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Research Papers

J. Turbomach. 2018;140(4):041001-041001-12. doi:10.1115/1.4038736.

In this work, the flows inside a high-pressure turbine (HPT) vane and stage are studied with a delayed detached eddy simulation (DDES) code. The fundamental nozzle/blade interaction is investigated with special attention paid to the development and transportation of the vane wake vortices. There are two motivations for this work. First, the extreme HPT operation conditions, including both transonic Mach numbers and high Reynolds numbers, impose a great challenge to modern computational fluid dynamics (CFD), especially for scale-resolved simulation methods. An accurate and efficient high-fidelity CFD solver is very important for a thorough understanding of the flow physics and the design of more efficient HPT. Second, the periodic wake vortex shedding is an important origin of turbine losses and unsteadiness. The wake and vortices not only cause losses themselves, but also interact with the shock wave (under transonic working condition), pressure waves, and have a strong impact on the downstream blade surface (affecting boundary layer transition and heat transfer). Based on one of our previous DDES simulations of a HPT vane, this work further investigates the development and length characteristics of the wake vortices, provides explanations for the length characteristics, and reveals the transportation of the wake vortices in the downstream rotor passages along with its impact on the downstream aero-thermal performance.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041002-041002-8. doi:10.1115/1.4038870.

The mean flow field in a smooth rotating channel was measured by particle image velocimetry (PIV) under the effect of buoyancy force. In the experiments, the Reynolds number, based on the channel hydraulic diameter (D) and the bulk mean velocity (Um), is 10,000, and the rotation numbers are 0, 0.13, 0.26, 0.39, and 0.52, respectively. The four channel walls are heated with indium tin oxide (ITO) heater glass, making the density ratio (d.r.) about 0.1 and the maximum value of buoyancy number up to 0.27. The mean flow field was simulated on a three-dimensional (3D) reconstruction at the position of 3.5 < X/D < 6.5, where X is along the mean flow direction. The effect of Coriolis force and buoyancy force on the mean flow was taken into consideration in the current work. The results show that the Coriolis force pushes the mean flow to the trailing side, making the asymmetry of the mean flow with that in the static conditions. On the leading surface, due to the effect of buoyancy force, the mean flow field changes considerably. Comparing with the case without buoyancy force, separated flow was captured by PIV on the leading side in the case with buoyancy force. More details of the flow field will be presented in this work.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041003-041003-11. doi:10.1115/1.4038839.

Turbomachinery active subspace performance maps are two-dimensional (2D) contour plots that illustrate the variation of key flow performance metrics with different blade designs. While such maps are easy to construct for design parameterizations with two variables, in this paper, maps will be generated for a fan blade with twenty-five design variables. Turbomachinery active subspace performance maps combine active subspaces—a new set of ideas for dimension reduction—with fundamental turbomachinery aerodynamics and design spaces. In this paper, contours of (i) cruise efficiency, (ii) cruise pressure ratio (PR), (iii) maximum climb flow capacity, and (iv) sensitivity to manufacturing variations are plotted as objectives for the fan. These maps are then used to infer pedigree design rules: how best to increase fan efficiency; how best to desensitize blade aerodynamics to the impact of manufacturing variations? In the present study, the former required both a reduction in PR and flow capacity—leading to a reduction of the strength of the leading edge bow wave—while the latter required strictly a reduction in flow capacity. While such pedigree rules can be obtained from first principles, in this paper, these rules are derived from the active subspaces. This facilitates a more detailed quantification of the aerodynamic trade-offs. Thus, instead of simply stating that a particular design is more sensitive to manufacturing variations; or that it lies on a hypothetical “efficiency cliff,” this paper seeks to visualize, quantify, and make precise such notions of turbomachinery design.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041004-041004-10. doi:10.1115/1.4038826.

