Research Papers

J. Turbomach. 2019;141(3):031001-031001-11. doi:10.1115/1.4042215.

Sand ingestion and deposition in gas turbine engine components can lead to several operational hazards. This paper discusses a physics-based model for modeling the impact, deposition, and sticking of sand particles to surfaces. The collision model includes both normal and tangential components of impact. The normal collision model divides the impact process into three stages, the elastic stage, the elastic–plastic stage, and full plastic stage, and the recovery process is assumed to be fully elastic. The adhesion loss in the recovery stage is described using Timoshenko's model and Tsai's model, and shows that the two models are consistent under certain conditions. Plastic deformation losses of surface asperities are also considered for particle–wall collisions. The normal impact model is supplemented by an impulse-based tangential model, which includes both sliding and rolling frictions. Sand properties are characterized by size and temperature dependencies. The predicted coefficient of restitution (COR) of micron-sized sand particles is in very good agreement with experimental data at room temperature and at higher temperatures from 1073 K to 1340 K. The predicted COR decreases rapidly at temperatures above 1340 K. There is a strong interplay between the size-dependent properties of micron sand particles and the temperature dependency of yield stress on the collision and deposition characteristics. This is the first physics-based high temperature model including translation and rotation of micron-sized sand particles with sliding and rolling modes in the gas turbine literature.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031002-031002-9. doi:10.1115/1.4042202.

Intakes of reduced length have been proposed with the aim of producing aero-engines with higher efficiency and reduced weight. As the intake length decreases, it is expected that stronger effects of the fan on the flow over the intake lip will be seen. If the effects of the fan cannot be ignored, a low-cost but still accurate fan model is of great importance for designing a short-intake. In this paper, a low order rotor/stator model, the immersed boundary method with smeared geometry (IBMSG), has been further developed and validated on a rig test case. The improved IBMSG is more robust than the original. The rig test case used for validation features a low-pressure compression system with a nonaxisymmetric inflow, which is representative of the inlet condition of an aero-engine at its cruise condition. Both the fan and the outlet guide vanes (OGVs) are modeled using IBMSG. A detailed analysis is carried out on the flow both upstream and downstream of the fan. After validating the IBMSG method against the rig test case, a short-intake case, coupled with a fan designed for the next generation of aero-engines, is further investigated. It is found that compared with the intake-alone case, the inflow distortion at the fan face is significantly reduced by the presence of fan. Due to this increased interaction between the fan and the flow over the intake lip, accounting for the effects of the downstream fan is shown to be essential when designing a short intake.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031003-031003-8. doi:10.1115/1.4042066.

Experimental data describing laminar separation bubbles developing under strong adverse pressure gradients, typical of ultra-high-lift turbine blades, have been analyzed to define empirical correlations able to predict the main features of the separated flow transition. Tests have been performed for three different Reynolds numbers and three different free-stream turbulence intensity levels. For each condition, around 4000 particle image velocimetry (PIV) snapshots have been acquired. A wavelet-based intermittency detection technique, able to identify the large scale vortices shed as a consequence of the separation, has been applied to the large amount of data to efficiently compute the intermittency function for the different conditions. The transition onset and end positions, as well as the turbulent spot production rate, are evaluated. Thanks to the recent advancements in the understanding on the role played by Reynolds number and free-stream turbulence intensity on the dynamics leading to transition in separated flows, guest functions are proposed in the paper to fit the data. The proposed functions are able to mimic the effects of Reynolds number and free-stream turbulence intensity level on the receptivity process of the boundary layer in the attached part, on the disturbance exponential growth rate observed in the linear stability region of the separated shear layer, as well as on the nonlinear later stage of completing transition. Once identified the structure of the correlation functions, a fitting process with own and literature data allowed us to calibrate the unknown constants. Results reported in the paper show the ability of the proposed correlations to adequately predict the transition process in the case of separated flows. The correlation for the spot production rate here proposed extends the correlations proposed in literature for attached (by-pass like) transition process, and could be used in γ–Reϑ codes, where the spot production rate appears as a source term in the intermittency function transport equation.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031004-031004-11. doi:10.1115/1.4042310.

