Figure 1 - uploaded by Erdal Oktay
Content may be subject to copyright.
Source publication
The purpose of this paper is to validate an unstructured 3D Euler flow solver, USER3D, against the experimental data and to compare it with a classical, structured, Euler solver, FLU3M of ONERA, on two different missile geometries at a Mach number of 2 and at various angles of attack up to 20 ° . The first geometry is a conventional missile with an...
Contexts in source publication
Context 1
... geometry for the conventional missile is given in figure 1. The nose section is a tangent ogive followed by a cylindrical body and four straight tail fins. ...
Context 2
... details of the unconventional missile geometry are given in figure 10. The grid generated for this unconventional missile having a lenticular cross-section is given in figure 11. ...
Context 3
... details of the unconventional missile geometry are given in figure 10. The grid generated for this unconventional missile having a lenticular cross-section is given in figure 11. The total number of cells for this grid is 297 778 with 55 667 nodes and 605 379 faces for half body. ...
Context 4
... variation of the normal force coefficient C N , with angle of attack is indicated in figure 12. It is observed that both of the numerical predictions coincide perfectly well with the experimental data for all of the angles of attack studied. ...
Context 5
... is observed that both of the numerical predictions coincide perfectly well with the experimental data for all of the angles of attack studied. Variation of the surface pressure coefficient C p for the cross sectional plane at x/L = 0.915 is given in figure 13. The comparison of pressure distributions from the present computations with those of FLU3M and the experimental data reveals a perfect agreement on the windward side whereas on the leeward side, both of the computational methods differ from the experimental measurements. ...
Context 6
... same behaviour was also discussed by d'Espiney [5], where the plateau region of the experimental pressure data is attributed to the separated flow. As for the location of the vortex core at this axial location, x/L = 0.915, the predictions of the two computational methods do agree with each other as shown in figure 14, and are not very far from the experiments. It can be added that the strong vortices which appear, are originated from the lateral sharp edges of the body, and not from a separation on a smooth surface (like for the first test case), which can explain why the inviscid solutions are relatively satisfying. ...
Citations
... It can be seen from the compatibility between the computational and experimental results. In addition, it can be proven that computational modeling is quite accurate, fast, and reliable on both conventional and unconventional missiles [7]. In the study conducted by DeSprito, it can be seen that the computational method can give very accurate results compared to experimental data. ...
... Computational Fluid Dynamics (CFD) has become a very important and useful design tool in aerodynamics [1,2]. The theoretical basis of CFD includes fluid mechanics, numerical analysis, computer technology and other fields. ...
Over the last decades, the discontinuous Galerkin (DG) method has demonstrated its excellence in accurate, higher-order numerical simulations for a wide range of applications in aerodynamics simulations. However, the development of practical, computationally accurate flow solvers for industrial applications is still in the focus of active research, and applicable boundary conditions and fluxes are also very important parts. Based on curvilinear DG method, we have developed a flow solver that can be used for solving the three-dimensional subsonic, transonic and hypersonic inviscid flows on unstructured meshes. The development covers the geometrical transformation from the real curved element to the rectilinear reference element with the hierarchical basis functions and their gradient operation in reference coordinates up to full third order. The implementation of solid wall boundary conditions is derived by the contravariant velocities, and an enhanced algorithms of Harten-Lax-van Leer with contact (HLLC) flux based on curved element is suggested. These new techniques do not require a complex geometric boundary information and are easy to implement. The simulation of subsonic, transonic and hypersonic flows shows that the linear treatment can limit the accuracy at high order and demonstrates how the boundary treatment involving curved element overcomes this restriction. In addition, such a flow solver is stable on a reasonably coarse meshes and finer ones, and has good robustness for three-dimensional flows with various geometries and velocities. For engineering practice, a reasonable accuracy can be obtained at reasonably coarse unstructured meshes.
... The grid generation is an essential as well as one of the most time-consuming steps in CFD. Though the unstructured mesh has proven to be reliable, and advantageous [22] however, a structured mesh was adopted according to the expected flow pattern. The computational mesh was generated using ICEM® CFD. ...
