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The purpose of this paper is to investigate and improve a new method of unstructured rotational dynamic overset grids, which can be used to simulate the unsteady flows around rotational parts of aircraft.

The computational domain is decomposed into two sub-domains, namely, the rotational sub-domain which contains the rotational boundaries, and the stationary sub-domain which contains the remainder flow field including the stationary boundaries. The artificial boundaries and restriction boundaries are used as the restriction condition to generate the entire computational grid, and then the overset grids are established according to the radius parameters of artificial boundaries set previously. The deformation of rotational boundary is treated by using the linear spring analogy method which is suitable for the dynamic unstructured grid. The unsteady Navier-Stokes/Euler equations are solved separately in the rotational sub-domain and stationary sub-domain, and data coupling is accomplished through the overlapping area. The least squares method is used to interpolate the flow variables for the artificial boundary points with a higher calculating precision. Implicit lower-upper symmetric-Gauss-Seidel (LU-SGS) time stepping scheme is implemented to accelerate the inner iteration during the unsteady simulation.

The airfoil steady flow, airfoil pitching unsteady flow, three-dimensional (3-D) rotor flow field, rotor-fuselage interaction unsteady flow field and the flutter exciting system unsteady flow field are numerically simulated, and the results have good agreements with the experimental data. It is shown that the present method is valid and efficient for the prediction of complicated unsteady problems which contain rotational dynamic boundaries.

The results are entirely based on computational fluid dynamics (CFD), and the 3D simulations are based on the Euler equations in which the viscous effect is ignored. The current work shows further applicable potential to simulate unsteady flow around rotational parts of aircraft.

The current study can be used to simulate the two-dimensional airfoil pitching, 3-D rotor flow field, rotor-fuselage interaction and the flutter exciting system unsteady flow. The work will help the aircraft designer to get the unsteady flow character around rotational parts of aircraft.

A new type of rotational dynamic overset grids is presented and validated, and the current work has a significant contribution to the development of unstructured rotational dynamic overset grids.

This paper aims to study the fluid flow and forced convection heat transfer from an isothermal circular cylinder with control rods in the laminar unsteady flow regime.

The overset grid method was used for accurate simulation of the unsteady flows around different arrangements of the cylinders. Grid generation for overset grids was performed using a general orthogonal boundary fitted coordinate system. The method of solution was based on a finite volume discretization of the Navier-Stokes equations. Simulations were carried out for the Prandtl numbers of 0.7 and 7.0 with the Reynolds numbers ranging from 60 to 300.

The results indicate that the performance of multiple control rods depends strongly on the spacing ratio. Furthermore, in a manner similar to the flow patterns, four different thermal regimes were recognized based on the variations of mean Nusselt number versus G/D, as the thermal regimes follow the categories of flow regimes at different diameter ratios. However, for different Prandtl numbers, no single trend of heat transfer variation versus the spacing ratio exists for same regime.

Few studies have been conducted to investigate the heat transfer characteristics from control rods. The results of this study provide a comprehensive knowledge on the dynamical and thermal behavior of the flow around multiple cylinders.

Turbomachinery, including pumps, are mainly designed to extract/produce energy from/to the flow. A major challenge in the numerical simulation of turbomachinery is the inlet flow rate, which is routinely treated as a known boundary condition for simulation purposes but is properly a dependent output of the solution. As a consequence, the results from numerical simulations may be erroneous due to the incorrect specification of the discharge flow rate. Moreover, the transient behavior of the pumps in their initial states of startup and final states of shutoff phases has not been studied numerically. This paper aims to develop a coupled procedure for calculating the transient inlet flow rate as a part of the solution via application of the control volume method for linear momentum. Large eddy simulation of a four-blade axial hydraulic pump is carried out to calculate the forces at every time step. The sharp interface immersed boundary method is used to resolve the flow around the complex geometry of the propeller, stator and the pipe casing. The effect of the spurious pressure fluctuations, inherent in the sharp interface immersed boundary method, is damped by local time-averaging of the forces. The developed code is validated by comparing the steady-state volumetric flow rate with the experimental data provided by the pump manufacturer. The instantaneous and time-averaged flow fields are also studied to reveal the flow pattern and turbulence characteristics in the pump flow field.

The authors use control volume analysis for linear momentum to simulate the discharge rate as part of the solution in a large eddy simulation of an axial hydraulic pump. The linear momentum balance equation is used to update the inlet flow rate. The sharp interface immersed boundary method with dynamic Smagorinsky sub-grid stress model and a proper wall model is used.

The steady-state volumetric flow rate has been computed and validated by comparing to the flow rate specified by the manufacturer at the simulation conditions, which shows a promising result. The instantaneous and time averaged flow fields are also studied to reveal the flow pattern and turbulence characteristics in the pump flow field.

