Research Papers

J. Turbomach. 2017;139(8):081001-081001-10. doi:10.1115/1.4035831.

Experimental measurements of overall cooling effectiveness conducted on a high-pressure turbine vane in a warm rig flow are scaled to engine conditions in this paper. A new theory for the scaling of turbine metal temperatures in cooled compressible flows has been applied, based on the principle of superposition, and demonstrated analytically and numerically in a previous paper. The analysis employs a definition of overall cooling effectiveness based on a new recovery and redistribution temperature, which makes it independent of the temperature boundary conditions of the hot and cold flow streams. This enables the vane external wall temperatures to be scaled to engine conditions by varying, in a fixed aerodynamic field, the mainstream-to-coolant temperature ratio. Experimental validation of the theory is provided in this article. Measurements were conducted in the Annular Sector Heat Transfer Facility, which employs fully cooled nozzle guide vanes, production parts of a civil aviation engine currently in service. Mainstream Mach and Reynolds numbers, inlet turbulence intensity, and coolant-to-mainstream total pressure ratio (and thus momentum flux ratio) are all matched to engine conditions. Full-coverage overall cooling effectiveness distributions, acquired by infrared thermography, are presented for a range of mainstream-to-coolant temperature ratios between 1.05 and 1.22 and subsequently scaled to engine conditions by an iterative procedure. In reducing to practice the principles of the new scaling theory, it is demonstrated that direct validation of turbine cooling system performance is possible in experiments at lower than engine temperatures.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081002-081002-11. doi:10.1115/1.4035833.

This paper describes the use of the free-form-deformation (FFD) parameterization method to create a novel blade shape for a highly loaded, transonic axial compressor. The novel geometry makes use of precompression (via an S-shaping of the blade around midspan) to weaken the shock and improve the aerodynamic performance. It is shown how free-form-deformation offers superior flexibility over traditionally used parameterization methods. The novel design (produced via an efficient optimization method) is presented and the resulting flow is analyzed in detail. The efficiency benefit is over 2%, surpassing other results in the literature for the same geometry. The precompression effect of the S-shape is analyzed and explained, and the entropy increase across the shock (along the midblade line) is shown to be reduced by almost 80%. Adjoint surface sensitivity analysis of the datum and optimized designs is presented, showing that the S-shape is located in the region predicted to be most significant for changes in efficiency. Finally, the off-design performance of the blade is analyzed across the rotor characteristics at various speeds.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081003-081003-13. doi:10.1115/1.4035662.

The major techniques for measuring jet noise have significant drawbacks, especially when including engine installation effects such as jet–flap interaction noise. Numerical methods including low order correlations and Reynolds-averaged Navier–Stokes (RANS) are known to be deficient for complex configurations and even simple jet flows. Using high fidelity numerical methods such as large eddy simulation (LES) allows conditions to be carefully controlled and quantified. LES methods are more practical and affordable than experimental campaigns. The potential to use LES methods to predict noise, identify noise risks, and thus modify designs before an engine or aircraft is built is a possibility in the near future. This is particularly true for applications at lower Reynolds numbers such as jet noise of business jets and jet-flap interaction noise for under-wing engine installations. Hence, we introduce our current approaches to predicting jet noise reliably and contrast the cost of RANS–numerical-LES (RANS–NLES) with traditional methods. Our own predictions and existing literature are used to provide a current guide, encompassing numerical aspects, meshing, and acoustics processing. Other approaches are also briefly considered. We also tackle the crucial issues of how codes can be validated and verified for acoustics and how LES-based methods can be introduced into industry. We consider that hybrid RANS–(N)LES is now of use to industry and contrast costs, indicating the clear advantages of eddy resolving methods.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081004-081004-12. doi:10.1115/1.4035663.

Prior to the detailed design of components, turbomachinery engineers must guide a mean-line or throughflow design toward an optimum configuration. This process requires a combination of informed judgement and low-order correlations for the principle sources of loss. With these requirements in mind, this paper examines the impact of key design parameters on endwall loss in turbines, a problem which remains poorly understood. This paper presents a parametric study of linear cascades, which represent a simplified model of real-engine flow. The designs are nominally representative of the low-pressure turbine blades of an aero-engine, with varying flow angles, blade thickness, and suction surface lift styles. Reynolds-averaged Navier–Stokes (RANS) calculations are performed for a single aspect ratio (AR) and constant inlet boundary layer thickness. To characterize the cascades studied, the two-dimensional design space is examined before studying endwall losses in detail. It is demonstrated that endwall loss can be decomposed into two components: one due to the dissipation associated with the endwall boundary layer and another induced by the secondary flows. This secondary-flow-induced loss is found to scale with a measure of streamwise vorticity predicted by classical secondary flow theory.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081005-081005-12. doi:10.1115/1.4035664.

