scholarly journals Building blocks for a leading edge high-order flow solver

PAMM ◽  
2017 ◽  
Vol 17 (1) ◽  
pp. 129-132 ◽  
Author(s):  
Immo Huismann ◽  
Jörg Stiller ◽  
Jochen Fröhlich
2022 ◽  
Author(s):  
Keith Obenschain ◽  
Yu Yu Khine ◽  
Robert Rosenberg ◽  
Raghunandan Mathur ◽  
Gopal Patnaik ◽  
...  
Keyword(s):  

2013 ◽  
Vol 17 (2) ◽  
pp. 255-270 ◽  
Author(s):  
Wei Cao ◽  
Chuan-fu Xu ◽  
Zheng-hua Wang ◽  
Lu Yao ◽  
Hua-yong Liu

Author(s):  
Feng Wang ◽  
Luca di Mare

Abstract Turbomachinery blade rows can have non-uniform geometries due to design intent, manufacture errors or wear. When predictions are sought for the effect of such non-uniformities, it is generally the case that whole assembly calculations are needed. A spectral method is used in this paper to approximate the flow fields of the whole assembly but with significantly less computation cost. The method projects the flow perturbations due to the geometry non-uniformity in an assembly in Fourier space, and only one passage is required to compute the flow perturbations corresponding to a certain wave-number of geometry variation. The performance of this method on transonic blade rows is demonstrated on a modern fan assembly. Low engine order and high engine order geometry non-uniformity (e.g. “saw-tooth” pattern) are examined. The non-linear coupling between the flow perturbations and the passage-averaged flow field is also demonstrated. Pressure variations on the blade surface and the potential flow field upstream of the leading edge from the proposed spectral method and the direct whole assembly solutions are compared. Good agreement is observed on both quasi-3D and full 3D cases. A lumped approach to compute deterministic fluxes is also proposed to further reduce the computational cost of the spectral method. The spectral method is formulated in such a way that it can be easily implemented into an existing harmonic flow solver by adding an extra source term, and can be potentially used as an efficient tool for aeromechanical and aeroacoustics design of turbomachinery blade rows.


2013 ◽  
Vol 23 (04) ◽  
pp. 1340011 ◽  
Author(s):  
FAISAL SHAHZAD ◽  
MARKUS WITTMANN ◽  
MORITZ KREUTZER ◽  
THOMAS ZEISER ◽  
GEORG HAGER ◽  
...  

The road to exascale computing poses many challenges for the High Performance Computing (HPC) community. Each step on the exascale path is mainly the result of a higher level of parallelism of the basic building blocks (i.e., CPUs, memory units, networking components, etc.). The reliability of each of these basic components does not increase at the same rate as the rate of hardware parallelism. This results in a reduction of the mean time to failure (MTTF) of the whole system. A fault tolerance environment is thus indispensable to run large applications on such clusters. Checkpoint/Restart (C/R) is the classic and most popular method to minimize failure damage. Its ease of implementation makes it useful, but typically it introduces significant overhead to the application. Several efforts have been made to reduce the C/R overhead. In this paper we compare various C/R techniques for their overheads by implementing them on two different categories of applications. These approaches are based on parallel-file-system (PFS)-level checkpoints (synchronous/asynchronous) and node-level checkpoints. We utilize the Scalable Checkpoint/Restart (SCR) library for the comparison of node-level checkpoints. For asynchronous PFS-level checkpoints, we use the Damaris library, the SCR asynchronous feature, and application-based checkpointing via dedicated threads. Our baseline for overhead comparison is the naïve application-based synchronous PFS-level checkpointing method. A 3D lattice-Boltzmann (LBM) flow solver and a Lanczos eigenvalue solver are used as prototypical applications in which all the techniques considered here may be applied.


2021 ◽  
pp. 1-27
Author(s):  
Feng Wang ◽  
Luca di Mare

Abstract Turbomachinery blade rows can have non-uniform geometries due to design intent, manufacture errors or wear. When predictions are sought for the effect of such non-uniformities, it is generally the case that whole assembly calculations are needed. A spectral method is used in this paper to approximate the flow fields of the whole assembly but with significantly less computation cost. The method projects the flow perturbations due to the geometry non-uniformity in an assembly in Fourier space. Only one passage is required to compute the flow perturbations corresponding to a certain wave-number of geometry variation. The performance of this method on transonic blade rows is demonstrated on a modern fan assembly. Low and high engine order geometry non-uniformity (e.g. “saw-tooth” pattern) are examined. The non-linear coupling between the flow perturbations and the passage-averaged flow field is also demonstrated. Pressure variations on the blade surface and the potential flow field upstream of the leading edge from the proposed method and the direct whole assembly solutions are compared. Good agreement is observed on both quasi-3D and full 3D cases. A lumped approach to compute deterministic fluxes is also proposed to further reduce the computational cost of the spectral method. The spectral method is formulated in such a way that it can be easily implemented into an existing harmonic flow solver by adding an extra source term, and can be used as an efficient tool for aeromechanical and aeroacoustics design of turbomachinery blade rows.


2010 ◽  
Vol 229 (19) ◽  
pp. 6715-6731 ◽  
Author(s):  
Qibing Li ◽  
Kun Xu ◽  
Song Fu

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