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The purpose of this paper is to study the structure and dynamic development of a pair of co-rotating trailing vortices, during their formation, interaction and merging, using detailed experimental measurements of the velocity and vorticity fields.

The vortices were generated using two half wings (NACA0030) positioned at equal and opposite angles of attack at the entrance of the test section of an open-circuit, subsonic, wind tunnel. Velocity vector measurements were obtained at Re_c = 133,000, on cross-plane grids at several distances from the trailing edges of the wings, using an in-house developed four-sensor hot wire anemometer probe.

The results include cross-plane contour plots of the mean and fluctuating velocity as well as mean vorticity fields. Each of these variables is affected in a different way, providing complementary information on the development of the flow field. After shedding, the two vortices are swept along the stream-wise direction and spiral around each other, thereby developing a braid of two vortices, which then deforms the external flow field. Gradually, the interaction with the external flow field links both vortices together until the final merging and the formation of a new stable linear vortex emerges.

Trailing vortices have been rendered particularly important during the past decades, because of increasing traffic density of very heavy aircrafts and several plane “incidents”, which were attributed to the action of the vortex wake.

The presented results provide information on the evolution and merging of a pair of vortices formed by a closely spaced differential wing configuration. The vortices interact almost immediately after shedding as expected in flap–flap or flap–wing vortices interaction.

The purpose of this paper is to present the results of the preliminary design and optimization of the air-intake system and the engine nacelle. The work was conducted as part of an integration process of a turboprop engine in a small aircraft in a tractor configuration.

The preliminary design process was performed using a parametric, interactive design approach. The parametric model of the aircraft was developed using the PARADES™ in-house software. The model assumed a high level of freedom concerning shaping all the components of aircraft important from the point of view of the engine integration. Additionally, the software was used to control the fulfillment of design constraints and to analyze selected geometrical properties. Based on the developed parametric model, the preliminary design was conducted using the interactive design and optimization methodology. Several concepts of the engine integration were investigated in the process. All components of the aircraft propulsion system were designed simultaneously to ensure their compliance with each other.

The concepts of the engine integration were modified according to changes in the design and technological constraints in the preliminary design process. For the most promising configurations, computational fluid dynamics (CFD) computations were conducted using commercial Reynolds-averaged Navier–Stokes solver FLUENT™ (ANSYS). The simulations tested the flow around the nacelle and inside the air-delivery system which consists of the air-intake duct, the foreign-particles separator and the auxiliary ducts delivering air to the cooling and air-conditioning systems. The effect of the working propeller was modeled using the Virtual Blade Model implemented in the FLUENT code. The flow inside the air-intake system was analyzed from the point of view of minimization of pressure losses in the air-intake duct, the quality of air stream delivered to the engine compressor and the effectiveness of the foreign particles separator.

Based on results of the CFD analyses, the final concept of the turboprop engine integration has been chosen.

The presented results of preliminary design process are valuable to achieve the final goal in the ongoing project.

This paper aimed to study the simulation and describe the turbulent fluid flow through a symmetrical tube with a propeller disk set inside it. The Navier–Stokes equations with the model of turbulence (k-ω) are used to describe this problem in space and time.

The propeller disk is represented by the distribution of the vector of velocities along its radius. The main purpose is to describe the boundary conditions at the inlet, at the outlet and special compatible conditions for the simulation of the propeller disk on the both sides. A one-side modification of the Riemann problem is used for the boundary value conditions. Total pressure and total density values and the angle of attack equal to zero are to be used preferentially at the inlet, whereas pressure should be used at the outlet. At the back side of the propeller disk, it is advantageous to use total density and total pressure distributions coming from the distribution of axial velocities on the disk and the total state values at the inlet, with extra-added velocities of rotation. At the front side of the propeller disk, it is preferable to use the distribution of the flowing mass known from the state values computed on the disk.

This set of boundary conditions allows simulation of the air flow twisting behind the propeller/fan including increases in the corresponding pressure.

The advantage of this approach is the possibility to solve axial cuts of air ducts. Similarly, it is possible to solve air flow around the engine nacelle of the propeller aircraft. By this approach, it is possible to separate the design of the axial cut air duct from the propeller solution.

