynamics (CFD)/conjugate heat transfer (CHT) calculations and are linearly interpolated over the duration of the transient heating or cooling process. In the new method, this interpolation procedure was modified to achieve an exponential behavior of the heat transfer coefficients over the transient process in order to enable a sufficient accuracy. Additionally, a second method was developed. In this method, the specific heat capacity of the solid state is reduced by a “speed up factor” to shorten the duration of the transient heating or cooling process. With the shortened processes, the computing times can be reduced significantly. After the calculations, the resulting times are transferred into realistic heating or cooling times by multiplying them with the speed up factor. The results of both methods are evaluated against experimental data and against the results of a numerical method known from literature. The methods show a good agreement with those data.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091002-091002-12. doi:10.1115/1.4036106.

Measurements of the mass/heat transfer coefficients on the blade and end wall surfaces of a linear turbine cascade are compared to numerical predictions using the standard shear stress transport (SST) closure and the SST model in combination with the Reθγ transition model (SST-TRANS). Experiments were carried out in a wind tunnel test section composed of five large-scale turbine blades, using the naphthalene sublimation technique. Two cases were tested, with exit Reynolds number of 600,000 and inlet turbulence values of 0.2% and 4%, respectively. The main secondary flow features, consisting of the horseshoe vortex system, the passage vortex, and the corner vortices, are identified and their influence on heat/mass transfer is analyzed. Numerical simulations were carried out to match the conditions of the experiments. Results show that large improvements are obtained with the introduction of the Reθγ transition model. In particular, excellent agreement with the experiments is found, for the whole spanwise extension of the blade, on the pressure surface. On the suction surface, performance is very good in the highly three-dimensional region close to the end wall, but some weaknesses appear in predicting the location of transition in the two-dimensional region. On the end wall surface, the SST model in combination with the transition model produces satisfactory results, greatly improved compared to the standard SST model.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091003-091003-11. doi:10.1115/1.4036190.

Low aspect ratio vanes, often the result of overall engine architecture constraints, create strong secondary flows and high end-wall loss. In this paper, a splitter concept is demonstrated that reduces secondary flow strength and improves stage performance. An analytic conceptual study, corroborated by inviscid computations, shows that the total secondary kinetic energy (SKE) of the secondary flow vortices is reduced when the number of passages is increased and, for a given number of vanes, when the inlet end-wall boundary layer is evenly distributed between the passages. Viscous computations show that, for this to be achieved in a splitter configuration, the pressure-side leg of the low aspect ratio vane horseshoe vortex, must enter the adjacent passage (and not “jump” in front of the splitter leading edge). For a target turbine application, four vane designs were produced using a multi-objective optimization approach. These designs represent current practice for a low aspect ratio vane, a design exempt from thickness constraints, and two designs incorporating splitter vanes. Each geometry is tested experimentally, as a sector, within a low-speed turbine stage. The vane designs with splitter geometries were found to reduce the measured secondary kinetic energy, by up to 85%, to a value similar to the design exempt from thickness constraints. The resulting flow field was also more uniform in both the circumferential and radial directions. One splitter design was selected for a full annulus test where a mixed-out loss reduction, compared to the current practice design, of 15.3% was measured and the stage efficiency increased by 0.88%.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091004-091004-10. doi:10.1115/1.4036107.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091005-091005-11. doi:10.1115/1.4036008.

Ash particle deposition in a high-pressure turbine stage was numerically investigated using steady Reynolds-averaged Navier-Stokes (RANS) and unsteady Reynolds-averaged Navie-Stokes (URANS) methods. An inlet temperature profile consisting of Gaussian nonuniformities (hot streaks) was imposed on the vanes, with vane cooling simulated using a constant vane wall temperature. The steady case utilized a mixing plane at the vane–rotor interface, while a sliding mesh was used for the unsteady case. Corrected speed and mass flow were matched to an experiment involving the same geometry, so that the flow solution could be validated against measurements. Particles ranging from 1 to 65 μm were introduced into the vane domain, and tracked using an Eulerian–Lagrangian tracking model. A novel particle rebound and deposition model was employed to determine particles' stick/bounce behavior upon impact with a surface. Predicted impact and capture distributions for different diameters were compared between the steady and unsteady methods, highlighting effects from the circumferential averaging of the mixing plane. The mixing plane simulation was found to generally under predict impact and capture efficiencies compared with the unsteady calculation, as well as under predict particle temperature upon impact with the blade surface. Quantitative impact and capture efficiency trends with the Stokes number are discussed for both the vane and blade, with companion qualitative distributions for the different Stokes regimes.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091006-091006-8. doi:10.1115/1.4036109.

Cooling flow behavior is investigated within the multiple serpentine passages with turbulators on the leading and trailing walls of an axial gas turbine blade operating at design-corrected conditions with accurate external flow conditions. Pressure and temperature measurements at midspan within the passages are obtained using miniature butt-welded thermocouples and miniature Kulite pressure transducers. These measurements, as well as airfoil surface pressure field data from a full computational fluid dynamics (CFD) simulation, are used as boundary conditions for a model that provides quantitative values of film-cooling blowing ratio for each film-cooling hole on the blade. The model accounts for the continuously changing cross-sectional area and shape of the channels, frictional pressure loss, convective heat transfer from the solid portion of the blade, massflow reduction as coolant bleeds out through film-cooling or impingement holes, compressibility effects, and the effects of blade rotation. The results of the model provide detailed coolant ejection information for a film-cooled rotating turbine airfoil operating at design-corrected conditions and also account for the highly variable freestream conditions on the airfoil. While these values are commonly known for simpler experimental geometries, they have previously either been unknown or estimated crudely for full-stage experiments of this nature. The better-quantified cooling parameters provide a bridge for better comparison with the wealth of film-cooling work already reported for simplified geometries. The calculation also shows the significant range in blowing ratio that can arise even among a single row of cooling holes associated with one of the turbulated passages, due to significant changes in both coolant and local freestream massfluxes.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091007-091007-10. doi:10.1115/1.4036200.

