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The purpose of this paper is to introduce a unique technique to couple the two-integral boundary layer solutions to a generic inviscid solver in an iterative fashion.

The boundary layer solution is obtained using the two-integral method to solve displacement thickness point by point with a local Newton method, at a fraction of the cost of a conventional mesh-based, full viscous solution. The boundary layer solution is coupled with an existing inviscid solver. Coupling occurs by moving the wall to a streamline at the computed boundary layer thickness and treating it as a slip boundary, then solving the flow again and iterating. The Goldstein singularity present when solving boundary layer equations is overcome by solving an auxiliary velocity equation along with the displacement thickness.

The proposed method obtained favourable results when compared with the analytical solutions for flat and inclined plates. Further, it was applied to modelling the flow around a NACA0012 airfoil and yielded results similar to those of the widely used XFOIL code.

A unique method is proposed for coupling of the boundary layer solution to the inviscid flow. Rather than the traditional transpiration boundary condition, mesh movement is employed to simulate the boundary layer thickness in a more physically meaningful way. Further, a new auxiliary velocity equation is presented to circumvent the Goldstein singularity.

The study of the electro-osmotic forces (EOF) in the flow of the boundary layer has been a topic of interest in biomedical engineering and other engineering fields. The purpose of this paper is to develop an innovative mathematical model for electro-osmotic boundary layer flow. This type of fluid flow requires sophisticated mathematical models and numerical simulations.

The effect of EOF on the boundary layer Williamson fluid model containing a gyrotactic microorganism through a non-Darcian flow (Forchheimer model) is investigated. The problem is formulated mathematically by a system of non-linear partial differential equations (PDEs). By using suitable transformations, the PDEs system is transformed into a system of non-linear ordinary differential equations subjected to the appropriate boundary conditions. Those equations are solved numerically using the finite difference method.

The boundary layer velocity is lower in the case of non-Newtonian fluid when it is compared with that for a Newtonian fluid. The electro-osmotic parameter makes an increase in the velocity of the boundary layer. The boundary layer velocity is lower in the case of non-Darcian fluid when it is compared with Darcian fluid and as the Forchheimer parameter increases the behavior of the velocity becomes more closely. Entropy generation decays speedily far away from the wall and an opposite effect occurs on the Bejan number behavior.

The present outcomes are enriched to give valuable information for the research scientists in the field of biomedical engineering and other engineering fields. Also, the proposed outcomes are hopefully beneficial for the experimental investigation of the electroosmotic forces on flows with non-Newtonian models and containing a gyrotactic microorganism.

THE drag of a streamline body, such as an airship hull or an aeroplane fuselage, with its axis in the direction of the relative wind, is almost entirely due to skin friction. The drag component clue to surface pressure is, in general, small and relatively unimportant. Unfortunately, in the range of Reynolds numbers obtainable in most wind tunnels, the drag due to skin friction is abnormally sensitive to the degree of initial turbulence in the wind tunnel stream and to the shape of the model, particularly near the bow. Wind tunnel tests on models of airship hulls at a Reynolds number RL= U0L/v (where L is the length of the model and Uo the velocity in the free stream) of the order of 10° show wide variations in the values obtained for the drag coefficients of the same model in different wind tunnels or for different models in the same tunnel. At Reynolds numbers of about 4 × 108, appropriate to a full‐scale airship, it is impossible to carry out a reliable series of tests, but there is reason to believe that careful streamlining is not important and that a shape which is usually considered a “poor” streamline will have as low a resistance on the full scale as the shapes adopted for “R100” and “R101,” which, in the wind tunnel, appeared to be the best obtainable. From the structural point of view the “good” streamline shape proved to be a distinct disadvantage owing to the restricted gas volume in the bow and stem and the consequent lack of lift to balance the weights of fins and mooring gear, which produced increased bending moments throughout the ship.

To provide an effective numerical method for analysis and design of aerodynamic characteristics of unmanned aerial vehicles basing on commercial package VSAERO.

Calculation was made by VSAERO package, which is based on a classical panel method enhanced on boundary layer method. Paper explains how to use efficiently VSAERO package, which utilizes advanced CAD techniques, in modern designing of unmanned aircraft.

