Accepted Manuscripts

Li He and Junsok Yi
J. Turbomach   doi: 10.1115/1.4036765
A general issue in turbomachinery flow computations is how to capture and resolve two kinds of unsteadiness efficiently and accurately: a) deterministic disturbances with temporal and spatial periodicities linked to blade count and rotational speed, and b) non-deterministic disturbances including turbulence and self-excited coherent patterns (e.g. vortex shedding, shear layer instabilities, etc.) with temporal and spatial wave lengths unrelated to blade count and rotational speed. In particular, the high cost of large eddy simulations (LES) is further compounded by the need to capture the deterministic unsteadiness of bladerow interactions in computational domains with large number of blade passages. The present work addresses this challenge by developing a multi-scale solution approach. The framework is based on an ensemble-averaging to split deterministic and nondeterministic disturbances. The two types of disturbances can be solved in suitably selected computational domains and solvers respectively. The local fine mesh is used for non-deterministic turbulence eddies and vortex shedding, while the global coarse mesh is for deterministic unsteadiness. A key enabler is that the unsteady stress terms of the non-deterministic disturbances are obtained only in a small set of blade passages and propagated to the whole domain with many more passages by a block spectral mapping. This distinctive multi-scale treatment makes it possible to achieve a high resolution URANS/LES solution in a multi-passage/whole annulus domain very efficiently. The method description will be followed by test cases demonstrating the validity and potential of the proposed methodology.
TOPICS: Flow (Dynamics), Turbochargers, Turbomachinery, Blades, Vortex shedding, Turbulence, Eddies (Fluid dynamics), Stress, Waves, Resolution (Optics), Shear (Mechanics), Annulus, Large eddy simulation, Computation
Konrad Bamberger and Thomas Carolus
J. Turbomach   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. 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 CFD-trained artificial neural networks (ANN). The ANNs were trained with steady state CFD results of approximately 14,000 distinct impellers. After this considerable onetime effort, each new fan optimization can be performed within a few minutes. It is shown that the ANNs are reliably applicable to all typical design points of axial fans according to Cordier's diagram. Moreover, an extension of the design space towards the classic realm of mixed-flow or even centrifugal fans is observed. It is also shown that the optimization method successfully handles geometrical constraints. 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 as well as a quantification of the improvement over fans designed with classic methods. Both investigations are supported by flow filed 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.
TOPICS: Turbochargers, Fans, Optimization, Design, Artificial neural networks, Computational fluid dynamics, Flow (Dynamics), Impellers, Engineering prototypes, Evolutionary algorithms, Steady state
Tianyu Pan, Qiushi Li, Zhiping Li and Yifang Gong
J. Turbomach   doi: 10.1115/1.4036646
Partial surge is a new type of instability inception in the form of axisymmetric low-frequency disturbance located in the hub region and has been observed in transonic axial flow compressors. Previous studies on the evolution of instability in a transonic axial flow compressor at different rotor speeds found that partial surge occurs and leads to full compressor flow instability at high rotor speeds but not at low rotor speeds, and the blade loading at the hub increases with the rotor speed. A hypothesis is first made that the level of blade loading in the hub region could be highly correlated to the occurrence of partial surge. Experiments and numerical simulation are then conducted to test this hypothesis when the radial distribution of blade loading near the stall point is varied by introducing inlet distortion (i.e. alternately mounting specially designed screens at the inlet of the compressor). Both the experimental results of instability evolution and the numerical results of radial distribution of blade loading show that high hub loading near the stall point is the necessary condition for the occurrence of partial surge. In addition, the general effects of radial loading distribution on the type of stall inception are presented and discussed.
TOPICS: Turbochargers, Axial flow, Surges, Rotors, Blades, Compressors, Computer simulation, Flow instability, Stall inception
Zhang Min, Liu Yan, Zhang Tianlong, Zhang Mengchao and Ying He
J. Turbomach   doi: 10.1115/1.4036647
This paper presents a continued study on a previously investigated winglet-shroud (WS) geometry for a linear turbine cascade. A plain tip and a full shroud tip are experimentally and numerically examined as the datum cases. Various width (w) of double-side winglets (DSW) involving 3%, 5%, 7% and 9% of the blade pitch (p) is numerically investigated. It is observed that the DSW cases do not alter the flow fields including the separation bubble and reattachment flow within the tip gap region, even for the case with the broadest width (w/p = 9%). Meanwhile, the horse-shoe vortex near the casing is not generated even for the case with the smallest width (w/p=3%). Larger width of the DSW geometry is indeed able to improve the aerodynamic performance, but in a slight degree. With the w/p increasing from 3% to 9%, the mass-averaged total pressure loss coefficient over an exit plane is just reduced by 2.61%. Therefore, a favorable width of w/p=5% is chosen to design the WS structure. Based on this, three locations of the partial shroud (linkage segment) are devised, which are located near the leading edge, the middle and close to the trailing edge respectively. Results illustrate that all three WS cases have advantages in lessening the aerodynamic loss over the DSW arrangement, but with the linkage segment located in the middle having optimal effect. This conclusion verifies the feasibility of the previously studied WS configuration.
TOPICS: Turbochargers, Cascades (Fluid dynamics), Optimization, Turbines, Geometry, Linkages, Flow (Dynamics), Separation (Technology), Bubbles, Vortices, Blades, Design, Pressure
Technical Brief  
Santosh Patil, Ivana D. Atanasovska and Saravanan Karuppanan
J. Turbomach   doi: 10.1115/1.4030242
The aim of this paper is to provide a new viewpoint of friction factor for contact stress calculations of gears. The idea of friction factor has been coined, for the calculation of contact stresses along the tooth contact for different helical gear pairs. Friction factors were developed by evaluating contact stresses with and without friction for different gear pairs. In this paper, 3D Finite Element Method (FEM) and Lagrange Multiplier algorithm has been used to evaluate the contact stresses. Initially, a spur gear FE model was validated with the theoretical analysis under frictionless condition, which is based on Hertz's contact theory. Then, similar FE models were constructed for 5, 15, 25 and 35 deg. helical gear pairs. The contact stresses of these models were evaluated for different coefficients of friction. These results were employed for the development of friction factor.

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