Problems in turbomachinery computational fluid dynamics (CFD) are often characterized by nonlinear and discontinuous responses. Ensuring the reliability of uncertainty quantification (UQ) codes in such conditions, in an autonomous way, is challenging. In this work, we suggest a new approach that combines three state-of-the-art methods: multivariate Padé approximations, optimal quadrature subsampling (OQS), and statistical learning. Its main component is the generalized least-squares multivariate Padé–Legendre (PL) approximation. PL approximations are globally fitted rational functions that can accurately describe discontinuous nonlinear behavior. They need fewer model evaluations than local or adaptive methods and do not cause the Gibbs phenomenon like continuous polynomial chaos methods. A series of modifications of the Padé algorithm allows us to apply it to arbitrary input points instead of optimal quadrature locations. This property is particularly useful for industrial applications, where a database of CFD runs is already available, but not in optimal parameter locations. One drawback of the PL approximation is that it is nontrivial to ensure reliability. To improve stability, we suggest to couple it with OQS. Our reasoning is that least-squares errors, caused by an ill-conditioned design matrix, are the main source of error. Finally, we use statistical learning methods to check smoothness and convergence. The resulting method is shown to efficiently and correctly fit thousands of partly discontinuous response surfaces for an industrial film cooling and shock interaction problem using only nine CFD simulations.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041005-041005-11. doi:10.1115/1.4038877.

Slot film cooling in an accelerating boundary layer with high freestream turbulence is studied numerically using large eddy simulations (LES). Calculations are done for a symmetrical leading edge geometry with the slot fed by a plenum populated with pin fins. The synthetic eddy method is used to generate different levels of turbulence and length scales at the inflow cross-plane. Calculations are done for a Reynolds number of 250,000 and freestream turbulence levels of 0.7%, 3.5%, 7.8%, and 13.7% to predict both film cooling effectiveness and heat transfer coefficient over the test surface. These conditions correspond to the experimental measurements of (Busche, M. L., Kingery, J. E., and Ames, F. E., 2014, “Slot Film Cooling in an Accelerating Boundary Layer With High Free-Stream Turbulence,” ASME Paper No. GT2014-25360.) Numerical results show good agreement with measurements and show the observed decay of thermal effectiveness and increase of Stanton number with turbulence intensity. Velocity and turbulence exiting the slot are nonuniform laterally due to the presence of pin fins in the plenum feeding the slot which creates a nonuniform surface temperature distribution. No transition to fully turbulent boundary layer is observed throughout the numerical domain. However, freestream turbulence increases wall shear stress downstream driving the velocity profiles toward the turbulent profile and counteracts the laminarizing effects of the favorable pressure gradient. The effective Prandtl number decreases with freestream turbulence. The temperature profiles deviate from the self-similar profile measured under low freestream turbulence condition, reflecting the role of the increased diffusivity in the boundary layer at higher freestream turbulence.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041006-041006-12. doi:10.1115/1.4038689.

The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trends with exit Mach number ranging from 0.8 to 1.4. Assessments using steady Reynolds-averaged Navier--Stokes equations (RANS) computation of the flow in the two turbine blades, complemented with control volume analyses and loss modeling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge (TE) loss, and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus, RANS is an adequate tool for determining the loss variations with exit isentropic Mach number and the loss model serves as an effective tool to interpret both the computational and the experimental data. The measured loss plateau in blade 1 for exit Mach number of 1–1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach number; this plateau is absent in blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For exit Mach number from 0.85 to 1, the higher loss associated with shock system in blade 1 is due to the larger divergent angle downstream of the throat than that in blade 2. However, when exit Mach number is between 1.00 and 1.30, blade 2 has higher shock loss. For exit Mach number above an approximate value of 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing exit Mach number, hence yielding a loss plateau in transonic-supersonic regime.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041007-041007-13. doi:10.1115/1.4038875.