This paper presents an investigation of the effects of ported shroud (PS) self-recirculating casing treatment used in turbocharger centrifugal compressors for increasing the operable range. The investigation consists of computing three-dimensional flow in a representative centrifugal compressor with and without PS at various levels of approximations in flow physics and geometrical configuration; this provides an enabler for establishing the causal link between PS flow effects and compressor performance changes. It is shown that the main flow path perceives the PS flow as a combination of flow actuations that include injection and removal of mass flow and injection of axial momentum and tangential momentum. A computational model in which the presence of the PS is replaced by imposed boundary conditions (BCs) that reflect the individual flow actuations has thus been formulated and implemented. The removal of a fraction of the inducer mass flow has been determined to be the dominant flow actuation in setting the performance of PS compressors. Mass flow removal reduces the flow blockage associated with the impeller tip leakage flow and increases the diffusion in the main flow path. Adding swirl to the injected flow in the direction opposite to the wheel rotation results in an increase of the stagnation pressure ratio and a decrease of the efficiency. The loss generation in the flow path has been defined to rationalize efficiency changes associated with PS compressor operation.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081006-081006-11. doi:10.1115/1.4035834.

The state-of-the-art design of turbomachinery components is based on Reynolds-averaged Navier–Stokes (RANS) solutions. RANS solvers model the effects of turbulence and boundary layer transition and therefore allow for a rapid prediction of the aerodynamic behavior. The only drawback is that modeling errors are introduced to the solution. Researchers and computational fluid dynamics developers are working on reducing these errors by improved model calibrations which are based on experimental data. These experiments do not typically, however, offer detailed insight into three-dimensional flow fields and the evolution of model quantities in an actual machine. This can be achieved through a direct step-by-step comparison of model quantities between RANS and direct numerical simulation (DNS). In the present work, the experimentally obtained model correlations are recomputed based on DNS of the same turbine profile simulated by RANS. The actual local values are compared to the modeled RANS results, providing information about the source of model deficits. The focus is on the transition process on the blade suction side (SS) and on evaluating the development of turbulent flow structures in the blade's wake. It is shown that the source of disagreement between RANS and DNS can be traced back to three major deficiencies that should be the focus of further model improvements.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081007-081007-12. doi:10.1115/1.4035832.

At the large scale turbine rig (LSTR) at Technische Universität Darmstadt, Darmstadt, Germany, the aerothermal interaction of combustor exit flow conditions on the subsequent turbine stage is examined. The rig resembles a high pressure turbine and is scaled to low Mach numbers. A baseline configuration with an axial inflow and a swirling inflow representative for a lean combustor is modeled by swirl generators, whose clocking position toward the nozzle guide vane (NGV) leading edge can be varied. A staggered double-row of cylindrical film cooling holes on the endwall is examined. The effect of swirling inflow on heat transfer and film cooling effectiveness is studied, while the coolant mass flux rate is varied. Nusselt numbers are calculated using infrared thermography and the auxiliary wall method. Boundary layer, turbulence, and five-hole probe measurements as well as numerical simulations complement the examination. The results for swirling inflow show a decrease of film cooling effectiveness of up to 35% and an increase of Nusselt numbers of 10–20% in comparison to the baseline case for low coolant mass flux rates. For higher coolant injection, the heat transfer is on a similar level as the baseline. The differences vary depending on the clocking position. The turbulence intensity is increased to 30% for swirling inflow.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081008-081008-11. doi:10.1115/1.4035920.

