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The purpose of this paper is to present a fast, economical and practical method for mathematical modeling of aerodynamic characteristics of rectangular wing in ground (WIG) effect.

Reynolds averaged Navier–Stokes (RANS) equations were converted to Bernoulli equation by reasonable assumptions. Also, Helmbold’s equation has been developed for calculation of the slope of wing lift coefficient in ground effect by defining equivalent aspect ratio (ARe). Comparison of present work results against the experimental results has shown good agreement.

A practical mathematical modeling with lower computational time and higher accuracy was presented for calculating aerodynamic characteristics of rectangular WIG effect. The relative error between the present work results and the experimental results was less than 8 per cent. Also, the accuracy of the proposed method was checked by comparing with the numerical methods. The comparison showed fairly good accuracy.

Aerodynamic surfaces in ground effect were used for reducing wetted surface and increasing speed in high-speed marine and novel aeronautical vehicles.

The proposed method is useful for investigation of aerodynamic performance of WIG vehicles and racing boats with aerodynamic surfaces in ground effect.

The proposed method has reduced the computational time significantly as compared to numerical simulation that allows conceptual design of the WIG crafts and is also economical.

The nose length is the key design parameter affecting the aerodynamic performance of high-speed maglev train, and the horizontal profile has a significant impact on the aerodynamic lift of the leading and trailing cars Hence, the study analyzes aerodynamic parameters with multi-objective optimization design.

The nose of normal temperature and normal conduction high-speed maglev train is divided into streamlined part and equipment cabin according to its geometric characteristics. Then the modified vehicle modeling function (VMF) parameterization method and surface discretization method are adopted for the parametric design of the nose. For the 12 key design parameters extracted, combined with computational fluid dynamics (CFD), support vector machine (SVR) model and multi-objective particle swarm optimization (MPSO) algorithm, the multi-objective aerodynamic optimization design of high-speed maglev train nose and the sensitivity analysis of design parameters are carried out with aerodynamic drag coefficient of the whole vehicle and the aerodynamic lift coefficient of the trailing car as the optimization objectives and the aerodynamic lift coefficient of the leading car as the constraint. The engineering improvement and wind tunnel test verification of the optimized shape are done.

Results show that the parametric design method can use less design parameters to describe the nose shape of high-speed maglev train. The prediction accuracy of the SVR model with the reduced amount of calculation and improved optimization efficiency meets the design requirements.

Compared with the original shape, the aerodynamic drag coefficient of the whole vehicle is reduced by 19.2%, and the aerodynamic lift coefficients of the leading and trailing cars are reduced by 24.8 and 51.3%, respectively, after adopting the optimized shape modified according to engineering design requirements.

In this paper, experimental and numerical results of a lambda wing have been compared. The purpose of this paper is to study the behaviour of lambda wings using a CFD tool and to consider different numerical models to obtain the most accurate results. As far as the consideration of numerical methods is concerned, the main focus is on the evaluation of computational methods for an accurate prediction of contingent leading edge vortices’ path and the flow separation occurring because of the burst of these vortices on the wing.

Experimental tests are performed in a closed-circuit wind tunnel at the Reynolds number of 6 × 105 and angles of attack (AOA) ranging from 0 to 10 degrees. Investigated turbulence models in this study are Reynolds Averaged Navior–Stokes (RANS) models in a steady state. To compare the accuracy of the turbulence models with respect to experimental results, sensitivity study of these models has been plotted in bar charts.

The results illustrate that the leading edge vortex on this lambda wing is unstable and disappears soon. The effect of this disappearance is obvious by an increase in local drag coefficient in the junction of inner and outer wings. Streamlines on the upper surface of the wing show that at AOA higher than 8 degrees, the absence of an intense leading edge vortex leads to a local flow separation on the outer wing and a reverse in the flow.

Results obtained from the behaviour study of transition (TSS) turbulence model are more compatible with experimental findings. This model predicts the drag coefficient of the wing with the highest accuracy. Of all considered turbulence models, the Spalart model was not able to accurately predict the non-linearity of drag and pitching moment coefficients. Except for the TSS turbulence model, all other models are unable to predict the aerodynamic coefficients corresponding to AOA higher than 10 degrees.

The presented results in this paper include lift, drag and pitching moment coefficients in various AOA and also the distribution of aerodynamic coefficients along the span.

The presented results include lift, drag and pitching moment coefficients in various AOA and also aerodynamic coefficients distribution along the span.

The purpose of this study is investigate the transient aerodynamic characteristics of high-speed vehicle with body roll motion under crosswind condition to improve aerodynamic stability.

An overset mesh was used to simulate the rolling motion of the vehicle body. A wind tunnel experiment was conducted to validate the numerical method.

The results revealed that the vehicle’s aerodynamic characteristics changed periodically with the body’s periodic motion. In the absence of crosswind, the pressure distribution on the left and right sides of the vehicle body was symmetrical, and the speed streamline flowed to the rear of the vehicle in an orderly manner. The maximum aerodynamic lift observed in the transient simulation was −0.089, which is approximately 0.70 times that of the quasi-static simulation experiment. In addition, the maximum aerodynamic side force observed in the transient simulation was 0.654, which is approximately 1.25 times that of the quasi-static simulation experiment.

The aerodynamic load varies periodically with the vehicle body’s cyclic motion. However, the extreme values of the aerodynamic load do not occur when the vehicle body is at its highest or lowest position. This phenomenon is primarily attributed to the mutual interference of airflow viscosity and the hysteresis effect in the flow field, leading to the formation of a substantial vortex near the wheel. Consequently, the aerodynamic coefficient at each horizontal position becomes inconsistent during the periodic rolling of the vehicle body.