The nonuniformity of the flow field induced by a nonaxisymmetric volute significantly degrades the stability of a turbocharger centrifugal compressor. In this paper, a nonaxisymmetric vaned diffuser is investigated as a nonaxisymmetric flow control method using both three-dimensional computational fluid dynamics (CFD) and experiment. The numerical study first focuses on the relationship between the flow field and the static pressure distortion, and the steady CFD results indicate that the positive static pressure gradient in the rotating direction facilitates flow separation in the vaned diffuser and induces a nonuniform flow field. A nonaxisymmetric flow control method with variable stagger and solidity of the vaned diffuser is developed to suppress the flow separation, and the guideline of the method suggests narrowing flow passages where the flow separates or closing diffuser vanes upstream of flow separations. Steady CFD also presents the flow field of the investigated turbocharger centrifugal compressor with volute, and flow separation is found in the flow passages near the volute tongue. Under the guidance of the nonaxisymmetric flow control method, several nonaxisymmetric vaned diffusers are designed to make the flow field uniform, which are believed to be beneficial for compressor stability. Finally, an experiment is carried out to validate the positive effects of the nonaxisymmetric vaned diffuser for stability improvement. The test data show that Non-AxisVD (with a nonaxisymmetric vaned diffuser) extends the stable flow range (SFR) of the compressor by 26% compared with the AxisVD (with an axisymmetric vaned diffuser), at the cost of acceptable decreases in the maximum total pressure ratio and peak efficiency.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041008-041008-10. doi:10.1115/1.4038833.

An optimal design of film cooling is a key factor in the effort of producing high-efficiency gas turbine. Understanding of the fluid dynamics interaction between cooling holes can help engineers to improve overall thermal effectiveness. Correct prediction through modeling is a very complex problem since multiple phenomena are involved such as mixing, turbulence, and heat transfer. The present work performs an investigation of different cooling configurations ranging from single hole up to two rows. The main objective is to evaluate the double-rows interaction and the effect on film cooling. Strong nonlinear effects are underlined by different simulations, while varying blowing ratio (BR) and geometrical configuration of cooling holes. Meanwhile an initial analysis is performed using flat plate geometry, verification and validation is then extended to realistic stage of high pressure (HP) turbine. Multiple cooling holes configurations are embedded on the pressure side (PS) and suction side (SS) of the single stage. The main outcome is the verification of the thermal effectiveness improvement obtained by cooling jets interaction of multiple rows design. The effects of curvature surface and frame of reference rotation are also evaluated, underlying the differences with standard flat plate test cases.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2018;140(4):041009-041009-11. doi:10.1115/1.4038873.

This paper concentrates on the stall inception analysis of transonic compressors with chordwise and axial sweep. A new prediction approach of stall inception is developed based on global stability analysis and immersed boundary theory, which makes it possible to take both the concrete blade geometry and the complicated base flow into consideration. The prediction of stall inception boils down to an eigenvalue problem. Spectral collocation method is adopted to discretize the eigenvalue equations and the eigenvalues are solved by using singular value decomposition method. The developed prediction approach is validated on two different typical transonic compressors, a single stage compressor and an isolated-rotor compressor, which shows a good agreement with the experimental data. The latter is adopted as a baseline rotor for the investigation of chordwise and axial sweep. By adjusting the stacking line of the baseline rotor, a series of swept rotors are modeled and the stall inception behavior of them is predicted by using the developed approach. The comparison of stall inception behaviors between these rotors is presented, and in combination with steady flow analyses, the effects of sweep features on the stall inception in transonic compressors are discussed.

Commentary by Dr. Valentin Fuster

Errata

J. Turbomach. 2018;140(4):047001-047001-1. doi:10.1115/1.4039142.

Table 1 in “Erratum: ‘Time-Dependent Deposition Characteristics of Fine Coal Fly Ash in a Laboratory Gas Turbine Environment’” contained significant errors in the chemical species labels and some minor errors in the calculated mol % listed under the “Newly measured” column. These errors significantly misrepresented the composition of the ash used in this study. The minor changes in the calculated mol %, however, only had a minimal effect on the calculated critical viscosity and subsequent results and no change on the interpretation and conclusions drawn from the results. The correct composition data are presented below in Table 1.

Topics: Coal , Gas turbines , Fly ash
Commentary by Dr. Valentin Fuster

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