Finite Element Analysis (FEA) is a powerful and well recognized tool used in the analysis of heat transfer problems. However, FEA can only analyze solid bodies and, by necessity thermal analysis with FEA is limited to conductive heat transfer. The other two types of heat transfer: convection and radiation must by approximated by boundary conditions. Modeling all three mechanisms of heat transfer without arbitrary assumption requires a combined use of FEA and Computational Fluid Dynamics (CFD).
Aircraft equipment is operated in a wide range of external conditions, which, with a certain combination of environmental parameters, can lead to icing of the engine internal elements. Due to icing, the engine components performance change what leads to decrease in thrust, gas dynamic stability, durability, etc. Safe aircraft operation and its desired performance may be lost as a result of such external influence. Therefore, it is relevant to study the possibilities of reducing the icing effect with the help of a special engine control. The focus of this paper is to determine control methods of an aircraft gas turbine engine addressing this problem. The object of the study is a modern commercial turbofan with a bypass ratio of about 9. In this paper analysis of the effect of ice crystal icing on the engine components performance is conducted.
In this study we examined numerically the electrostatic spray transfer processes in the rotary bell spray applicator, which is this case implemented in a full 3D representation. The algorithm implemented and developed for this simulation includes airflow, spray dynamics, tracking of paint droplets and an electrostatic modularized solver to present atomization and in-flight spray phenomena for the spray forming procedure. The algorithm is implemented using the OpenFOAM package. The shaping airflow is simulated via an unsteady 3D compressible Navier-Stokes method. Solver for particle trajectory was developed to illustrate the process of spray transport and also the interaction of airflow and particle that is solved by momentum coupling. As the numerical results in this paper indicates dominant operating parameter voltage setting, further the charge to mass ratio and air-paint flow rate deeply effect the spray shape and the transfer efficiency (TE).
With the actual tendency of space exploration, hypersonic flight have gain a significant relevance, taking the attention of many researchers over the world. This work aims to present a numerical tool to solve hypersonic gas dynamic flows for space propulsion geometries. This will be done by validating the code using two well-known hypersonic test cases, the double cone and the hollow cylinder flare. These test cases are part of NATO Research and Technology Organization Working Group 10 validation of hypersonic flight for laminar viscous-inviscid interactions. During the validation process several important flow features of hypersonic flow are captured and compared with available CFD and numerical data. Special attention is taken to the phenomenon of vibrational excitation of the molecules. Different vibrational non-equilibrium models are used and compared with the available data. The pressure and the heat flux along the surfaces are also analyzed.
The paper presents the numerical approach to simulation and optimization of A350 S19 splice assembly process. The main goal is to reduce the number of installed temporary fasteners while preventing the gap between parts from opening during drilling stage. The numerical approach includes computation of residual gaps between parts, optimization of fastener pattern and validation of obtained solution on input data generated on the base of available measurements. The problem is solved with ASRP (Assembly Simulation of Riveting Process) software. The described methodology is applied to the optimization of the robotized assembly process for A350 S19 section.
The demonstrator project RACER is developed under the leadership of Airbus Helicopters Group within a large European partnership and concerns the development of new VTOL formula in order to fill the mobility gap between conventional helicopters and airplanes. Thus, RACER is a compound rotorcraft including wings and propellers. The new wing arrangement suggested by Airbus Helicopters Groups is defined as a staggered bi-plane configuration with an upper and a lower straight wing at each side of the helicopter, both being interconnected at their outermost tips, forming a triangular framework. Responsible for the design, manufacturing and assembly of the wings is ASTRAL consortium consisted of GE Aviation and University of Nottingham. The identification of the best strategy to assemble the joined wing configuration is quite challenging. In order to ensure that the final wing assembly will fit to the fuselage, a jig that simulates the fuselage was suggested by Airbus Helicopters Group.
ASRP (Assembly Simulation of Riveting Process) software is a special tool for modelling assembly process for large scale airframe parts. On the base of variation simulation, ASRP provides a convenient way to analyze, verify and optimize the arrangement of temporary fasteners. During the airframe assembly process certain criteria on the residual gap between parts must be fulfilled. The numerical approach realized in ASRP allows one to evaluate the quality of contact on every stage of the assembly process and solve verification and optimization problems for temporary fastener patterns. The paper is devoted to description of several specialized approaches that combine statistical analysis of measured data and numerical simulation using high-performance computing for optimization of fastener patterns, calculation of forces in fasteners needed to close initial gaps and identification of hazardous areas in junction regions.
The demanded development towards various emission reduction goals set up by several institutions forces the aerospace industry to think about new technologies and alternative aircraft configurations. With these alternative aircraft concepts, the landing gear layout is also affected. Turbofan engines with very high bypass ratios could increase the diameter of the nacelles extensively. In this case, mounting the engines above the wing could be a possible arrangement to avoid an exceedingly long landing gear. Thus, the landing gear could be shortened and eventually mounted at the fuselage instead of the wings. Other technologies such as high aspect ratio wings have an influence on the landing gear integration as well. To assess the difference, especially in weight, between the conventional landing gear configuration and alternative layouts a method is developed based on preliminary structural designs of the different aircraft components, namely landing gear, wing and fuselage.
Flow separation is among the major causes of aerodynamic drag experience by wings. Vortex generators are regularly used as a means of flow separation control in wings, their use leading to delayed flow separation and drag reduction. A disadvantage of external vortex generators has been observed to be high momentum loss and inefficiency in vortex generation. Internal vortex generators minimize the penalty of momentum loss and generate vortices closer to the surface. In this paper, the impact of the length of internal vortex generators on the aerodynamic characteristics of a wing have been investigated. Internal vortex generators have been placed at 30% chord distance on the suction side of a NACA 0012 airfoil. Analysis is carried out using the Computational Fluid Dynamics software ANSYS Fluent. The length of the vortex has been varied between H and 5H, H being the thickness of the boundary layer, at air flow Reynolds Number between 1,000,000 and 5,000,000.
Vortex generators are aerodynamic devices generally used to delay local air separation and stalling. Conventional vortex generators are external and located normal to the surface with a yaw angle against the flow. However, external vortex generators lead to high momentum loss in the boundary layer, producing inefficient vortices which separate from the surface. They hence do not reenergise the boundary layer to a large extent, in order to allow for delayed flow separation. In order to reduce this loss, internal vortex generators may be used. The effect of internal vortex generators has been investigated on a NACA 0012 airfoil using the Computational Fluid Dynamics software ANSYS Fluent. As the effect of a vortex on the boundary layer is inherently three-dimensional, the numerical analysis of an internal vortex generator is limited to a three-dimensional simulation of the flow.
This paper raises a coupling system of aircraft environmental control and fuel tank inerting based on membrane separation. The system applies a membrane dehumidifier to replace water vapor removal unit of heat regenerator, condenser and water separator, which is widely used in conventional aircraft environmental control system nowadays. Water vapor can travel across the membrane wall under its pressure difference without phase change, so the dehumidification process consumes no cooling capacity and the cooling capacity of the system increases. This paper first compares the thermodynamic properties of environmental control system based on membrane dehumidification and the environmental control system based on condensation. The results show that the membrane dehumidification system has bigger cooling capacity and lighter weight.