Manufacturing and assembly variation can lead to shifts in the inlet flow incidence angles of a rotating turbine airfoil row. Understanding the sensitivity of the adiabatic film cooling effectiveness to a range of inlet conditions is necessary to verify the robustness of a cooling design. In order to investigate the effects of inlet flow incidence angles, adiabatic and overall effectiveness data were measured in a low speed linear cascade at 0 deg and 10 deg of the designed operating condition. Tests were completed at an inlet Reynolds number of Re = 120,000 and a turbulence intensity of Tu = 5% at the leading edge of the test article. Particle image velocimetry was used to verify the incident flow angle for each angle studied. The test section was first adjusted so that the pressure distribution and stagnation line of the airfoil matched those predicted by an aerodynamic computational fluid dynamics (CFD) model. IR thermography was then used to measure the adiabatic effectiveness levels of the fully cooled airfoil model with nine rows of shaped holes of varying construction and feed delivery. Measurements were taken over a range of blowing ratios and at a density ratio of DR = 1.23. This process was repeated for the two incidence angles measured, while the inlet pressure to the airfoil model was held constant for these incidence angle changes. Differences in laterally adiabatic effectiveness across the airfoil model were most evident in the showerhead, with changes as large as 0.2. The effect persisted most strongly at s/D = ±35 downstream of the stagnation row of holes, but was visible over the whole viewable area of 160 $s/D$. The effect was due to the stagnation line affecting the film at the showerhead row. Due to this effect, the showerhead was investigated in detail, including the effects of the stagnation line shift as well as the influence of the incidence angle on the overall effectiveness of the showerhead region. It was found that the stagnation line has the tendency to dramatically increase the near-hole adiabatic effectiveness levels when positioned within the breakout footprint of the hole. The effect persisted for the overall effectiveness study, since the hole spacing for this particular configuration was wide enough that the through hole convection was not completely dominant. This is the first study to present measured effectiveness values over both the pressure- and suction-side surfaces of a fully cooled airfoil for appreciably off-nominal incidence angles as well as examine adiabatic and overall effectiveness levels for a conical stagnation row of holes.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091008-091008-10. doi:10.1115/1.4036201.

In this paper, experiments and numerical modeling are used to quantify the effects of clearance and eccentricity on compressor performance and to examine the influence of each on flow distribution and stall margin. A change in the size of the tip-clearance gap influences the pressure rise and the stall margin of a compressor. Eccentricity of the tip-clearance gap then further exacerbates the negative effects of increasing tip-clearance. There are few studies in the literature dealing with the combined effect of clearance and eccentricity. There is also little guidance for engine designers, who have traditionally used rules of thumb to quantify these effects. One such rule states that the stall margin of an eccentric machine will be equal to that of a concentric machine with uniform clearance equal to the maximum eccentric clearance. In this paper, this rule of thumb is checked using experimental data and found to be overly pessimistic. In addition, eccentric clearance causes a variation in axial velocity around the circumference of the compressor. The current study uses a three-dimensional model which demonstrates the importance of radial flow gradients in capturing this redistribution. Flow redistribution has been treated analytically in the past, and for this reason, previous modeling has been restricted to two dimensions. The circumferential variation in axial velocity is also examined in terms of the local stability of the flow by considering the stalling flow coefficient of an equivalent axisymmetric compressor with the same local tip-clearance. The large clearance sector of the annulus is found to operate beyond its equivalent axisymmetric stall limit, which means that the small clearance sector of the annulus must be stabilizing the large clearance sector. An improved rule of thumb dealing with the effects of eccentricity is presented.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(9):091009-091009-8. doi:10.1115/1.4036296.

A class of problems in turbomachinery is characterized by unsteady interactions at low reduced frequencies. These interactions are often the result of perturbations with length-scale on the order of the machine circumference and examples include axial compressors operating with inlet distortion, fans with downstream pylons, and turbine rotors downstream of midframe struts. Typically, this unsteadiness is accompanied by higher frequency fluctuations caused by perturbations with a length-scale on the order of a blade pitch. Conventional numerical analysis of this class of problem requires computations with a time step governed by the high-frequency content but a greatly reduced run time could be achieved if the time step was dictated solely by the low reduced frequency, long length-scale, interaction of interest. In this paper, a filtering mixing plane technique is proposed that removes unwanted short length-scale perturbations at the interfaces between blade rows. This approach gives the user control over the amount of mixing that occurs at these interfaces with the limits being fully mixed-out to pitchwise uniformity (conventional mixing plane) or no mixing (conventional sliding plane). By choosing to retain only enough harmonics to resolve the low reduced frequency interaction of interest, an order of magnitude reduction in run time can be achieved.

Commentary by Dr. Valentin Fuster