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Research Papers

J. Turbomach. 2017;139(11):111001-111001-10. doi:10.1115/1.4036764.

This article discusses the development, application, and validation of an optimization method for the impellers of axial fans. The method is supposed to be quick, accurate, and applicable to optimization at an extensive range of design points (DPs). Optimality here means highest possible total-to-static efficiency for a given design point and is obtained by an evolutionary algorithm in which the target function is evaluated by computational fluid dynamics (CFD)-trained artificial neural networks (ANN) of the multilayer perceptron (MLP) type. The MLPs were trained with steady-state CFD (i.e., Reynolds-averaged Navier–Stokes (RANS)) results of approximately 14,000 distinct impellers. After this considerable one-time effort to generate the CFD dataset, each new fan optimization can be performed within a few minutes. It is shown in this article that the MLPs are reliably applicable to all typical design points of axial fans according to Cordier's diagram. Moreover, an extension of the design space toward the classic realm of mixed-flow or even centrifugal fans is observed. It is also shown that the optimization method successfully handles geometrical and operational constraints proving the high degree of universality of the method. Another focus of this article is on the application of the newly developed optimization method to numerous design points. This yields two major findings: the estimation of maximum achievable total-to-static efficiency as a function of the targeted design point (with and without geometrical constraints) as well as a quantification of the improvement over fans designed with classic methods. Both investigations are supported by flow field analyses to aerodynamically explain the findings. Experimental validation of the method was performed with a total of nine prototypes. The positive correlation between MLP, CFD, and experiment successfully validates the methodology.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(11):111002-111002-11. doi:10.1115/1.4037056.

Winglet tips are promising candidates for future high-pressure turbine rotors. Many studies found that the design of the suction-side winglet is the key to the aerodynamic performance of a winglet tip, but there is no general agreement on the exact design philosophy. In this paper, a novel suction-side winglet design philosophy in a turbine cascade is introduced. The winglets are obtained based on the near-tip flow field of the datum tip geometry. The suction-side winglet aims to reduce the tip leakage flow particularly in the front part of the blade passage. It is found that on the casing endwall, the pressure increases in the area where the winglet is used. This reduces the tip leakage flow in the front part of the blade passage and the pitchwise pressure gradient on the endwall. As a result, the size of the tip leakage vortex reduces. A surprising observation is that the novel optimized winglet tip design eliminates the passage vortex and results in a further increasing of the efficiency. The tip leakage loss of the novel winglet tip is 18.1% lower than the datum cavity tip, with an increase of tip surface area by only 19.3%. The spanwise deflection of the winglet due to the centrifugal force is small. The tip heat load of the winglet tip is 17.5% higher than that of the cavity tip. Numerical simulation shows that in a turbine stage, this winglet tip increases the turbine stage efficiency by 0.9% mainly by eliminating the loss caused by the passage vortex at a tip gap size of 1.4% chord compared with a cavity tip.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(11):111003-111003-11. doi:10.1115/1.4037126.

This paper presents an experimental investigation of the concept of using the combustor transition duct wall to shield the nozzle guide vane leading edge. The new vane is tested in a high-speed experimental facility, demonstrating the improved aerodynamic and thermal performance of the shielded vane. The new design is shown to have a lower average total pressure loss than the original vane, and the heat transfer on the vane surface is overall reduced. The peak heat transfer on the vane leading edge–endwall junction is moved further upstream, to a region that can be effectively cooled as shown in previously published numerical studies. Experimental results under engine-representative inlet conditions showed that the better performance of the shielded vane is maintained under a variety of inlet conditions.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(11):111004-111004-17. doi:10.1115/1.4037028.