During aerodynamic analysis of unmanned aircraft the computing cycle is repeated many times until the required accuracy is obtained and when the best performance of an aircraft is achieved. Design process depends on the number of iterations. If the preliminary configuration (the so‐called starting design point) is well selected and the aerodynamic analysis is completed in a relatively short time, then the overall design time will be shortened.

The panel method is very useful tool in spite of different limitations. For example, the Reynolds number has to be sufficiently high, angles of attack and sideslip have to be small enough. Computational process is relatively fast and the accuracy depends on the geometry representation. The boundary layer included into the computational model improves the accuracy of aerodynamic calculations. This methodology is limited to subsonic and low transonic speeds.

A very useful source of computational information and patterns to follow, especially for engineering students and engineers dealing with aerodynamic of unmanned aviation. Surface panel geometry can be transferred from UNIGRAPHICS via IGES files or can be generated from scratch using SPING or PEP software.

This paper offers a practical help for designers planning to develop a new unmanned platform. VSAERO package appeared to be a very useful tool for aerodynamic calculation in the full cycle design activity. This software utilizes the panel method enhanced on a boundary layer model for determination of the fundamental aerodynamic characteristic of an arbitrary aircraft. Presented paper shows a very efficient way how to compute the aerodynamics necessary for design of a new MALE class UAV.

Wall-modeled large eddy simulation (LES) is a practical tool for solving wall-bounded flows with less computational cost by avoiding the explicit resolution of the near-wall region. However, its use is limited in flows that have high non-equilibrium effects like separation or transition. This study aims to present a novel methodology of using high-fidelity data and machine learning (ML) techniques to capture these non-equilibrium effects.

A precursor to this methodology has already been tested in Radhakrishnan et al. (2021) for equilibrium flows using LES of channel flow data. In the current methodology, the high-fidelity data chosen for training includes direct numerical simulation of a double diffuser that has strong non-equilibrium flow regions, and LES of a channel flow. The ultimate purpose of the model is to distinguish between equilibrium and non-equilibrium regions, and to provide the appropriate wall shear stress. The ML system used for this study is gradient-boosted regression trees.

The authors show that the model can be trained to make accurate predictions for both equilibrium and non-equilibrium boundary layers. In example, the authors find that the model is very effective for corner flows and flows that involve relaminarization, while performing rather ineffectively at recirculation regions.

Data from relaminarization regions help the model to better understand such phenomenon and to provide an appropriate boundary condition based on that. This motivates the authors to continue the research in this direction by adding more non-equilibrium phenomena to the training data to capture recirculation as well.

The purpose of this paper is to propose a new finite element method (FEM) solving strategy for efficient analysis of the challenging edge effect problem in plate structures. Its main ideas are to develop special-purpose plate element models to effectively simulate the behaviors in the plate’s edge zones near free/SS1 edges.

These new elements are developed based on the hybrid-Trefftz element method. During their construction procedures, the analytical solutions of the edge effect problem, which are in exponential forms, are used to enhance the interior displacement fields. Besides, the Lagrangian multipliers are introduced into the modified hybrid-Trefftz functional for considering the stress resultant constraints at free/SS1 edges. Thus, these elements theoretically possess the abilities to correctly capture the very steep gradients of the resultant distributions in the boundary layers.

These new specialized hybrid-Trefftz plate elements can very efficiently solve the edge effect problem with high accuracy, even when distorted meshes are used. Moreover, because these elements’ construction procedures contain only boundary integrals, the computation expense for accurately integrating the exponential trial functions can be considerably saved.

This work presents an alternative novel idea for using the FEM to more effectively handle the local stress problems by incorporating the use of the analytical trial functions.

The purpose of this paper is to reshape a fast-jet electronics pod’s external geometry to ensure compliance with aircraft pylon load limits across its carriage envelope while adhering to onboard system constraints and fitment specifications.

Initial geometric layout determination used empirical methods. Performance approximation on the aircraft with added fairings and stabilising fin configurations was conducted using a panel code. Verification of loads was done using a full steady Reynolds-averaged Navier–Stokes solver, validated against published wind tunnel test data. Acceptable load envelope for the aircraft pylon was defined using two already-certified stores with known flight envelopes.

