Accepted Manuscripts

Dr. Beni Cukurel
J. Turbomach   doi: 10.1115/1.4040565
There is a mistake in Eq. 5 which could misguide the reader; the correct formulation is included in the draft file.
TOPICS: Cooling, Turbochargers, Channel flow, Errors, Film cooling
Georgios Ntanakas, Marcus Meyer and Kyriakos Giannakoglou
J. Turbomach   doi: 10.1115/1.4040564
In turbomachinery, the steady adjoint method has been successfully used for the computation of derivatives of various objective functions with respect to design variables in gradient-based optimization. However, the continuous advances in computing power and the accuracy limitations of the steady state assumption lead towards the transition to unsteady CFD computations in the industrial design process. Previous work on unsteady adjoint for turbomachinery applications almost exclusively rely upon frequency-domain methods, for both the flow and adjoint equations. In contrast, in this paper, the development the discrete adjoint to the URANS solver for 3D multi-row applications, in the time-domain, is presented. The adjoint equations are derived along with the adjoint to the 5-stage Runge-Kutta scheme. Communication between adjacent rows is achieved by the adjoint sliding interface method. An optimization workflow that uses unsteady flow and adjoint solvers is presented and tested in two cases, with objective functions accounting for the transient flow in a turbine vane and the periodic flow in a compressor three-row setup.
TOPICS: Turbochargers, Turbomachinery, Shape optimization, Unsteady flow, Flow (Dynamics), Optimization, Computation, Steady state, Turbines, Industrial design, Compressors, Computational fluid dynamics, Design, Workflow, Accounting
Guoping Xia, Gorazd Medic and Thomas Praisner
J. Turbomach   doi: 10.1115/1.4040113
Current design-cycle Reynolds-Averaged Navier-Stokes-based CFD methods have the tendency to over-predict corner-stall events for axial-flow compressors operating at off-design conditions. This shortcoming has been demonstrated even in simple single-row cascade configurations. Here we report on the application of hybrid RANS/LES predictions for simulating the corner-stall data from the linear compressor cascade work conducted at Ecole Centrale de Lyon. This benchmark data set provides detailed loss information while also revealing a bimodal behavior of the separation which, not surprisingly, is also not well modeled by RANS. The hybrid RANS/LES (or DES) results presented here predict bimodal behavior similar to the data only when special treatment is adopted to resolve the leading-edge endwall region where the horseshoe vortex forms. The horseshoe vortex is shown to be unstable, which produces the bimodal instability. The DES simulation without special treatment or refinement in the horseshoe vortex region fails to predict the bimodal instability, and thus the bimodal behavior of the separation. This in turn causes a gross over-prediction in the scale of the corner-stall. The horseshoe vortex region is found to be unstable with rolling of the tertiary vortex over the secondary vortex and merging with the primary horseshoe vortex. With these flow dynamics realized in the DES simulations, the corner stall characteristics are found to be in better agreement with the experimental data, as compared to RANS and standard DES approaches
TOPICS: Compressors, Simulation, Turbochargers, Cascades (Fluid dynamics), Corners (Structural elements), Reynolds-averaged Navier–Stokes equations, Vortices, Separation (Technology), Design, Flow (Dynamics), Axial flow, Cycles, Computational fluid dynamics
Sina C. Stapelfeldt and Mehdi Vahdati
J. Turbomach   doi: 10.1115/1.4040110
This paper examines the factors which can result in discrepancies between rig tests and numerical predictions of the flutter boundary for fan blades. Differences are usually attributed to the deficiency of CFD models for resolving the flow at off-design conditions. This work was initiated as a result of inconsistencies between the flutter prediction of two rig fan blades. A new set of flutter computations for both blades with varying detail in the computational models were performed to investigate these inconsistencies. The results of this work indicate that the previous discrepancies between CFD and tests were due to: 1. Differences in the effectiveness of the acoustic liner in attenuating the pressure wave created by the blade vibration as a result of differences in flutter frequencies between the two fan blades. 2. Differences in the level of unintentional mistuning of the two fan blades due to manufacturing tolerances. In the second part of this research, the effects of blade misstaggering and inlet temperature on aerodynamic damping were investigated. The data presented in this paper clearly show that manufacturing and environmental uncertainties can play an important role in the flutter stability of a fan blade. They demonstrate that aeroelastic similarity is not necessarily achieved if only aerodynamic properties and the traditional aeroelastic parameters, reduced frequency and mass ratio, are maintained. This emphasises the importance of engine-representative models, in addition to an accurate and validated CFD code, for the reliable prediction of the flutter boundary.
TOPICS: Engines, Turbochargers, Flutter (Aerodynamics), Blades, Computational fluid dynamics, Manufacturing, Waves, Pressure, Stability, Flow (Dynamics), Temperature, Acoustics, Damping, Design, Vibration, Computation, Uncertainty
Yanfeng Zhang, Shuzhen Hu, Ali Mahallati, Xue Feng Zhang and Edwards Vlasic
J. Turbomach   doi: 10.1115/1.4039936
The present work, a continuation of a series of investigations on the aerodynamics of aggressive inter-turbine ducts (ITD), is aimed at providing detailed understanding of the flow physics and loss mechanisms in four different ITD geometries. A systematic experimental and computational study was carried out by varying duct outlet-to-inlet area ratios and mean rise angles while keeping the duct length-to-inlet height ratio, Reynolds number and inlet swirl constant in all four geometries. The flow structures within the ITDs were found to be dominated by the boundary layer separation and counter-rotating vortices in both the casing and hub regions. The duct mean rise angle determined the severity of adverse pressure gradient in the casing's first bend whereas the duct area ratio mainly governed the second bend's static pressure rise. The combination of upstream wake flow and the first bend's adverse pressure gradient caused the boundary layer to separate and intensify the strength of counter-rotating vortices. At high mean rise angle, the separation became stronger at the casing's first bend and moved farther upstream. At high area ratios, a two-dimensional separation appeared on the casing and resulted in increased loss. Pressure loss penalties increased significantly with increasing duct mean rise angle and area ratio.
TOPICS: Aerodynamics, Turbochargers, Turbines, Ducts, Separation (Technology), Pressure, Flow (Dynamics), Boundary layers, Pressure gradient, Vortices, Physics, Wakes, Reynolds number
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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