A laboratory experimental method and an analysis technique are presented for evaluation of individual film-cooling row flow capacity characteristics. The method is particularly suited to complex systems such as hot section nozzle guide vanes (NGV) with lossy feed system characteristics. The method is believed to be both more accurate and more experimentally efficient than previous techniques. The new analysis technique uses an experimentally calibrated network model to represent the complex feed system and replaces the need for internal loss measurements, which are both demanding and inaccurate. Experiments are performed in the purpose-built University of Oxford Coolant Capacity Rig (CCR), a bench-top, blow-down type facility with atmospheric back-pressure. The design of the CCR is informed by the requirements to assess engine-scale film-cooled components rapidly, accurately, and precisely. Improvements in the experimental method include a differential mass flow rate measurement method (which eliminates the effect of leaks and minimizes the number of rows that must be blanked, ensuring that the internal coupling is as close as possible to the engine condition) and a variable bypass flow which ensures the mass flow measurement nozzle always operates within its calibrated range. We demonstrate the method using two high-pressure (HP) NGV designs: an engine part with relatively uncoupled (in terms of internal loss) cooling rows; and a laser-sintered part with highly coupled cooling rows. We show that the individual-row flow capacity of a high-pressure nozzle guide vane (HPNGV) can be evaluated in the CCR in a single day to a 2σ precision of approximately 0.5% and a 2σ accuracy (bias) of 0.6%. The importance of performing individual-row capacity measurements is demonstrated: failure to scale flow capacity on a row-by-row basis introduces an error of 30% in the engine situation.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(11):111005-111005-11. doi:10.1115/1.4037029.

This paper describes the effects of coolant-to-mainstream density ratio and specific heat capacity flux ratio (the product of blowing ratio and specific heat capacity ratio) on the overall cooling effectiveness of high pressure (HP) turbine vanes. Experimental measurements have been conducted at correct engine-matched conditions of Mach number, Reynolds number, turbulence intensity, and coolant-to-mainstream momentum flux ratio. Vanes tested were fully cooled production parts from an engine currently in service. A foreign gas mixture of SF6 and Ar was selected for injection as coolant in the facility so that density and blowing ratios were also matched to the engine situation. The isentropic exponent of the foreign gas mixture coincides with that of air. Full-coverage surface maps of overall cooling effectiveness were acquired by an infrared (IR) thermography technique at a range of mainstream-to-coolant temperature ratios. Measurements were subsequently scaled to engine conditions by employing a new theory based on the principle of superposition and a recovery and redistribution temperature demonstrated in previous papers. It is shown that the two aerodynamically matched situations of air- and foreign-gas-cooled experiments give virtually the same effectiveness trends and patterns. Actual levels differ, however, on account of specific heat capacity flux ratio differences. The effect is described and quantified by a one-dimensional analytical model of the vane wall. Differences in Biot number with respect to engine conditions are discussed as they also influence the scaling of turbine metal temperatures.

Commentary by Dr. Valentin Fuster
J. Turbomach. 2017;139(11):111006-111006-8. doi:10.1115/1.4037127.

This paper will compare two approaches of sensitivity analysis, namely (i) the adjoint method which is used to obtain an initial estimate of the geometric sensitivity of the gas-washed surfaces to aerodynamic quantities of interest and (ii) a Monte Carlo type simulation with an efficient sampling strategy. For both approaches, the geometry is parameterized using a modified NACA parameterization. First, the sensitivity of those parameters is calculated using the linear (first-order) adjoint model. Since the effort of the adjoint computational fluid dynamics (CFD) solution is comparable to that of the initial flow CFD solution and the sensitivity calculation is simply a postprocessing step, this approach yields fast results. However, it relies on a linear model which may not be adequate to describe the relationship between relevant aerodynamic quantities and actual geometric shape variations for the derived amplitudes of shape variations. Second, in order to better capture nonlinear and interaction effects, a Monte Carlo type simulation with an efficient sampling strategy is used to carry out the sensitivity analysis. The sensitivities are expressed by means of the coefficient of importance (CoI), which is calculated based on modified polynomial regression and therefore able to describe relationships of higher order. The methods are applied to a typical high-pressure compressor (HPC) stage. The impact of a variable rotor geometry is calculated by three-dimensional (3D) CFD simulations using a steady Reynolds-averaged Navier–Stokes model. The geometric variability of the rotor is based on the analysis of a set of 400 blades which have been measured using high-precision 3D optical measurement techniques.

Commentary by Dr. Valentin Fuster

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