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The main purpose of this study was to determine the basic aerodynamic characteristics of the airliner Tu-154M at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°.

Wind tunnel tests of the Tu-154M aircraft model at the scale 1:20 were performed in a low-speed wind tunnel T-3 by using a six-component internal aerodynamic balance. Several model configurations were also investigated.

The results of the presented studies showed that at the wide range of the overcritical angles of attack and sideslip angles, i.e. α = −900° ÷ 900° and β = −900° ÷ 900°, the Tu-154M aircraft flap deflection affected the values of the drag and lift coefficients and generally had no major effect on the values of the side force and pitching moment coefficients.

The model vibration which was the result of flow separation at high angles of attack was the wind tunnel test limitation.

Studies of the airliner aerodynamic characteristics at the wide range of the overcritical angles of attack and sideslip angles allow assessment of the aircraft aerodynamic properties during possible unexpected situations when the passenger aircraft is found to have gone beyond the conventional flight envelope.

There are no social implications of this study to report.

The presented wind tunnel test results of the airliner aerodynamic characteristics at overcritical angles of attack and sideslip angles is an original contribution to the existing not-too-extensive database available in the literature.

The attack angle of stencil printing squeegees with different geometries was analysed using finite element modelling.

A finite element model (FEM) was developed to determine the attack angle during the stencil printing. The material properties of the squeegee were included in the model according to the parameters of steel AISI 4340, and the model was validated by experimental measurements. Two geometric parameters were investigated; two different unloaded angles (45° and 60°) and four overhang sizes of the squeegee (6, 15, 20 and 25 mm).

It was found that the deflection of the blade is nearly homogenous along the length of the squeegee. This implies that the attack angle does not change significantly along the squeegee length. The results showed significant differences between the initial and the attack angle. For example, the angle of the squeegee with 15 mm overhang size and with 60° initial angle decreased by more than 5° for a specific squeegee force of 0.3 N/mm; resulting in an attack angle of 53.4°.

The attack angle during the printing is considerably lower than the initial angle as a result of the printing force. The papers, which were dealing with the numerical modelling of the stencil printing presumed that the squeegees were having their initial angle. This could have led to invalid numerical results. Therefore, we decided to investigate the attack angle during stencil printing for squeegees with different initial geometries to enhance the numerical modelling of stencil printing.

This study aims to numerically investigate the heat transfer and laminar forced and two-phase flow of Water/Cu nanofluid in a rectangular microchannel with oblique ribs with angle of attacks equal to 0-45°. This simulation was conducted in the range of Reynolds numbers of 5-120 in volume fractions of 0, 2 and 4 per cent of solid nanoparticles in three-dimensional space.

This study investigates the effect of the changes of angle of attack of rectangular rib on heat transfer and hydrodynamics of two-phase flow. This study was done in three-dimensional space and simulation was done with finite volume method. SIMPLEC algorithm and second-order discretization of equations were used to increase the accuracy of results. The usage of nanofluid, application of rips with different angles of attacks and using the two-phase mixture method is the distinction of this paper compared with other studies.

The results of this research revealed that the changing angle of attack of ribs is an effective factor in heat transfer enhancement. On the other hand, the existence of rib on the internal surfaces of a microchannel increases friction coefficient. By increasing the volume fraction of nanoparticles, due to the augmentation of fluid density and viscosity, the pressure drop increases significantly. For all of the angle of attacks studied in this paper, the maximum rate of performance evaluation criterion has been obtained in Reynolds number of 30 and the minimum amount of performance evaluation criterion was been obtained in Reynolds numbers of 5 and 120.

Many studies have been done in the field of heat transfer in ribbed microchannel. In this paper, the laminar flow in the ribbed microchannel Water/Cu nanofluid in a rectangular microchannel by using two-phase mixture method is numerically investigated with different volume fractions (0-4 per cent), Reynolds numbers (5-120) and angle of attacks of rectangular rib in the indented microchannel (0-45°).

As to different angles of attack and nonlinear problems caused by high temperatures in coexisting hypersonic aircraft, people mainly rely on fluid software for research but lack analysis of flow mechanisms. Owing to computational difficulties, few people use numerical algorithms to combine them for discussion. Hence, this study aims to make a deep inquiry into the laminar flow and heat transfer of compressible Newtonian fluid in hypersonic aircraft with small attack angles.