This approach has been used for new air duct designed on the operating conditions with Star-CCM+ solver.

The aim of this work was to focus on the creation of a mathematical model for the optimization problem and the methodology for selection of the most important parameters. The engine mount design optimization problem requires a selection of the most important parameters, which influence the final solution.

A two-tier optimization algorithm was considered. The geometry of the engine mount is optimized by first taking into account the sum of loads, and the weight is minimized next.

Several parameters were selected as the essential set of factors for the engine mount design and for optimization to get the minimal weight for the system of the engine and engine mount.

As a case study, the single turboprop engine attached to the front part of fuselage – the so-called tractor propeller configuration – was considered.

This paper contains the proposed methodology, which is going to be used in the efficient systems and propulsion for small aircraft (ESPOSA) project to optimize the newly designed engine mount.

This paper presents a new developed methodology that could be useful for engine mount design.

The purpose of this paper is to present the first-year results of the EU-supported GABRIEL project on the possible use of magnetic levitation (MagLev) technology to assist aircraft take-off and landing (ATOL).

Developing a radically new technology is a complex task. It is based on extensive expert analysis, use of technology identification evaluation and selection methods, principle of the design philosophies and development of the radically new technologies.

A possible solution of using the MagLev technology to assist ATOL was developed and defined, including several original ideas, such as the cart-sledge concept or the unconventional climb principle.

This is a typical “out-of-the-box” project without limitations on the developing new principles and technologies, but it is working on the development of a possible solution within the predictable technical and technological envelopes.

The developed concept should assess whether MagLev technology for the ATOL is feasible, cost-effective and safe.

The developed GABRIEL principle may significantly reduce the noise and chemical emissions in airport regions and increase the efficiency of the air transportation system.

The GABRIEL concept is the first concept for using the MagLev technology to assist the takeoff and landing processes related to the commercial civil aviation.

The purpose of this paper is to assess the validity of Weller's b-ω flamelet model for practical swirl-stabilized combustion applications.

Swirl-stabilized premixed flame behavior is investigated utilizing an atmospheric combustor test rig. Swirl number of the flow is 0.74 with a cold flow Reynolds number of 19,400 based on the hydraulic diameter at the inlet pipe. Operating condition corresponds to an equivalence ratio of 0.7 at a thermal load of 20.4 kW. Reacting flow was seeded with TiO₂ particles, and velocity distribution at the center plane was measured utilizing particle image velocimetry. These results serve as a validation dataset for numerical simulations. An open-source computational fluid dynamics (CFD) code library (OpenFOAM) is used for numerical computations. These unsteady Reynolds averaged Navier Stokes (RANS) computations were performed at the same load condition corresponding to experimental data. Parallel numerical simulations were carried out on 128 processor cores. To resolve turbulence, Menter's k-ω shear stress transport model was utilized; flame behavior, on the other hand, was described by Weller's b-ω flamelet model. A block-structured all-hexahedral mesh was used.

It is observed that two counter-rotating vortices in the main recirculation zone are responsible for flame stabilization. Weak secondary recirculation zones are also present at the sides above the dump plane. Flame front location was inferred from Mie scattering images. Experimental findings show that the flame anchors both on the tip of the center body and also at the rim of the outlet pipe. Numerical simulations capture the complex interactions between the flame and the turbulent flow. These results qualitatively agree with the flame structure observed experimentally.

Swirl-stabilized combustion systems are used in many practical applications ranging from aeroengines to land-based power generation systems. There are implications regarding the understanding of these combustion systems.

Better understanding of combustion systems contributes to better performing turbine engines and reduced emissions with implications for the entire society.

The paper provides experimental insight into the application of a combustion model for a flame configuration of practical interest.

The purpose of this paper is to present the problems of the electromechanical impedance (EMI), especially its applications for structural health monitoring of aircraft bolt-joints and innovative approach of EMI prediction at loosening of bolt-joints.

This experimental study includes the results of a full-scale test of the Mi-8 helicopter tail beam, particularly, its bolt-joints of a beam with other parts of the structure. One of the connecting frames of the tail beam was equipped with piezoelectric transducers (PZT) glued on the surface of the frame near the bolts. The bolts' loosening was investigated by using the EMI technology.