An approach is proposed for computing the volumetric flow rate as a coupled part of the flow solution, enabling the simulation of turbomachinery at all phases, including the startup/shutdown phase. To the best of the authors’ knowledge, this is the first large eddy simulation of a hydraulic pump to calculate the transient inlet flow rate as a part of the solution rather than specifying it as a fixed boundary condition. The method serves as a numerical framework for simulating problems incorporating complex shapes with moving/stationary parts at all regimes including the transient start-up and shut-down phases.

This study aims to present a moving grid method based on the manipulation of connections.

In this study, the grid’s connections were manipulated to simulate a released store’s displacement. The selected model in this research is the EGLIN test case. In the introduced method, connections are modified in specific nodes of the grid. Governing flow equations were solved with the finite volume method. The major characteristic of this technique is using the averaging method for calculating the flux of cells.

This method maintains the grid’s quality even in large displacements of the released store. The three-dimensional simulation was carried out in transonic and supersonic regimes. Comparison of the results with experimental data were highly satisfactory.

Using this moving grid method is recommended for simulating other models.

Prediction of store trajectory released from air vehicles is one of the most critical issues under study especially in the design of new stores.

The most prominent advantage of this method is maintaining the grid quality simultaneous with large displacements of the released store.

This paper aims to improve the computational efficiency and to achieve high-order accuracy for the computation of helicopter rotor unsteady flows in forward flight during the industrial preliminary design stage.

The integral arbitrary Lagrangian–Eulerian form of unsteady compressible Navier–Stokes equations with low Mach number preconditioned pseudo time terms based on non-inertial frame of reference undergoing rotating and translating was derived and discretized in the framework of multi-block structured finite volume grid using three types of spatial reconstruction schemes, i.e. the third-order accurate monotonic upwind scheme for conservation laws, the fifth-order accurate weighted essentially non-oscillatory and the fifth-order accurate weighted compact nonlinear schemes.

The results show that the present non-inertial computational method can obtain comparable results with other methods, such as the dynamic overset method, and make sure that the higher-order spatial schemes can significantly improve the tip vortex resolution.

The computational grid used by the present method remained static during the whole unsteady computation process, with only local deformations induced by blade cyclic pitch and other operating motions, which greatly reduced the complexity of grid motion and enhanced the efficiency and robustness.

The purpose of this paper is to develop parallel algorithms for moving boundary simulations by local remeshing and compose them to a fully parallel simulation cycle for the solution of problems with engineering interests.

The moving boundary problems are solved by unsteady flow computations coupled with six-degrees-of-freedom equations of rigid body motion. Parallel algorithms are developed for both computational fluid dynamics (CFD) solution and grid deformation steps. Meanwhile, a novel approach is developed for the parallelization of the local remeshing step. It inputs a distributed mesh after deformation, then marks low-quality elements to be deleted on the respective processors. After that, a parallel domain decomposition approach is used to repartition the hole mesh and then to redistribute the resulting sub-meshes onto all available processors. Then remesh individual sub-holes in parallel. Finally, the element redistribution is rebalanced.

If the CFD solver is parallelized while the remaining steps are executed in sequential, the performance bottleneck of such a simulation cycle is observed when the simulation of large-scale problem is executed. The developed parallel simulation cycle, in which all of time-consuming steps have been efficiently parallelized, could overcome these bottlenecks, in terms of both memory consumption and computing efficiency.

A fully parallel approach for moving boundary simulations by local remeshing is developed to solve large-scale problems. In the algorithm level, a novel parallel local remeshing algorithm is present. It repartitions distributed hole elements evenly onto all available processors and ensures the generation of a well-shaped inter-hole boundary always. Therefore, the subsequent remeshing step can fix the inter-hole boundary involves no communications.

The design of a space launcher requires some considerations about the unsteady loads and heat transfer occurring at the base of the structure. In particular, these phenomena are predominant during the early stage of the flight. This paper aims to evaluate the ability of the unstructured, high order finite-volume CFD solver FLUSEPA, developed by Airbus Safran Launchers, to accurately describe these phenomena.

This paper first performs a steady simulation on a base flow around a four-clustered rocket configuration. Results are compared with NASA experiments and Loci-CHEM simulations. Then, unsteady simulations of supersonic H2/air reacting mixing layer based on the experiment of Miller, Bowman and Mungal are performed. Three meshes with different cells number are used to study the impact of spatial resolution. Instantaneous and time-averaged concentrations are compared with the combined OH/acetone planar laser-induced fluorescence imaging from the experiment.

FLUSEPA satisfactorily predicts the base heat flux at the base of a four-clustered rocket configuration. NASA Loci-CHEM reactive simulations indicate that afterburning plays an important role and should not be neglected. The unsteady reactive computation of a supersonic mixing layer shows that FLUSEPA is also able to accurately predict flow structures and interactions. However, the complexity of the experiment and the lack of details concerning the facility prevents from obtaining satisfactory converged results.

