Search Results

Over the past few years, we have developed a unique approach to simulate aircraft icing numerically; we call this method morphogenetic modelling. Previously, we developed a successful two-dimensional version of the model; the objective of our present research is to show that the morphogenetic modelling approach can be extended to three-dimensional in-flight icing. In this paper, we focus on the simulation of three-dimensional, discrete rime structures forming on the fuselage of a helicopter. The numerical model consists of three components: an airflow solver, a drop trajectory solver, and a morphogenetic ice growth model. The velocity field of the flow is computed using the Euler equations, while the drop trajectories are computed using a Lagrangian approach. Computation of drop impact locations determines the local collision efficiency distribution. The morphogenetic model deals with the processes occurring on the impinging surface.

When conducting experiments in icing wind tunnels (IWTs), a significant question is to what extent the temperature of the water droplets generated by the spray system has converged to the static air temperature when the droplets impinge on the test object. This is a particularly important issue for large droplets, since the cooling rate of droplets decreases sharply with increasing diameter. In this paper, on the one hand, realistic droplet temperature distributions in the measurement section of the Rail Tec Arsenal IWT (located in Vienna) are computed by means of a numerical code which tracks the paths of the droplets from the spraying nozzle to the measurement section and simultaneously calculates their cooling rates. On the other hand, numerical icing simulations are performed to investigate to what extent the deviation of the droplet temperature from static air temperature influences icing and thermal anti-icing processes.

Aerodynamic effects induced from vectoring an exhaust jet are investigated using a well established thin-layer Reynolds averaged Navier-Stokes code. This multiple block code has been modified to allow for the specification of jet properties at a block face. The applicability of the resulting code for thrust vectoring applications is verified by comparing numerically and experimentally determined pressure coefficient distributions for a jet-wing afterbody configuration with a thrust-vectoring 2-D nozzle. Induced effects on the body and nearby wing from thrust vectoring are graphically illustrated.

In this study, computational methods are presented that compute ice accretion on multiple-element airfoils in specified icing conditions. The “Droplerian” numerical simulation method used is based on an Eulerian method for predicting droplet trajectories and the resulting droplet catching efficiency on the surface of the configuration. Flow field and droplet catching efficiency form input for Messinger's model for ice accretion. The droplet trajectory method has been constructed such that the solution of any flow-field simulation (e.g., potential-flow, Euler equations) can be used as input for the finite-volume solution method. On an unstructured grid the spatial distribution of droplet loading and droplet velocity are obtained. From these quantities the droplet catching efficiency is derived. Of special interest in this study are the Supercooled Large Droplets (SLD). The simulation of SLD requires a specific splashing model.

The nonlinear development of stationary crossflow vortices over a 45° swept NLF(2)-0415 airfoil is studied. Previous investigations indicate that the linear stability theories are unable to accurately describe the transitional flow over crossflow-dominated configurations. In recent years the development of nonlinear parabolized stability equations (NPSE) has opened new pathways toward understanding transitional boundary-layer flows. This is because the elegant inclusion of nonlinear and nonparallel effects in the NPSE allows accurate stability analyses to be performed without the difficulties and overhead associated with direct numerical simulations (DNS). Numerical (NPSE) results are presented here and compared with experimental results obtained at the Arizona State University Unsteady Wind Tunnel (ASUUWT) for the same configurations.

Even if the probability of a combustion chamber burn-through is unlikely in modern aeroengines, the behavior of the engine has to be considered in term of airworthiness and safety. The whole 3D flow field inside the bypass duct of two aeroengines, BR710 and BR715 were simulated numerically in the event of a combustion chamber burn-through. The simulations showed the effect of different burn-through hole sizes on the bending of the torch and on the maximum temperature of the bypass duct. Utilizing the latest CFD methods the hypothetical case of an additional failure of the bypass duct structure was also simulated. These simulations also showed that both engines retain their integrity for the various burn-through scenarios.

The aerodynamic/propulsive interaction between a subsonic jet exhausting perpendicularly through a flat plate into a crossflow is investigated numerically using an approximately factored, partially flux-split, implicit solver for the three-dimensional, thin-layer Navier-Stokes equations. This algorithm is applied to flows with a range of jet-to-crossflow velocity ratios between 4 and 8. The computations model the jet trajectory, the contrarotating vortex pair and the wake region near the plate downstream of the jet orifice. Both qualitative and quantitative agreement with the existing experimental database are demonstrated. Flow visualization is instructive for understanding the physics of this flowfield.