It was demonstrated that loosening of the bolt-joint produces a significant and statistically stable change of the EMI metric. Presumably, both the small shift of resonance frequencies and the EMI magnitude and resistance change are caused mainly by damping variation at the bolt-joint loosening. In this analytical study, the 2D model of a constrained PZT is proposed. In contrast with the existing model, the modal decomposition analysis is used as a universal mean to express the dynamic properties and dynamic responses of both the transducer and the host structure. This approach, together with the finite element modal analysis, allows simulation of any complex system “PZT-host structure”. The model can be easily transformed also to the 3D one. The bolt-joint of the Mi-8 helicopter with the EMI measurement system was simulated by using the developed 2D model. The simulation results satisfactorily correspond to the test.

The results of this research can be used for implementation in the structural health monitoring of bolt-joints and other aerospace structural components.

The new experimental results on aircraft real bolt-joints were obtained. Especially significant is the original 2D model of the electromechanical impedance, based on the modal decomposition method, which can significantly improve the accuracy and the realistic description of the dynamic interaction between PZT and structure, as well as the dynamic response to the appearance of structural damage.

The purpose of this paper is to propose a solution of the engine bypass ratio choice problem of a very light jet (VLJ) class aircraft using the multiple objective optimization (MOO) method.

The work focuses on the choice of one of the most essential parameters of the jet engine, that is its bypass ratio. The work presents the methodology of optimal designing using the multitask character of the matter which is based on the mathematical model of optimization in the concept of the set theory. To make an optimal choice of the chosen parameter, a complete computational model of an aircraftwas made (aerodynamic, power unit, performance and cost) and then the method that allows to choose the bypass ratio was selected, regardingmultiple estimating criteria of the obtained solutions. The presented method can be used at the concept design state for determining the chosen and most important technical parameters of the aircraft.

The way to design a competing aircraft is to choose its design parameters, including the power unit, by using the advanced methods of MOO. The main aim of the work was to demonstrate a method of selecting chosen parameters of the transport aircraft at the preliminary design stage. The work focuses on the choice of bypass ratio of the jet engine of the VLJ. The method could be helpful at the preliminary design stage of a new aircraft to selection of other design parameters.

The exemplary calculations were made for 50 different transport tasks to take into account different performance conditions of the aircraft. The presented method can be used at the concept design state for determining the chosen and most important technical parameters of the aircraft.

The work shows a practical possibility to implement the proposed method. The presented method could be helpful at the preliminary design stage of a new aircraft to select its design parameters. The results of the analyses are a separate point for further research and studies.

The work shows a practical possibility to implement the proposed approach for design problems at early stages of product development.

The purpose of this article is to present numerical investigations of flow control with piezoelectric actuators on a backward facing step (BFS) and fluidic vortex generators on a NACA0015 aerofoil for the reattachment and separation control through the manipulation of the Reynolds stresses.

The unsteady flow phenomena associated with both devices are simulated using Spalart–Allmaras-based hybrid Reynolds averaged Navier-Stokes (RANS)/large eddy simulation (LES) models (detached eddy simulation (DES), delayed detached eddy simulation (DDES) and improved delayed detached eddy simulation (IDDES)), using an in-house computational fluid dynamics (CFD) solver. Results from these computations are compared with experimental observations, enabling their reliable assessment through the detailed investigation of the Reynolds stresses and also the separation and reattachment.

All the hybrid RANS/LES methods investigated in this article predict reasonable results for the BFS case, while only IDDES captures the separation point as measured in the experiments. The oscillating surface flow control method by piezoelectric actuators applied to the BFS case demonstrates that the Reynolds stresses in the controlled case decrease, and that a slightly nearer reattachment is achieved for the given actuation. The fluidic vortex generators on the surface of the NACA0015 case force the separated flow to fully reattach on the wing. Although skin friction is increased, there is a significant decrease in Reynolds stresses and an increase in lift to drag ratio.

The value of this article lies in the assessment of the hybrid RANS/LES models in terms of separation and reattachment for the cases of the backward-facing step and NACA0015 wing, and their further application in active flow control.