This paper aims to present the results of aerodynamic calculation of impact the main rotor on the fuselage and the tail of a light gyroplane. This kind of vehicle is a type of rotorcraft which uses a non-powered rotor in autorotation to develop lift and engine-powered propeller to provide the thrust. Both of them disturb the flow around the gyroplane body (gyroplane fuselage and tail) and influence on its static stability. The main goal of the presented research was to find the magnitude of this influence. To measure this effect, the main stability derivatives changes of gyroplane body were investigated.

The CFD analysis of the complete gyroplane was made. Computation was performed for the model of gyroplane which was equipped with the two sub-models of the main rotor and the engine-powered propeller. Both of them were modelled as the actuator discs. This method allows to compute the aerodynamic impact of rotating components on the gyroplane body. All aerodynamic analysis was made by the MGAERO software. The numerical code of the software bases on the Euler flow model. Next, the resulting aerodynamic coefficients were used to calculate the most important stability derivatives of the gyroplane body.

The result obtained by computation presents the change in the most important aerodynamic coefficients and stability derivatives of the gyroplane body caused by the impact of its main rotor. Moreover, the result includes the change of the aerodynamic coefficients and stability derivatives caused by change of the main rotor configuration (change of rotation rate and angle of incidence) and change of the flight condition (gyroplane angle of attack sideslip angle and flight speed).

Analysis of the main rotor impact will be very useful for evaluation of dynamic stability of the light gyroplane. Moreover, the results will be helpful to design the horizontal and vertical tail for the light gyroplane.

This paper presents the method of the numerical analysis of the static stability of the light gyroplane’s body. The results of analysis present the impact of disturbance generated by the rotating main rotor on the static stability of the gyroplane body. Moreover, the impact of the main rotor configuration change and the flight condition change on the static stability were investigated too. The evaluation of the gyroplane’s body static stability was made by the stability derivatives. The methodology and obtained result will be very useful for analysis of the dynamic stability of the light gyroplanes. Moreover, the results will be helpful during design the main components of the gyroplane like vertical and horizontal tail.

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.

The purpose of this paper is to investigate the aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack using global analytic nonlinear aerodynamic model.

Aeroelastic behavior of a supersonic flight vehicle flying at moderate angles of attack is considered, using nonlinear aerodynamics and linear elastodynamics and structural models. Normal force distribution coefficient over the length of the vehicle and pitching moment coefficient are the main aerodynamic parameters used in the aeroelastic modeling. It is very important to have closed form analytical relations for these coefficients in the model. They are generated using global nonlinear multivariate orthogonal modeling functions in this work. Angle of attack and length of the vehicle are selected as independent variables in the first step. Local variation of angle of attack is applied to the analytical model and due to its variation along the body of the vehicle, equations of motion are finalized. Mach number is added to the independent variables to investigate its role on instability of the vehicle and the modified model is compared with the previous one in the next step. Thrust effect on the aeroelastic stability of the vehicle is analyzed at final stage.

It is shown that for the vehicles having simple configurations and low length to diameter ratios flying at low angles of attack, assuming normal force distribution coefficient linear relative to α is reasonable. It is concluded that vehicle's velocity and thrust has not great effect on its divergence dynamic pressure.

Based on the constructed model, a simulation code is generated to investigate the aeroelastic behavior of the vehicle. The resultant code is verified by investigating the static aeroelastic stability margin of the vehicle presented in the references. Mach number effect on the aeroelastic behavior of the vehicle is considered using modified aerodynamic model and is compared with the results. Data base for identifying aerodynamic coefficients is conducted using CFD code.

WE concluded Part II of this series with the remark that a different outlook is needed for problems of control surface flutter than for those of wing flutter. There are two reasons for this. Wing flutter must be investigated carefully early on in the design of an aircraft so as to provide a safe aircraft without a severe weight penalty, whereas the weight penalty of avoiding control surface flutter is usually small, although not negligible, and modifications can often be made at short notice, so it is important to make a full investigation as late as possible before flight when all the data are available in a reliable form. The second reason is that with wing flutter, as with aileron reversal and divergence, it is usual to think of safety margins in terms of forward speed or possibly wing torsional stiffness; with control surface flutter, on the other hand, quite different types of safety factor become the rule.

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.

The purpose of this paper is to evaluate the accuracy of linear and quasi-steady aerodynamic models of aircraft aerodynamic models when a small unmanned aerial system flies in the presence of strong wind and gust at a high angle of attack and a high sideslip angle.

Compatibility analysis were done to improve the quality of recorded flight test data. A robust method called fuzzy logic modeling is used to set up the aerodynamic models. The reduced frequency is used to represent the unsteadiness of the flow field according to Theodorsen’s theory. The work done by the aerodynamic moments on the motions is used as the criteria of stability.

In portions of flight, aircraft’s stability and control derivatives were unstable and nonlinear functions of airflow angles and angular rates. The roll angle had an important effect on unsteadiness of directional oscillatory damping derivatives. The pilot-induced oscillation and wing rock possibilities were investigated and dismissed so that the lateral directional oscillatory motion was classified as a nonlinear Dutch roll oscillation. Major modeling enhancements or real-time parameter identification are required for the control of a small unmanned aerial system in off-nominal conditions. The robustness tests of all-weather autopilot systems must be done with consideration of sign change.

Oscillatory damping derivatives were reconstructed using flight test data and the inadequacy of engineering level software in predicting this type of instability observed and demonstrated for a flight in the presence of wind shear and external disturbances.