Supercritical CO2 (sCO2) cycles are considered as a promising technology for next generation concentrated solar thermal, waste heat recovery, and nuclear applications. Particularly at small scale, where radial inflow turbines can be employed, using sCO2 results in both system advantages and simplifications of the turbine design, leading to improved performance and cost reductions. This paper aims to provide new insight toward the design of radial turbines for operation with sCO2 in the 100–200 kW range. The quasi-one-dimensional mean-line design code topgen is enhanced to explore and map the radial turbine design space. This mapping process over a state space defined by head and flow coefficients allows the selection of an optimum turbine design, while balancing performance and geometrical constraints. By considering three operating points with varying power levels and rotor speeds, the effect of these on feasible design space and performance is explored. This provides new insight toward the key geometric features and operational constraints that limit the design space as well as scaling effects. Finally, review of the loss break-down of the designs elucidates the importance of the respective loss mechanisms. Similarly, it allows the identification of design directions that lead to improved performance. Overall, this work has shown that turbine design with efficiencies in the range of 78–82% is possible in this power range and provides insight into the design space that allows the selection of optimum designs.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081009-081009-12. doi:10.1115/1.4035921.

A new model is proposed for predicting particle rebound and deposition in environments relevant to gas turbine engines. The model includes the following physical phenomena: elastic deformation, plastic deformation, adhesion, and shear removal. It also incorporates material property sensitivity to temperature and tangential-normal rebound velocity cross-dependencies observed in experiments. The model is well-suited for incorporation in computational fluid dynamics (CFD) simulations of complex gas turbine flows due to its algebraic (explicit) formulation. Model predictions are compared to coefficient of restitution data available in the open literature as well as deposition results from two different high-temperature turbine deposition facilities. While the model comparisons with experiments are in many cases promising, several key aspects of particle deposition remain elusive. The simple phenomenological nature of the model allows for parametric dependencies to be evaluated in a straightforward manner. It is hoped that this feature of the model will aid in identifying and resolving the remaining stubborn holdouts that prevent a universal model for particle deposition.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081010-081010-15. doi:10.1115/1.4036007.

This paper presents an experimental investigation of the rectangular diffusion hole. The effects of rectangular aspect ratio and lateral diffusion angle on film-cooling effectiveness were studied at a low-speed flat-plate experimental facility. The pressure-sensitive paint (PSP) measurement technique was employed to determine the adiabatic effectiveness. The experiments were performed at a density ratio of DR = 1.38 and a mainstream turbulence intensity of Tu = 3.5%. The blowing ratio was varied from M = 0.5 to M = 2.5. Three aspect ratios and three lateral diffusion angles were chosen to match the semicircle and straight-line sidewall shape of the rectangular cross section. A comparative investigation was performed among a typical fan-shaped hole and ten rectangular diffusion holes. The experimental results exhibited diversified film distribution patterns of the rectangular diffusion hole, including single-, bi-, and tripeak patterns. The overall cooling effectiveness increased with the increase of rectangular aspect ratio. The improved magnitude was amplified as blowing ratio increased. The holes with semicircle sidewall were shown to be more suitable for high blowing ratio conditions. The maximum increase of cooling effectiveness was over 70% compared to the fan-shaped hole. The reduction of the lateral diffusion angle affected the film distribution pattern significantly, thereby influencing the cooling effectiveness. To obtain a fixed coverage ratio of film hole row, the rectangular diffusion hole with a larger cross-sectional aspect ratio and a slightly smaller lateral diffusion angle is a preferred scheme.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(8):081011-081011-7. doi:10.1115/1.4035840.

Flutter is a self-excited and self-sustained aero-elastic instability, caused by the positive feedback between structural vibration and aerodynamic forces. A two-passage linear turbine cascade was designed, built, and tested to better understand the phenomena and collect data to validate numerical models. The cascade featured a center airfoil that had its pitch axis as a degree-of-freedom to enable coupling between the air flow and mechanical response in a controlled manner. The airfoil was designed to be excited about its pitch axis using an electromagnetic actuation system over a range of frequencies and amplitudes. The excitation force was measured with load cells, and the airfoil motion was measured with accelerometers. Extraordinary effort was taken to minimize the mechanical damping so that the damping effects of the airflow over the airfoil, that were of primary interest, would be observable. Assembling the cascade required specialized alignment procedures due to the tight clearances and large motion. The aerodynamic damping effects were determined by observing changes in the mechanical frequency response of the system. Detailed aerodynamic and mechanical measurements were conducted within a wide range of Mach numbers (Ma) from Ma = 0.10 to 1.20. Experimental results indicated that the aerodynamic damping increased from Ma = 0.10 to 0.65, dropped suddenly, and was then constant from Ma = 0.80 to 1.20. A flutter condition was identified in the interval between Ma = 0.65 and Ma = 0.80. The aerodynamic damping was also found to be independent of displacement amplitude within the tested range, giving credence to linear numerical approaches.

Commentary by Dr. Valentin Fuster

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