In this paper, on the basis of mass, momentum and energy conservation laws, the governing equations of the hypersonic boundary layer are established. Viscosity, specific heat capacity and thermal conductivity are considered nonlinear functions concerning temperature. In virtue of the MacCormack finite difference method, the stationary numerical solutions are solved directly, and the validity of the algorithm is verified.

The results demonstrate that at Mach number 5, compared to the 0° attack angle, the maximum temperature near-wall at the 3° attack angle increases by about 25%. An enjoyable phenomenon is discovered, where the position corresponding to the maximum wall shear force shifts back as the attack angle and Mach number increase. The relationship between the near-wall maximum temperature versus attack angle and Mach number is fitted through numerical calculation results.

Empirical formulas can be used to estimate heat transfer characteristics at small attack angles, which will guide the design of aircraft thermal protection systems.

The high lift devices are effective at high angle of attack to increase the coefficient of lift by increasing the camber. But it affects the low angle of attack aerodynamic performance by increasing the drag. Hence, they have made as a movable device to deploy only at high angles of attack, which increases the design and installation complexities. This study aims to focus on the comparison of aerodynamic efficiency of different conventional leading edge (LE) slat configurations with simple fixed bioinspired slat design.

This research analyzes the effect of LE slat on aerodynamic performance of CLARK Y airfoil at low and high angles of attack. Different geometrical parameters such as slat chord, cutoff, gap, width and depth of LE slat have been considered for the analysis.

It has been found that the LE slat configuration with slat chord 30% of airfoil chord, forward extension 8% of chord, dip 3% of chord and gap 0.75% of chord gives higher aerodynamic efficiency (C_l/Cd) than other LE slat configurations, but it affects the low angles of attack aerodynamic performance with the deployed condition. Hence, this optimum slat configuration is further modified by closing the gap between LE slat and the main airfoil, which is inspired by the marine mammal’s nose. Thus increases the coefficient of lift at high angles of attack due to better acceleration over the airfoil nose and as well enhances the aerodynamic efficiency at low angles of attack.

The two-dimensional computational analysis has been done for different LE slat’s geometrical parameters at low subsonic speed.

This bio-inspired nose design improves aerodynamic performance and increases the structural strength of aircraft wing compared to the conventional LE slat. This fixed design avoids the complex design and installation difficulties of conventional movable slats.

The findings will have significant impact on the fields of aircraft wing and wind turbine designs, which reduces the design and manufacturing complexities.

Different conventional slat configurations have been analyzed and compared with a simple fixed bioinspired slat nose design at low subsonic speed.

This paper aims to achieve an optimum flow separation control over the airfoil using a passive flow control method by introducing a bio-inspired nose near the leading edge of the National Advisory Committee for Aeronautics (NACA) 4 and 6 series airfoil. In addition, to find the optimised leading edge nose design for NACA 4 and 6 series airfoils for flow separation control.

Different bio-inspired noses that are inspired by the cetacean species have been analysed for different NACA 4 and 6 series airfoils. Bio-inspired nose with different nose length, nose depth and nose circle diameter have been analysed on airfoils with different thicknesses, camber and camber locations to understand the aerodynamic flow properties such as vortex formation, flow separation, aerodynamic efficiency and moment.

The porpoise nose design that has a leading edge with depth = 2.25% of chord, length = 0.75% of chord and nose diameter = 2% of chord, delays the flow separation and improves the aerodynamic efficiency. Average increments of 5.5% to 6° in the lift values and decrements in parasitic drag (without affecting the pitching moment) for all the NACA 4 and 6 series airfoils were observed irrespective of airfoil geometry such as different thicknesses, camber and camber location.

The two-dimensional computational analysis is done for different NACA 4 and 6 series airfoils at low subsonic speed.

This design improves aerodynamic performance and increases the structural strength of the aircraft wing compared to other conventional high lift devices and flow control devices. This universal leading edge flow control device can be adapted to aircraft wings incorporated with any NACA 4 and 6 series airfoil.