This paper describes observed modal oscillations arising from a feedback mechanism between an acoustic resonance in the exit flow channel and aerodynamic and aeroelastic disturbances in a transonic fan stage. During tests, the fan suffered from rotating stall and surge which were preceded by low frequency pressure fluctuations. Through a range of aerodynamic and aeromechanical instrumentations, it was possible to determine a clear chain of cause and effect, whereby geometrical asymmetries trigger local instabilities and modal oscillations through an interaction with the system acoustics. To the authors knowledge, this is the first time that modal oscillations occurring before stall are attributed to multiphysical interactions, showing that acoustic characteristics of the system can influence the aerodynamic as well as the aeromechanical stability of fans. This bears implications for the stability assessment of fans and compressors because first, the stability margin may be affected by standing waves generated in bypass ducts or combustion chambers, and second, geometrical variations of the rotor blades which are believed to be beneficial for aeromechanical stability may lead to complex coupling phenomena.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031005-031005-10. doi:10.1115/1.4042164.

The present paper describes the application of proper orthogonal decomposition (POD) to large eddy simulation (LES) of the T106A low-pressure-turbine profile with unsteady incoming wakes at two different flow conditions. Conventional data analysis applied to time averaged or phase-locked averaged flow fields is not always able to identify and quantify the different sources of losses in the unsteady flow field as they are able to isolate only the deterministic contribution. A newly developed procedure allows such identification of the unsteady loss contribution due to the migration of the incoming wakes, as well as to construct reduced order models that are able to highlight unsteady losses due to larger and/or smaller flow structures carried by the wakes in the different parts of the blade boundary layers. This enables a designer to identify the dominant modes (i.e., phenomena) responsible for loss, the associated generation mechanism, their dynamics, and spatial location. The procedure applied to the two cases shows that losses in the fore part of the blade suction side are basically unaffected by the flow unsteadiness, irrespective of the reduced frequency and the flow coefficient. On the other hand, in the rear part of the suction side, the unsteadiness contributes to losses prevalently due to the finer scale (higher order POD modes) embedded into the bulk of the incoming wake. The main difference between the two cases has been identified by the losses produced in the core flow region, where both the largest scale structures and the finer ones produces turbulence during migration. The decomposition into POD modes allows the quantification of this latter extra losses generated in the core flow region, providing further inputs to the designers for future optimization strategies.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031006-031006-9. doi:10.1115/1.4042165.

This paper presents the results of an experimental study on the transport of entropy waves within a research turbine stage, representative of the key aero-thermal phenomenon featuring the combustor-turbine interaction in aero-engines. The entropy waves are injected upstream of the turbine by a dedicated entropy wave generator (EWG) and are released in axial direction; they feature circular shape with peak amplitude in the center and exhibit sinusoidal-like temporal evolution over the whole wave area. The maximum over-temperature amounts to 7% of the undisturbed flow, while the frequency is 30 Hz. The entropy waves are released in four azimuthal positions upstream of the stage, so to simulate four different burner-to-stator blade clocking. Time-resolved temperature measurements were performed with fast microthermocouples (FTC); the flow and the pressure field upstream and downstream of the stator and the rotor was measured with five-hole pneumatic probes and fast-response aerodynamic pressure probes. The entropy waves are observed to undergo a relevant attenuation throughout their transport within the stator blade row, but they remain clearly visible at the stator exit and retain their dynamic characteristics. In particular, the total temperature distribution appears severely altered by burner-stator clocking position. At the stage exit, the entropy waves loose their coherence, appearing spread in the azimuthal direction to almost cover the entire pitch in the outer part of the channel, while being more localized below midspan. Despite the severe and unsteady interaction of the entropy waves within the rotor, they retain their original dynamic character. A comparison with measurements performed by injecting steady hot streaks is finally reported, remarking both differences and affinities. As a relevant conclusion, it is experimentally shown that entropy waves can be proficiently simulated by considering a succession of hot streaks of different amplitude.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031007-031007-11. doi:10.1115/1.4042070.