This study is the first step on the development of a cost-effective method aiming at predicting unsteady loads and heat transfer on space launchers using an unsteady and reactive model for the CDF calculations. It uses original techniques such as conservative CHIMERA-like overset grids, local re-centering of fluxes and local adaptive time-stepping to reduce computational cost while being robust and accurate.

Computational fluid dynamics (CFD)/computational structural dynamics (CSD) coupling analysis is an important method in the research of helicopter aeroelasticity due to its high precision. However, this method still suffers from some problems, such as wake dissipation and large computational cost. In this study, a new coupling method and a new air load correction method that combine the free wake model with the CFD/CSD method are proposed to maintain computational efficiency whilst solving the wake dissipation problem of the prior coupling methods.

A new coupling method and a new air load correction method that combine the free wake model with the CFD/CSD method are proposed. With the introduction of the free wake model, the CFD solver can adopt two-order accuracy schemes and fewer aerodynamic grids, thus maintaining computational efficiency whilst solving the wake dissipation problem of the prior coupling methods.

Compared with the predictions of the prior methods and flight test data, those of the proposed method are more accurate and closer to the test data. The difference between the two methods in high-speed forward flight is minimal.

Because of the chosen research approach, the research results may lack generalisability. Therefore, researchers are encouraged to test the proposed method further.

In this paper, a CFD/CSD/free wake coupling method is proposed to improve the computational accuracy of the traditional CFD/CSD coupled method and ensure the computational efficiency.

The purpose of this paper is to carry out an extensive numerical study in order to understand the flow structures and the resulting noise generated by a supersonic impinging jet on a flat plate. One of the parameters varied in this study is the distance between the jet exit plane and the flat plate.

Because of the unsteady nature of the problem a time-dependent computation is carried out using the detached eddy simulation turbulence model. The OVERFLOW 2 CFD code was used with a highly resolved grid and small time steps.

The authors found that as the separation distance increases, the dominant frequencies in the noise spectrum decrease. In addition, the relative strength of the various frequencies to each other changes with changing distance, indicating the changing modes of the jet. The CFD results indicate a strong interaction between the acoustic waves emanating from the impingement plate and the jet plume. This feedback mechanism is responsible for destabilizing the jet shear layer leading to the jet changing modes. The computed near field spectra, convection velocities of the jet vortical structures and mean jet centerline velocity profile are in good agreement with experimental measurements. The results also show very high sound pressure levels all over the impingement plate but especially near the impingement point. These levels, if sustained, are detrimental to both human operators as well as the surrounding structures.

Given the large-scale nature of the computations carried out, it is very costly to run the computations long enough to collect a good, statistically steady time sample to achieve a low frequency bandwidth resolution. Such a long time sample could actually improve the results in terms of frequency resolution and obtained an even better agreement with experiments. Off course there is always the issue of grid resolution as well, but given the good agreement with experiments that the authors obtained, the authors are confident in their results.

The practical implications of the results the authors obtained are significant in that, the authors now know that hybrid RANS-large eddy simulation methods can be used for this complex, unsteady engineering problems. In addition, the results also show the high noise level both on the impingement surface and in the surroundings of the jet. This could have a negative impact on the structural integrity of the flat surface.

Noisy environments are never desirable anywhere especially in places where human operations take place. Therefore, given the high noise levels obtained in the simulations and confirmed by experiments, any human presence around the jet will be harmful to hearing and precautions need to be taken.

This is a physics-based study; i.e. understanding the physical phenomena involved in supersonic jet impingement. Of particular interest is the interaction of the jet shear layer with the acoustic waves emanating from the impingement area. This feedback loop is found to be responsible for intensifying the instability of the jet shear layer.

This paper aims to assess the strengths and weaknesses of four thermodynamic models used in aircraft icing simulations to orient the development or the choice of an improved thermodynamic model.

Four models are compared to assess their capabilities: Messinger, iterative Messinger, extended Messinger and shallow water icing models. They have been implemented in the aero-icing framework, NSCODE-ICE, under development at Polytechnique Montreal since 2012. Comparison is performed over typical rime and glaze ice cases. Furthermore, a manufactured geometry with multiple recirculation zones is proposed as a benchmark test to assess the efficiency in runback water modeling and geometry evolution.

The comparison shows that one of the main differences is the runback water modeling. Runback modeling based on the location of the stagnation point fails to capture the water film behavior in the presence of recirculation zones on airfoils. However, runback modeling based on air shear stress is more suitable in this situation and can also handle water accumulation while the other models cannot. Also, accounting for the conduction through the ice layer is found to have a great impact on the final ice shape as it increases the overall freezing fraction.

This paper helps visualize the effect of different thermodynamic models implemented in the same aero-icing framework. Also, the use of a complex manufactured geometry highlights weaknesses not normally noticeable with classic ice accretion simulations. To help with the visualization, the ice shape is presented with the water layer, which is not shown on typical icing results.