The results would be of significant interest in the fields of aircraft design and wind turbine design, lowering the cost of energy and air travel for social benefits.

Different bio-inspired nose designs that are inspired by the cetacean species have been analysed for NACA 4 and 6 series airfoils and universal optimum nose design (porpoise airfoil) is found for NACA 4 and 6 series airfoils.

This paper focuses mainly on the experimental and in‐flight spin investigations for an executive light airplane, named I‐23 and built in the Institute of Aviation (Warsaw, Poland). It is a single‐engine, all composite, straight wing, retractable undercarriage, conventional configuration and flight control system airplane. In‐flight spin tests confirmed good rudder and elevator effectiveness for spin recovery in a wide range of positions of the center of gravity. A typical time history of a spin entry and the developed spin and recovery is shown as well.

This paper aims to analyze the influence of the changings in geometrical parameters on the aerodynamic performance of the control canard projectiles.

Because of the mentioned point, the range of projectiles increment has a considerable importance, and the design algorithm of a control canard projectile was first written. Then, were studied the effects of canard geometric parameters such as aspect ratio, taper ratio and deflectable nose on lift to drag coefficient ratio, static margin based on the slender body theory and cross section flow.

The code results show that aspect ratio increment, results in an increase in lift-to-drag ratio of the missile, but increase in canard taper ratio results in increasing of lift-to-drag ratio at 1° angle of attack, while during increasing the canard taper ratio up to 0.67 at 4° angle of attack, lift to drag first reaches to maximum and then decreases. Also, static margin decreases with canard taper ratio and aspect ratio increment. The developed results for this type of missile were compared with same experimental and computational fluid dynamic (CFD) results and appreciated agreement with other results at angles of attack between 0° and 6°.

To design a control canard missile, the effect of each geometric parameter of canard needs to be estimated. For this purpose, the suitable algorithm is used. In this paper, the effects of canard geometric parameters, such as aspect ratio, taper ratio and deflectable nose on lift-to-drag coefficient ratio and static margin, were studied with help of the slender body theory and cross-section flow.

The contribution of this paper is to predict the aerodynamic characteristics for the control canard missile. In this study, the effect of the design parameter on aerodynamic characteristics can be estimated, and the effect of geometrical characteristics has been analyzed with a suitable algorithm. Also, the best lift-to-drag coefficient for the NASA Tandem Control Missile at Mach 1.75 was selected at various angles of attack. The developed results for this type of missile were compared with same experimental and CFD results.

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.

The primary purpose of this paper is to investigate the characteristics of the unsteady flow field in the wake of Eppler‐361 airfoil undergoing harmonic pitch oscillation in both pre‐stall and post‐stall regimes.

Experimental measurements were carried out to study the characteristics of the unsteady flow field within the wake of an airfoil. All of the experiments were conducted in a low‐speed wind tunnel, and the velocity field was measured by a hot‐wire anemometry. The airfoil was given a harmonic pitching motion about its half chord axis at two reduced frequencies of 0.091 and 0.273. All experimental data were taken at the oscillation amplitude of 8°. During the experiments, the mean angle of attack was altered from 2.5 to 10° that this made it possible to study the wake in both pre‐stall and post‐stall regimes.

From the results, it can be concluded that different velocity profiles are formed in the wake at different phase angles. In addition, the hysteresis of the velocity field in the wake is captured between increasing and decreasing incidences. It is also found that the velocity field in the wake is strongly affected by the operating conditions of the airfoil, e.g. mean angle of attack, reduced frequency and instantaneous angle of attack. Huge variations in the profiles of the wake are observed at high instantaneous angles of attack when the mean angle of attack is 10°, i.e. when the airfoil experiences significant oscillations beyond the static stall. It is concluded that this is due to dynamic stall phenomenon.

Findings of the present study give valuable information, which can be used to characterize wakes of micro air vehicles, helicopter's rotor blades, and wind turbine blades. In addition to this, present findings can be used to predict dynamic stall of the above applications.

The paper is the first to investigate the unsteady wake of Eppler‐361 airfoil and to predict the dynamic stall phenomenon of this airfoil.