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (SH) (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole, while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031008-031008-12. doi:10.1115/1.4042248.

Accurate estimation of slip factor is of paramount importance to ensure centrifugal compressor work input is adequately predicted during the preliminary design process. However, variations in the flow field at impeller exit in both the pitchwise and spanwise directions complicate the evaluation procedure considerably. With the increasing implementation of engine downsizing technologies in the automotive sector, achieving a wide operating range has become a factor of prime importance for centrifugal compressors used in automotive turbocharging applications. As a result of the design features required to achieve this aim, modern impeller geometries have been shown to exhibit an approximately parabolic variation in slip factor across their respective operating maps. By comparison, traditional slip correlations typically exhibit a constant, or at best monotonic, relationship between slip factor and impeller exit flow coefficient. It is this lack of modeling fidelity which the current work seeks to address. In order to tackle these shortcomings, it is proposed that the impeller exit flow should be considered as being made up of three distinct regions: a region of recirculation next to the shroud providing aerodynamic blockage to the stage active flow, and a pitchwise subdivision of the active flow region into jet and wake components. It is illustrated that this hybrid approach in considering both spanwise and pitchwise stratification of the flow permits a better representation of slip factor to be achieved across the operating map. The factors influencing the relative extent of each of these three distinct regions of flow are numerous, requiring detailed investigations to successfully understand their sources and to characterize their extent. A combination of 3D computational fluid dynamics (CFD) data and gas stand test data for six automotive turbocharger compressor stages was employed to achieve this aim. Through application of the extensive interstage static pressure data gathered during gas stand testing at Queen's University Belfast, the results from the 3D CFD models were validated, thus permitting a more in-depth evaluation of the flow field in terms of locations and parameters that could not easily be measured under gas stand test conditions. Building on previous knowledge gained about the variation in shroud side recirculation with geometry and operating condition, the characteristic jet/wake flow structure emanating from the active flow region of the impeller was represented in terms of area and mass flow components. This knowledge allowed individual slip factor values for the jet and wake to be calculated and combined to give an accurate passage average value which exhibited the distinctive nonlinear variation in slip across the operating map which is frequently absent from existing modeling methods. Fundamental considerations of the flow phenomena in each region provided explanation of the results and permitted a modeling approach to be derived to replicate the trends observed in both the experimental data and the CFD simulations.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031009-031009-10. doi:10.1115/1.4042009.

The last 50 years has witnessed significant improvement in film cooling technologies while transpiration cooling is still not implemented in turbine airfoil cooling. Although transpiration cooling could provide higher cooling efficiency with less coolant consumption compared to film cooling, the fine pore structure and high porosity in transpiration cooling metal media always raised difficulties in conventional manufacturing. Recently, the rapid development of additive manufacturing (AM) has provided a new perspective to address such challenge. With the capability of the innovative powder bed selective laser metal sintering (SLMS) AM technology, the complex geometries of transpiration cooling part could be precisely fabricated and endued with improved mechanical strength. This study utilized the SLMS AM technology to fabricate the transpiration cooling and film cooling structures with Inconel 718 superalloy. Five different types of porous media including two perforated plates with different hole pitches, metal sphere packing, metal wire mesh, and blood vessel shaped passages for transpiration cooling were fabricated by EOS M290 system. One laidback fan-shaped film cooling coupon was also fabricated with the same printing process as the control group. Heat transfer tests under three different coolant mass flow rates and four different mainstream temperatures were conducted to evaluate the cooling performance of the printed coupons. The effects of geometry parameters including porosity, surface outlet area ratio, and internal solid–fluid interface area ratio were investigated as well. The results showed that the transpiration cooling structures generally had higher cooling effectiveness than film cooling structure. The overall average cooling effectiveness of blood vessel-shaped transpiration cooling reached 0.35, 0.5, and 0.57, respectively, with low (1.2%), medium (2.4%), and high (3.6%) coolant injection ratios. The morphological parameters analysis showed the major factor that affected the cooling effectiveness most was the internal solid–fluid interface area ratio for transpiration cooling. This study showed that additive manufactured transpiration cooling could be a promising alternative method for turbine blade cooling and worthwhile for further investigations.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031010-031010-9. doi:10.1115/1.4042166.

