Research Papers

J. Turbomach. 2018;140(12):121001-121001-8. doi:10.1115/1.4041061.

Gas turbine design has been characterized over the years by a continuous increase of the maximum cycle temperature, justified by a corresponding increase of cycle efficiency and power output. In such way, turbine components heat load management has become a compulsory activity, and then, a reliable procedure to evaluate the blades and vanes metal temperatures is, nowadays, a crucial aspect for a safe components design. In the framework of the design and validation process of high pressure turbine cooled components of the BHGE NovaLTTM 16 gas turbine, a decoupled methodology for conjugate heat transfer prediction has been applied and validated against measurement data. The procedure consists of a conjugate heat transfer analysis in which the internal cooling system (for both airfoils and platforms) is modeled by an in-house one-dimensional thermo-fluid network solver, the external heat loads and pressure distribution are evaluated through 3D computational fluid dynamics (CFD) analysis and the heat conduction in the solid is carried out through a 3D finite element method (FEM) solution. Film cooling effect has been treated by means of a dedicated CFD analysis, implementing a source term approach. Predicted metal temperatures are finally compared with measurements from an extensive test campaign of the engine in order to validate the presented procedure.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121002-121002-18. doi:10.1115/1.4041142.

Stall followed by surge in a high speed compressor can lead to violent disruption of flow, damage to the blade structures and, eventually, engine shutdown. Knowledge of unsteady blade loading during such events is crucial in determining the aeroelastic stability of blade structures; experimental test of such events is, however, significantly limited by the potential risk and cost associated. Numerical modeling, such as unsteady computational fluid dynamics (CFD) simulations, can provide a more informative understanding of the flow field and blade forcing during poststall events; however, very limited publications, particularly concerning multistage high speed compressors, can be found. The aim of this paper is to demonstrate the possibility of using CFD for modeling full-span rotating stall and surge in a multistage high speed compressor, and, where possible, validate the results against experimental measurements. The paper presents an investigation into the onset and transient behavior of rotating stall and surge in an eight-stage high speed axial compressor at off-design conditions, based on 3D Reynolds-averaged Navier–Stokes (URANS) computations, with the ultimate future goal being aeroelastic modeling of blade forcing and response during such events. By assembling the compressor with a small and a large exit plenum volume, respectively, a full-span rotating stall and a deep surge were modeled. Transient flow solutions obtained from numerical simulations showed trends matching with experimental measurements. Some insights are gained as to the onset, propagation, and merging of stall cells during the development of compressor stall and surge. It is shown that surge is initiated as a result of an increase in the size of the rotating stall disturbance, which grows circumferentially to occupy the full circumference resulting in an axisymmetric flow reversal.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121003-121003-9. doi:10.1115/1.4041290.

The effects of heat transfer between the compressor structure and the primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Nonadiabatic transient calculations show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation, with the remainder attributed to stage rematching. Heat transfer increases loading in the front stages and destabilizes the front block. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects only capture 1.2 points of the 8.0 point stall margin reduction. Based on this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121004-121004-10. doi:10.1115/1.4041375.

Windmilling requirements for aircraft engines often define propulsion and airframe design parameters. The present study is focused on two key quantities of interest during windmill operation: fan rotational speed and stage losses. A model for the rotor exit flow is developed that serves to bring out a similarity parameter for the fan rotational speed. Furthermore, the model shows that the spanwise flow profiles are independent of the throughflow, being determined solely by the configuration geometry. Interrogation of previous numerical simulations verifies the self-similar nature of the flow. The analysis also demonstrates that the vane inlet dynamic pressure is the appropriate scale for the stagnation pressure loss across the rotor and splitter. Examination of the simulation results for the stator reveals that the flow blockage resulting from the severely negative incidence that occurs at windmill remains constant across a wide range of mass flow rates. For a given throughflow rate, the velocity scale is then shown to be that associated with the unblocked vane exit area, leading naturally to the definition of a dynamic pressure scale for the stator stagnation pressure loss. The proposed scaling procedures for the component losses are applied to the flow configuration of Prasad and Lord (2010). Comparison of simulation results for the rotor-splitter and stator losses determined using these procedures indicates very good agreement. Analogous to the loss scaling, a procedure based on the fan speed similarity parameter is developed to determine the windmill rotational speed and is also found to be in good agreement with engine data. Thus, despite their simplicity, the methods developed here possess sufficient fidelity to be employed in design prediction models for aircraft propulsion systems.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121005-121005-10. doi:10.1115/1.4041233.