Re-lofting the pod’s geometry enabled meeting all geometric and pylon load constraints. However, due to the pod's large size, re-lofting alone was not adequate to respect aircraft/pylon load limitations. A flight restriction was imposed on the aircraft’s roll rate to reduce yaw and roll moments within allowable limits.

The geometry of an electronics pod was redesigned to maximise the permissible flight envelope on its carriage aircraft while respecting the safe carriage load limits determined for its store pylon. Aircraft carriage load constraints must be determined upfront when considering the design of fast-jet electronic pods.

A process for determining the unknown load constraints of a carriage aircraft by analogy is presented, along with the process of tailoring the geometry of an electronics pod to respect aerodynamic load and geometric constraints.

This paper aims to investigate the aerodynamic characteristics as well as static stability of wing-in-ground effect aircraft. The effect of geometrical characteristics, namely, twist angle, dihedral angle, sweep angle and taper ratio are examined.

A three-dimensional computational fluid dynamic code is developed to investigate the aerodynamic characteristics of the effect. The turbulent model is utilized for characterization of flow over wing surface.

The numerical results show that the maximum change of the drag coefficient depends on the angle of attack, twist angle and ground clearance, in a decreasing order. Also, it is found that the lift coefficient increases as the ground clearance, twist angle and dihedral angle decrease. On the other hand, the sweep angle does not have a significant effect on the lift coefficient for the considered wing section and Reynolds number. Also, as the aerodynamic characteristics increase, the taper ratio befits in trailing state.

To design an aircraft, the effect of each design parameter needs to be estimated. For this purpose, the sensitivity analysis is used. In this paper, the influence of all parameter against each other including ground clearance, angle of attack, twist angle, dihedral angle and sweep angle for the NACA 6409 are investigated.

As a summary, the contribution of this paper is to predict the aerodynamic performance for the cruise condition. In this study, the sensitivity of the design parameter on aerodynamic performance can be estimated and the effect of geometrical characteristics has been investigated in detail. Also, the best lift to drag coefficient for the NACA 6409 wing section specifies and two types of taper ratios in ground effect are compared.

Advancements in aerospace technologies, which rely on unsteady fluid dynamics, are being hindered by a lack of easy to use, computationally efficient unsteady computational fluid dynamics (CFD) software. Existing CFD platforms are capable of handling unsteady flapping, but the time, money and expertise required to run even a basic flapping simulation make design iteration and optimization prohibitively expensive for the average researcher.

In the present paper, a remedy to model the effects of viscosity is introduced to the original vortex method, in which the pitching moment amplitude grew over time for simulations involving multiple flapping cycles. The new approach described herein lumps far-field wake vortices to mimic the vortex decay, which is shown to improve the accuracy of the solution while keeping the pitching moment amplitude under control, especially for simulations involving many flapping cycles.

In addition to improving the accuracy of the solution, the new method greatly reduces the computation time for simulations involving many flapping cycles. The solution of the original vortex method and the new method are compared to published Navier–Stokes solver data and show very good agreement.

By utilizing a novel unsteady vortex method, which has been designed specifically to handle the highly unsteady flapping wing problems, it has been shown that the time to compute a solution is reduced by several orders of magnitude (Denda et al., 2016). Despite the success of the vortex method, especially for a small number of flapping cycles, the solution deteriorates as the number of flapping cycles increases due to the inherent lack of viscosity in the vortex method.

The objective of the present study is to investigate numerically the effects of thermal and buoyancy forces on both upward flow (UF) and downward flow (DF) of air in a vertical parallel‐plates channel. The plates are wetted by a thin liquid water film and maintained at a constant temperature lower than that of the air entering the channel.

The solution of the elliptical PDE modeling the flow field is based on the finite volume method.

Results show that buoyancy forces have an important effect on heat and mass transfers. Cases with evaporation and condensation have been investigated for both UF and DF. It has been established that the heat transfer associated with these phase changes (i.e. latent heat transfer) may be more or less important compared with sensible heat transfer. The importance of these transfers depends on the temperature and humidity conditions. On the other hand, flow reversal has been predicted for an UF with a relatively high temperature difference between the incoming air and the walls.

Contrary to most studies in channel heat and mass transfer with phase change, the mathematical model considers the full elliptical Navier‐Stokes equations. This allows one to compute situations of flow reversal.