Internal crossflow, or internal flow that is perpendicular to the overflowing mainstream, reduces film cooling effectiveness by disrupting the diffusion of coolant at the exit of axial shaped holes. Previous experimental investigations have shown that internal crossflow causes the coolant to bias toward one side of the diffuser and that the severity of the biasing scales with the inlet velocity ratio, VRi, or the ratio of crossflow velocity to the jet velocity in the metering section of the hole. It has been hypothesized and computationally predicted that internal crossflow produces an asymmetric swirling flow within the hole that causes the coolant to bias in the diffuser and that biasing contributes to ingestion of hot mainstream gas into the hole, which is undesirable. However, there are no experimental measurements as of yet to confirm these predictions. In the present study, in- and near-hole flow field and thermal field measurements were performed to investigate the flow structures and mainstream ingestion for a standard axial shaped hole fed by internal crossflow. Three different inlet velocity ratios of VRi = 0.24, 0.36, and 0.71 were tested at varying injection rates. Measurements were made in planes normal to the nominal direction of coolant flow at the outlet plane of the hole and at two downstream locations—x/d =0 and 5. The predicted swirling structure was observed for the highest inlet velocity ratio and flow within the hole was shown to scale with VRi. Ingestion within the diffuser was significant and also scaled with VRi. Downstream flow and thermal fields showed that increased biasing contributed to more severe jet detachment and coolant dispersion away from the surface.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031011-031011-8. doi:10.1115/1.4042210.

Film cooling can have a significant effect on the heat transfer coefficient (HTC) between the overflowing freestream gas and the underlying surface. This study investigated the influence of approach flow characteristics, including the boundary layer thickness and character (laminar and turbulent), as well as the approach flow Reynolds number, on the HTC. The figure of merit for this study was the HTC augmentation, that is, the ratio of HTCs for a cooled versus uncooled surface. A heated foil surface provided a known heat flux, allowing direct measurement of HTC and augmentation. The foil was placed both upstream and downstream of the film cooling holes, in order to generate an approaching thermal boundary layer, as representative of actual engine conditions. High-resolution IR thermography provided spatially resolved HTC augmentation data. An open-literature shaped-hole design was used, known as the 7-7-7 hole, in order to compare with existing results in the literature. A variety of blowing conditions were tested from M =0.5 to 3.0. Two elevated density ratios of DR = 1.20 and DR = 1.80 were used. The results indicated that turbulent boundary layer thickness had a modest effect on HTC augmentation, whereas a very high level of augmentation was observed for a laminar approach boundary layer. The presence of upstream heating greatly increased the HTC augmentation in the near-hole region, although these effects died out by 10–15 diameters from the holes.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031012-031012-9. doi:10.1115/1.4041738.

The current industrial standard for numerical simulations of axial compressors is the steady Reynolds-averaged Navier–Stokes (RANS) approach. Besides the well-known limitations of mixing planes, namely their inherent inability to capture the potential interaction and the wakes from the upstream blades, there is another flow feature which is lost, and which is a major accountable for the radial mixing: the transport of streamwise vorticity. Streamwise vorticity is generated for various reasons, mainly associated with secondary and tip-clearance flows. A strong link exists between the strain field associated with the vortices and the mixing augmentation: the strain field increases both the area available for mixing and the local gradients in fluid properties, which provide the driving potential for the mixing. In the rear compressor stages, due to high clearances and low aspect ratios, only accounting for the development of secondary and clearance flow structures, it is possible to properly predict the spanwise mixing. In this work, the results of steady and unsteady simulations on a heavy-duty axial compressor are compared with experimental data. Adopting an unsteady framework, the enhanced mixing in the rear stages is properly captured, in remarkable agreement with experimental distributions. On the contrary, steady analyses strongly underestimate the radial transport. It is inferred that the streamwise vorticity associated with clearance flows is a major driver of radial mixing, and restraining it by pitch-averaging the flow at mixing planes is the reason why the steady approach cannot predict the radial transport in the rear part of the compressor.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031013-031013-10. doi:10.1115/1.4041673.