The purpose of this paper is to initiate a 3-D non-spinning semi-circular missile model with a single generic planar fin and perform a computational analysis on it to understand the flow pattern around the fin. A freestream computational fluid dynamics analysis was done in the subsonic and supersonic Mach range, in which, the fluid behaviour was investigated using a two-equation turbulence model. A structured mesh was adopted to visualize the flow pattern around the fin. The aerodynamic coefficients were calculated for the model, and the predicted values were compared with the previous experimental results as well as the numerical results. This paper attempts to present all the possible flow visualizations which might help in better understanding of flow around missiles having planar fins. This work also attempts to establish a turbulent computational model for a single planar fin missile model for subsonic to supersonic range.
... The goal is to see the flow pattern about missile with ogive and parabolic shape noses and compare the results with other references. First, we report the results for an ogive shape nose which is also studied in ref [1] and ref [2] using USER3D and Flu3m methods. The second nose has studied is a parabolic shape which we have added in this article. ...
... The second nose has studied is a parabolic shape which we have added in this article. In this study first we consider ogive nose geometry and the numerical Roe method is applied for the same conditions as ref [1] and then comparison of these two results is presented. We report also the results for the parabolic nose. ...
... Our calculation shows how the results change for different parameters like angles of attack and nose shapes. The desired type of simple fin missile at Supersonic flight conditions at Mach 2 and the altitude of 22,800 meters has been considered and compared with the results of ref [1] and ref [2]. ...
This paper studied numerical analysis for a typical land to land missile with conventional geometry. The goal is to obtain the flow pattern about missile with ogive and parabolic shape noses and compare the results with other references. The computational Roe method is applied and it is found that parabolic nose shape at Mach number 2 has less drag coefficient than ogive shape one. However the normal force coefficient, CN is approximately equal in both nose shapes at different angles of attack. INTRODUCTION This paper studied numerical analysis for a typical land to land missile with conventional geometry. The boundary conditions is used for this work are the conditions for this supersonic missile in flight. This article tries to compute the results for different nose geometries. The goal is to see the flow pattern about missile with ogive and parabolic shape noses and compare the results with other references. First, we report the results for an ogive shape nose which is also studied in ref [1] and ref [2] using USER3D and Flu3m methods. The second nose has studied is a parabolic shape which we have added in this article. In this study first we consider ogive nose geometry and the numerical Roe method is applied for the same conditions as ref [1] and then comparison of these two results is presented. We report also the results for the parabolic nose. The nose of supersonic missiles especially at high Mach number plays important role in performance of missile and the possibility of severe heating effects beyond the speed of sound speed (Hypersonic). Therefore a study of aerodynamic of nose shape in supersonic missiles is required for increasing the speed up to desired Mach number. Our calculation shows how the results change for different parameters like angles of attack and nose shapes. The desired type of simple fin missile at Supersonic flight conditions at Mach 2 and the altitude of 22,800 meters has been considered and compared with the results of ref [1] and ref [2].
... As an illustration of the accuracy and efficiency enhancement of the adaptive wavelet method, the flow around a missile is computed as the first case. For the conventional missile geometry with an O-give nose, a cylindrical body and fins is taken [9] . All the present computations on the missile are performed at a fixed free stream M ∞ = 1.5, angle of attack of 5 α ° = and Reynolds number of 100,000. ...
An efficient adaptive wavelet method is proposed for the enhancement of computational efficiency of the Navier-Stokes equations.
The method is based on sparse point representation (SPR), which uses the wavelet decomposition and thresholding to obtain
a sparsely distributed dataset. The threshold mechanism is modified in order to maintain the spatial accuracy of a conventional
Navier-Stokes solver by adapting the threshold value to the order of spatial truncation error. The computational grid can
be dynamically adapted to a transient solution to reflect local changes in the solution. The flux evaluation is then carried
out only at the points of the adapted dataset, which reduces the computational effort and memory requirements. A stabilization
technique is also implemented to avoid the additional numerical errors introduced by the threshold procedure. The numerical
results of the adaptive wavelet method are compared with a conventional solver to validate the enhancement in computational
efficiency of Navier-Stokes equations without the degeneration of the numerical accuracy of a conventional solver.