An increasingly common experimental method allows determination of the overall effectiveness of a film cooled turbine component. This method requires the Biot number of the experimental model to match that of the engine component such that the nondimensional surface temperature, ϕ, is matched to that of the engine component. The matched Biot number requirement effectively places a requirement on the thermal conductivity of the model and the traditional implementation places no requirement on the model's density or specific heat. However, such is not the case if such a model is exposed to unsteadiness in the flow such as with film cooling unsteadiness. In this paper, we develop an additional nondimensional parameter that must also be theoretically matched to conduct overall effectiveness experiments with unsteady film cooling. Since finding suitable materials with an acceptable combination of thermodynamic properties for a typical low temperature experiment can be difficult, simulations were conducted to determine the impact of imperfectly matched parameters achievable with common materials. Because the disparity between the diffusion and the unsteadiness time scales can hinder numerical simulation, a novel analytical solution to the heat equation with relevant unsteady Robin type boundary conditions is developed. Particular solutions are examined to determine the sensitivity of the temperature response of a turbine blade (or a model of one) to its material properties and the form of the unsteady variation in the convection parameters. It is shown that it is possible to obtain useful experimental results even with imperfectly matched parameters.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121006-121006-11. doi:10.1115/1.4041373.

A simple nondimensional model to describe the flutter onset of labyrinth seals is presented. The linearized mass and momentum integral equations for a control volume which represents the interfin seal cavity, retaining the circumferential unsteady flow perturbations created by the seal vibration, are used. First, the downstream fin is assumed to be choked, whereas in a second step the model is generalized for unchoked exit conditions. An analytical expression for the nondimensional work-per-cycle is derived. It is concluded that the stability of a two-fin seal depends on three nondimensional parameters, which allow explaining seal flutter behavior in a comprehensive fashion. These parameters account for the effect of the pressure ratio, the cavity geometry, the fin clearance, the nodal diameter (ND), the fluid swirl velocity, the vibration frequency, and the torsion center location in a compact and interrelated form. A number of conclusions have been drawn by means of a thorough examination of the work-per-cycle expression, also known as the stability parameter by other authors. It was found that the physics of the problem strongly depends on the nondimensional acoustic frequency. When the discharge time of the seal cavity is much greater than the acoustic propagation time, the damping of the system is very small and the amplitude of the response at the resonance conditions is very high. The model not only provides a unified framework for the stability criteria derived by Ehrich (1968, “Aeroelastic Instability in Labyrinth Seals,” ASME J. Eng. Gas Turbines Power, 90(4), pp. 369–374) and Abbot (1981, “Advances in Labyrinth Seal Aeroelastic Instability Prediction and Prevention,” ASME J. Eng. Gas Turbines Power, 103(2), pp. 308–312), but delivers an explicit expression for the work-per-cycle of a two-fin rotating seal. All the existing and well-established engineering trends are contained in the model, despite its simplicity. Finally, the effect of swirl in the fluid is included. It is found that the swirl of the fluid in the interfin cavity gives rise to a correction of the resonance frequency and shifts the stability region. The nondimensionalization of the governing equations is an essential part of the method and it groups physical effects in a very compact form. Part I of the paper details the derivation of the theoretical model and draws some preliminary conclusions. Part II of the corresponding paper analyzes in depth the implications of the model and outlines the extension to multiple cavity seals.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121007-121007-8. doi:10.1115/1.4041377.

The dimensionless model presented in part I of the corresponding paper to describe the flutter onset of two-fin rotating seals is exploited to extract valuable engineering trends with the design parameters. The analytical expression for the nondimensional work-per-cycle depends on three dimensionless parameters of which two of them are new. These parameters are simple but interrelate the effect of the pressure ratio, the height, and length of the interfin geometry, the seal clearance, the nodal diameter (ND), the fluid swirl velocity, the vibration frequency, and the torsion center location in a compact and intricate manner. It is shown that nonrelated physical parameters can actually have an equivalent impact on seal stability. It is concluded that the pressure ratio can be stabilizing or destabilizing depending on the case, whereas the swirl of the flow is always destabilizing. Finally, a simple method to extend the model to multiple interfin cavities, neglecting the unsteady interaction among them, is described.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121008-121008-14. doi:10.1115/1.4041294.

This paper presents a detailed experimental and numerical study on the effects of upstream step geometry on the endwall secondary flow and heat transfer in a transonic linear turbine vane passage with axisymmetric converging endwalls. The upstream step geometry represents the misalignment between the combustor exit and the nozzle guide vane endwall. The experimental measurements were performed in a blowdown wind tunnel with an exit Mach number of 0.85 and an exit Re of 1.5×106. A high freestream turbulence level of 16% was set at the inlet, which represents the typical turbulence conditions in a gas turbine engine. Two upstream step geometries were tested for the same vane profile: a baseline configuration with a gap located 0.88Cx (43.8 mm) upstream of the vane leading edge (upstream step height = 0 mm) and a misaligned configuration with a backward-facing step located just before the gap at 0.88Cx (43.8 mm) upstream of the vane leading edge (step height = 4.45% span). The endwall temperature history was measured using transient infrared thermography, from which the endwall thermal load distribution, namely, Nusselt number, was derived. This paper also presents a comparison with computational fluid dynamics (CFD) predictions performed by solving the steady-state Reynolds-averaged Navier–Stokes with Reynolds stress model using the commercial CFD solver ansysfluent v.15. The CFD simulations were conducted at a range of different upstream step geometries: three forward-facing (upstream step geometries with step heights from −5.25% to 0% span), and five backward-facing, upstream step geometries (step heights from 0% to 6.56% span). These CFD results were used to highlight the link between heat transfer patterns and the secondary flow structures and explain the effects of upstream step geometry. Experimental and numerical results indicate that the backward-facing upstream step geometry will significantly enlarge the high thermal load region and result in an obvious increase (up to 140%) in the heat transfer coefficient (HTC) level, especially for arched regions around the vane leading edge. However, the forward-facing upstream geometry will modestly shrink the high thermal load region and reduce the HTC (by ∼10% to 40% decrease), especially for the suction side regions near the vane leading edge. The aerodynamic loss appears to have a slight increase (0.3–1.3%) because of the forward-facing upstream step geometry but is slightly reduced (by 0.1–0.3%) by the presence of the backward upstream step geometry.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121009-121009-12. doi:10.1115/1.4041601.