Electrostatic Rotary Bell Sprayers (ERBSs) have been widely used in the painting industry, especially in the automotive and aerospace industries, due to their superior performance. The effects of the applied voltage and paint droplet charge values on the spraying pattern and coating Transfer Efficiency (TE) in the ERBS, including a high-voltage ring for spray cloud control, have been studied numerically in a wide range of droplet size distribution. A 3D Eulerian-Lagrangian numerical analysis is implemented under the framework of the OpenFOAM package. The fluid dynamics of turbulence, primary and secondary breakup procedures are modeled using a large eddy simulation (LES) model, Rosin-Rammler distribution, and modified TAB approach, respectively.

The main objective of this work is to investigate, by means of numerical simulations, the performance of the engine nacelle ventilation cooling system of a helicopter under hover and forward flight conditions, and to propose a simplified method of evaluating the performance based on rotor downwash flow by taking the synthetical effect of engine nacelle, exhaust ejector and external flow of a helicopter into account. For the engine nacelle of a helicopter, an integrated model of the nacelle and exhaust ejector was set up including the domain of external flow. The unstructured grid and finite volume method were applied for domains and control equations discreteness, and the standard k-ε model was applied for solving turbulent control equations. Using the business CFD software, the flow field and the temperature field in the nacelle were calculated for single inlet scheme and double inlets scheme, total up to 9 schemes. The performance of the exhaust ejector was computed.

A numerical method to simulate aeroacoustic fields around automobiles is proposed in the present paper. The proposed method can be used to compute sound emissions directly in both far fields and near fields. Sound passes through body structures near A-pillars and rear-view mirrors. The direct predictions of the sound to passengers therefore require solutions of acoustic near fields. Most aeroacoustics simulations around automobiles are based on Lighthill's analogy. Strictly speaking, Lighthill's analogy is not consistent in near fields because near fields are not governed by a simple wave equation. In the present paper, a proper approach is proposed to achieve further progress in the simulation of aeroacoustic fields around automobiles. The difficulties occur because the sound pressure is much smaller than the vortical flow pressure.

The problems connected with the simulation of APUs with different configuration are considered. The techniques studied allow to set up end to develop fast, flexible and reliable numerical computer codes. The paper shows the results of calculations for both design-point and off design conditions. Moreover diagnostic and fault simulation are carried out and discussed. At last direct and inverse transient working simulations have been performed.

Hypersonic flows over simple 3-D bodies and a space vehicle are simulated using a real-gas Navier- Stokes code under an equilibrium air assumption. This code is based on 3-D upwind flux splitting scheme with generalized Roe's Riemann solver. The real-gas effect is incorporated using the VEG (Variable Equivalent Gamma) method [1]*. The equivalent gamma and other thermodynamic properties are calculated using empirical curve fits [2]. Numerical simulations are conducted for flow fields around a spherical blunt body, a spherical-nose cylinder, and a cone-cylinder as simple configurations, and HOPE (H-orbiting plane: Japanese spaceplane) as a practical plane configuration. Flow conditons are Mach numbers of 7.72, 15.0 for the blunt bodies, 6.86 for the cone-cylinder, and 15.0 for the HOPE. Computed pressure and density distributions are presented. Results for simple configuration cases are compared with experimental data for the code validation.

Ambient pressure and temperature are two main factors affecting the engine performance. As altitude increases, the air volume and air temperature entering the cylinder per cycle decrease due to the lowering of atmospheric pressure and temperature, which directly affects the engine performance. As a result, engine performance in the plateau environment degrades while the power, economy, and emission performance of the engine significantly deteriorate. This paper focuses on the simulation and parameter optimization of the combustion process of non-road methanol engines, and 1D simulation is for BSFC (Brake Specific Fuel Consumption) prediction while 3D simulation is for soot and NOx (Nitrogen Oxides) predictions. Discusses, analyzes and predicts the feasibility of non-road methanol engines for high altitude conditions. Especially the application of high proportion of methanol in non-road methanol engines at high altitudes.