KeywordsNavier-Stokes equations-Adaptive wavelet method-Convergence acceleration-Sparse point representation
Since 2019, Riga Technical University, Institute of Aeronautics, in partnership with Cryogenic and vacuum systems Ltd (Ventspils), executed the European Regional Development Fund Project No. 1.1.1.1/18/A/133 “Prototype development of transportable in multimodal traffic mobile space testing facility Metamorphosis.” In the process of realization of the project, the analysis of the technical failure risk was provided. Most risk components were detached, and several measures for risk prevention were performed. During the testing of the prototype, it turned out that significant vibration and shock loads acted on the prototype during movement, so some design changes were made. The development of the variational vacuum system calculation model allowed the project team to find the solutions and successfully finish the tests, proofing the concept of the mobile thermal vacuum test facility. This article describes the problem and the search for a solution and shows its implementation.
Some Results of the Mobile Space Testing Facility Metamorphosis Prototype Design, Development and Test
Sergey Kravchenko(1), Nikolay Kuleshov(2), Ilmars Blumbergs(3*), Natalia Kravchenko(4), Vladimir Shestakov(5), Daria Panova(6), Aya Medany(7)
(*) Corresponding author
Authors' affiliations
DOI: https://doi.org/10.15866/irease.v16i3.22588
Abstract
Since 2019, Riga Technical University, Institute of Aeronautics, in partnership with Cryogenic and Vacuum Systems Ltd (Ventspils), executed the European Regional Development Fund Project No. 1.1.1.1/18/A/133 “Prototype development of transportable in multimodal traffic mobile space testing facility Metamorphosis.” In the process of realization of the project, the analysis of the technical failure risk was provided. Most risk components were detached, and several measures for risk prevention were performed. During the testing of the prototype, it turned out that significant vibration and shock loads acted on the prototype during movement, so some design changes were made. The development of the variational vacuum system calculation model allowed the project team to find the solutions and successfully finish the tests, proofing the concept of the mobile thermal vacuum test facility. This article describes the problem and the search for a solution and shows its implementation.
Copyright © 2023 Praise Worthy Prize - All rights reserved.
Keywords: Mobile Thermal Vacuum Test Complex; Computational Model; Vacuum System
A method of predicting pitch and roll damping derivatives of missile geometries with fins using an unsteady RANS (Reynolds-averaged Navier-Stokes) solver was presented. A three-dimensional structured RANS solver based on arbitrary Lagrangian-Eulerian(ALE) formulation with a dynamically deforming mesh algorithm was used and validated by the wind tunnel and ballistic range data available in the literature. Roll and pitch damping derivatives were calculated from load history of the unsteady flow around the model. A standard research configuration, known as the basic finner, was studied under forced pitching and rolling conditions. Pitching and rolling motions with oscillation were analyzed at supersonic Mach numbers ranging from 1.5 to 2.5. Predicted results show good agreement with the available wind tunnel data.
A method to predict pitch and roll damping derivatives of missile geometries using an unsteady Euler solver is presented. A three-dimensional unstructured Euler solver based on the arbitrary Lagrangian - Eulerian (ALE) formulation with a dynamically deforming mesh algorithm is used and validated with the wind tunnel and ballistic range data available in the literature. Roll and pitch damping derivatives are calculated from load history of the unsteady flow around the missile. To validate the applicability of the present method, a standard research configuration, known as the Basic Finner, is studied under forced pitching and rolling conditions. Pitching and rolling motions with constant rates are analyzed at supersonic Mach numbers ranging from 1.5 to 2.5. Predicted results showed good agreement with the available wind tunnel data.