The main centrifugal compressor performance criteria are pressure ratio, efficiency, and wide flow range. The relative importance of these criteria, and therefore the optimum design balance, varies between different applications. Vaned diffusers are generally used for high-performance applications as they can achieve higher efficiencies and pressure ratios, but have a reduced operating range, in comparison to vaneless diffusers. Many impeller-based casing treatments have been developed to enlarge the operating range of centrifugal compressors over the last decades but there is much less information available in open literature for diffuser focused methods, and they are not widely adopted in commercial compressor stages. The development of aerodynamic instabilities at low mass flow rate operating conditions can lead to the onset of rotating stall or surge, limiting the stable operating range of the centrifugal compressor stage. More understanding of these aerodynamic instabilities has been established in recent years. Based on this additional knowledge, new casing treatments can be developed to prevent or suppress the development of these instabilities, thus increasing the compressor stability at low mass flow rates. This paper presents a novel vaned diffuser casing treatment that successfully increased the stable operating range at low mass flow rates and high pressure ratios. Detailed experimental measurements from a high pressure ratio turbocharger compressor stage combined with complementary CFD simulations were used to examine the effect of the new diffuser casing treatment on the compressor flow field and led to the improvement in overall compressor stability. A detailed description of how the new casing treatment operates is presented within the paper.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121010-121010-7. doi:10.1115/1.4041811.

Determination of a scalable Nusselt number (based on “adiabatic heat transfer coefficient”) has been the primary objective of the most existing heat transfer experimental studies. Based on the assumption that the wall thermal boundary conditions do not affect the flow field, the thermal measurements were mostly carried out at near adiabatic condition without matching the engine realistic wall-to-gas temperature ratio (TR). Recent numerical studies raised a question on the validity of this conventional practice in some applications, especially for turbine blade. Due to the relatively low thermal inertia of the over-tip-leakage (OTL) flow within the thin clearance, the fluids' transport properties vary greatly with different wall thermal boundary conditions and the two-way coupling between OTL aerodynamics and heat transfer cannot be neglected. The issue could become more severe when the gas turbine manufacturers are making effort to achieve much tighter clearance. However, there has been no experimental evidence to back up these numerical findings. In this study, transient thermal measurements were conducted in a high-temperature linear cascade rig for a range of tip clearance ratio (G/S) (0.3%, 0.4%, 0.6%, and 1%). Surface temperature history was captured by infrared thermography at a range of wall-to-gas TRs. Heat transfer coefficient (HTC) distributions were obtained based on a conventional data processing technique. The profound influence of tip surface thermal boundary condition on heat transfer and OTL flow was revealed by the first-of-its-kind experimental data obtained in the present experimental study.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(12):121011-121011-8. doi:10.1115/1.4041809.

The interaction of flow and film-cooling effectiveness between jets of double-jet film-cooling (DJFC) holes on a flat plate is studied experimentally. The time-averaged flow field in several axial positions (X/d = −2.0, 1.0, and 5.0) is obtained through a seven-hole probe. The downstream film-cooling effectiveness on the flat plate is measured by pressure sensitive paint (PSP). The inclination angle (θ) of all the holes is 35 deg, and the compound angle (β) is ±45 deg. Effects of the spanwise distance (p = 0, 0.5d, 1.0d, 1.5d, and 2.0d) between the two interacting jets of DJFC holes are studied, while the streamwise distance (s) is kept as 3d. The blowing ratio (M) varies as 0.5, 1.0, 1.5, and 2.0. The density ratio (DR) is maintained at 1.0. Results show that the interaction between the two jets of DJFC holes has different effects at different spanwise distances. For a small spanwise distance (p/d = 0), the interaction between the jets presents a pressing effect. The downstream jet is pressed down and kept attached to the surface by the upstream one. The effectiveness is not sensitive to blowing ratios. For mid-spanwise distances (p/d = 0.5 and 1.0), the antikidney vortex pair dominates the interaction and pushes both of the jets down, thus leading to better coolant coverage and higher effectiveness. As the spanwise distance becomes larger (p/d ≥ 1.5), the pressing effect almost disappears, and the antikidney vortex pair effect is weaker. The jets separate from each other and the coolant coverage decreases. At a higher blowing ratio, the interaction between the jets of DJFC holes happens later.

Commentary by Dr. Valentin Fuster

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