Dynamic stall phenomena have been investigated numerically by solving incompressible Navier-Stokes equations with a third- order upwind scheme in order to reveal the flow structure and mechanism of dynamic stall. The separated flows around a wing section at static attack angle are calculated and compared with the experiments which are conducted by the present authors. The results show excellent agreements with the experiments and the numerical scheme is proved to work well. Separated flows around oscillating airfoil in pitch are calculated by using the scheme and a moving mesh system. The flow conditions are selected from our experiments. The calculated separated region is small in pitching-up process and it becomes large in a pitching-down process. Quite different characteristics of flow patterns between in a pitching-up and pitching-down processes are obtained.

Analysis of tandem wing aircraft configurations has been of interest to the aerospace community since the early 1970's. The theoretical performance gains from the use of two similarly-sized wings make this unusual configuration an enticing option for future aircraft designs. In this investigation, a two-dimensional Navier-Stokes analysis previously developed for internal flow geometries has been extended to external flow geometries. The modified flow analysis was validated against two sets of experimental data. A series numerical simulations were then performed for a tandem-airfoil configuration in which the stagger (chord-wise distance between the mid-chord of each airfoil) was varied. At each stagger position, the aerodynamic flow field was investigated at several negative and positive incidence angles. The predicted results indicate that (for moderate-to-large stagger distances) the aft airfoil performs similar to the fore airfoil at lower angles of attack.

An axisymmetric slender body comprising an ogive-cylinder (or ellipsoid -cylinder) shell enclosing an engine is analysed at various incidences to an incompressible uniform free stream for several jet flow rates on the basis of integral equation method.The pressure distributions, normal force, pitching moment, drag and thrust are calculated. The results with and without the effect of the exhaust jet flow are presented to show the influences of the jet mass flow rates on the aerodynamic properties.

A numerical method is developed to analyse the flow around an axisymmetric ogive-cylinder and an ellipsoid-cylinder bodies undergoing harmonic pitching motion in an uniform air free stream. The pressure distributions along the lengthwise and over the circumference of the body are calculated with fineness ratio of 3:1 at mean angle of incidence 0° and 5°. Results are presented for a range of frequency parameters and various mean angle of incidence in order to show the influence of the frequency parameter and the mean angle of incidence on the aerodynamic properties.

This paper studies the level of confidence and applicability of CFD simulations using steady-state Reynolds-Averaged Navier-Stokes (RANS) in predicting aerodynamic performance losses on swept-wings due to contamination with ice accreted in-flight. The wing geometry selected for the study is the 65%-scale Common Research Model (CRM65) main wing, for which NASA Glenn Research Center’s Icing Research Tunnel has generated experimental ice shapes for the inboard, mid-span, and outboard sections. The reproductions at various levels of fidelity from detailed 3D scans of these ice shapes have been used in recent aerodynamic testing at the Office National d’Etudes et Recherches Aérospatiales (ONERA) and Wichita State University (WSU) wind tunnels. The ONERA tests were at higher Reynolds number range in the order of 10 million, while the WSU tests were in the order of 1 million.

An overset grid approach is used to analyze a 3-element trapezoidal wing high-lift configuration. A new software system was developed to automate the overset computational fluid dynamics process. A three-dimensional grid resolution study is conducted, and comparisons of numerical results are made to experimental data which were obtained after the simulations. Comparisons between numerical and experimental data are in good agreement for the lift coefficient over a wide range of angles of attack, up to and including CLmax. Comparisons of chordwise distributions of the pressure coefficient between numerical and experimental data are in good agreement for all three elements, except the lift is under-predicted for the tip region when the wing is near CLmax.

This work aims to numerically investigate the aerodynamic characteristics of a multi-element airfoil NACA 23012. The investigation was conducted through Computational Fluid Dynamics (CFD), using ANSYS FLUENT software. The Navier-Stokes equations were solved for turbulent, incompressible flow using k-epsilon model and SIMPLE algorithm. The study was carried out for both take-off / landing conditions and the results were compared to experimental data of the NACA 23012 from wind tunnel tests. The experimental and computational results for drag and lift coefficients match effectively up to pre-stall attack angles. The pressure coefficients, velocity distribution, and wall Y+ data were presented for different angles of attack (0 deg, 4 deg, and 8 deg). The CFD analysis could help acquire a closer and detailed understanding of airfoil performance, which is usually not easy through normal experimentation.