Most studies of turbine airfoil film cooling in laboratory test facilities have used relatively large plenums to feed flow into the coolant holes. However, a more realistic inlet condition for the film cooling holes is a relatively small channel. Previous studies have shown that the film cooling performance is significantly degraded when fed by perpendicular internal crossflow in a smooth channel. In this study, angled rib turbulators were installed in two geometric configurations inside the internal crossflow channel, at 45 deg and 135 deg, to assess the impact on film cooling effectiveness. Film cooling hole inlets were positioned in both prerib and postrib locations to test the effect of hole inlet position on film cooling performance. A test was performed independently varying channel velocity ratio and jet to mainstream velocity ratio. These results were compared to the film cooling performance of previously measured shaped holes fed by a smooth internal channel. The film cooling hole discharge coefficients and channel friction factors were also measured for both rib configurations with varying channel and inlet velocity ratios. Spatially averaged film cooling effectiveness is largely similar to the holes fed by the smooth internal crossflow channel, but hole-to-hole variation due to inlet position was observed.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031014-031014-8. doi:10.1115/1.4041700.

This paper describes the development and initial application of an adjoint harmonic balance (HB) solver. The HB method is a numerical method formulated in the frequency domain which is particularly suitable for the simulation of periodic unsteady flow phenomena in turbomachinery. Successful applications of this method include unsteady aerodynamics as well as aeroacoustics and aeroelasticity. Here, we focus on forced response due to the interaction of neighboring blade rows. In the simulation-based design and optimization of turbomachinery components, it is often helpful to be able to compute not only the objective values—e.g., performance data of a component—themselves but also their sensitivities with respect to variations of the geometry. An efficient way to compute such sensitivities for a large number of geometric changes is the application of the adjoint method. While this is frequently used in the context of steady computational fluid dynamics (CFD), it becomes prohibitively expensive for unsteady simulations in the time domain. For unsteady methods in the frequency domain, the use of adjoint solvers is feasible but still challenging. The present approach employs the reverse mode of algorithmic differentiation (AD) to construct a discrete adjoint of an existing HB solver in the framework of an industrially applied CFD code. The paper discusses implementational issues as well as the performance of the adjoint solver, in particular regarding memory requirements. The presented method is applied to compute the sensitivities of aeroelastic objectives with respect to geometric changes in a turbine stage.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2019;141(3):031015-031015-12. doi:10.1115/1.4041699.

A compressor blade integrated with circumferential groove casing treatment (CGCT) is optimized in this study. A hybrid aerodynamic optimization algorithm that combines the differential evolution (DE) with a radial basis function (RBF) response surface is used for the multi-objective optimization via the computational fluid dynamics (CFD) analysis. The sweep and lean distributions are optimized to pursue the maximum total pressure ratio and adiabatic efficiency at the design point. Constraints on the choking mass flow rate and the near-stall compression ratio are imposed to ensure the off-design performance. The performance is improved much more with the blade-CGCT integrated optimization than with the blade-only optimization. The stall margin of the blade-only optimized blade with CGCT added as an afterthought can be even worse than the baseline blade. The CGCT-removal test for the blade-CGCT integrated optimization result further verifies that the superior performance of the blade-CGCT integrated optimization is obtained via optimizing the coupling between the effects of the sweep and lean on the blade loading and the effects of the CGCT on the flow blockage.

Commentary by Dr. Valentin Fuster

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