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Aerodynamics of paragliders is very complicated aeroelastic phenomena. The purpose of this work is to quantify the amount of aerodynamic drag related to the flexible nature of a paraglider wing.

The laboratory testing on scaled models can be very difficult because of problems in the elastic similitude of such a structure. Testing of full-scale models in a large facility with a large full-scale test section is very expensive. The degradation of aerodynamic characteristics is evaluated from flight tests of the paraglider speed polar. All aspects of the identification such as pilot and suspension lines drag and aerodynamics of spanwise chambered wings are discussed. The drag of a pilot in a harness was estimated by means of wind tunnel testing, computational fluid dynamics (CFD) solver was used to estimating smooth wing lift and drag characteristics.

The drag related to the flexible nature of the modern paraglider wing is within the range of 4-30 per cent of the total aerodynamic drag depending on the flight speed. From the results, it is evident that considering only the cell opening effect is sufficient at a low-speed flight. The stagnation point moves forwards towards the nose during the high speed flight. This causes more pronounced deformation of the leading edge and thus increased drag.

This paper deals with a detailed analysis of specific paraglider wing. Although the results are limited to the specific geometry, the findings help in the better understanding of the paraglider aerodynamics generally.

The data obtained in this paper are not affected by any scaling problems. There are only few experimental results in the field of paragliders on scaled models. Those results were made on simplified models at very low Reynolds number. The aerodynamic drag characteristics of the pilot in the harness with variable angles of incidence and Reynolds numbers have not yet been published.

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.

The purpose of this paper is to examine the aerodynamic effects of rear spoiler geometry on a sports car. Today, due to economical, safety and even environmental concerns, vehicle aerodynamics play a much more significant role in design considerations and rear spoilers play a major role in this area.

A 2-D vehicle geometry of a race car is created and solved using the computational fluid dynamics (CFD) solver FLUENT version 6.3. The aerodynamic effects are analyzed under various vehicle speeds with and without a rear spoiler. The main results are compared to a wind tunnel experiment conducted with 1/18 replica of a Nascar.

By the CFD analysis, the drag coefficient without the spoiler is calculated to be 0.31. When the spoiler is added to the geometry, the drag coefficient increases to 0.36. The computational results with the spoiler are compared with the experimental data, and a good agreement is obtained within a 5.8 percent error band. The uncertainty associated with the experimental results of the drag coefficient is calculated to be 6.1 percent for the wind tunnel testing. The sources of discrepancies between the experimental and numerical results are identified and potential improvements on the model and experiments are provided in the paper. Furthermore, in the CFD model, it is found that the addition of the spoiler caused a decrease in the lift coefficient from 0.26 to 0.05.

This paper examines the effects of rear spoiler geometry on vehicle aerodynamic drag by comparing the CFD analysis with wind tunnel experimentation and conducting an uncertainty analysis to assess the reliability of the obtained results.

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 to explore the effects of aerospike and hemispherical aerodisks on flow characteristics and drag reduction in supersonic flow over a blunt body. Specifically, the study aims to analyze the impact of varying the length of the cylindrical rod in the aerospike (ranging from 0.5 to 2.0 times the diameter of the blunt body) and the diameter of the hemispherical disk (ranging from 0.25 to 0.75 times the blunt body diameter). CFD simulations were conducted at a supersonic Mach number of 2 and a Reynolds number of 2.79 × 10⁶.

ICEM CFD and ANSYS CFX solver were used to generate the three-dimensional flow along with its structures. The flow structure and drag coefficient were computed using Reynolds-averaged Navier–Stokes equation model. The drag reduction mechanism was also explained using the idea of dividing streamline and density contour. The performance of the aero spike length and the effect of aero disk size on the drag are investigated.

The separating shock is located in front of the blunt body, forming an effective conical shape that reduces the pressure drag acting on the blunt body. It was observed that extending the length of the spike beyond a specific critical point did not impact the flow field characteristics and had no further influence on the enhanced performance. The optimal combination of disk and spike length was determined, resulting in a substantial reduction in drag through the introduction of the aerospike and disk.

To predict the accurate results of drag and to reduce the simulation time, a hexa grid with finer mesh structure was adopted in the simulation.

The blunt nose structures are primarily employed in the design of rockets, missiles, and re-entry capsules to withstand higher aerodynamic loads and aerodynamic heating.

For the optimized size of the aero spike, aero disk is also optimized to use the benefits of both.

The purpose of this paper is to investigate the optimal average drag coefficient of the Ahmed body for mixed platoon driving under crosswind and no crosswind conditions using the response surface optimization method. This study has extraordinary implications for the planning of future intelligent transportation.

First, the single vehicle and vehicle platoon models are validated. Second, the configuration with the lowest average drag coefficient under the two conditions is obtained by response surface optimization. At the same time, the aerodynamic characteristics of the mixed platoon driving under different conditions are also analyzed.

The configuration with the lowest average drag coefficient under no crosswind conditions is 0.3 L for longitudinal spacing and 0.8 W for lateral spacing, with an average drag coefficient of 0.1931. The configuration with the lowest average drag coefficient under crosswind conditions is 10° for yaw angle, 0.25 L for longitudinal spacing, and 0.8 W for lateral spacing, with an average drag coefficient of 0.2251. Compared to the single vehicle, the average drag coefficients for the two conditions are reduced by 25.1% and 41.3%, respectively.

This paper investigates the lowest average drag coefficient for mixed platoon driving under no crosswind and crosswind conditions using a response surface optimization method. The computational fluid dynamics (CFD) results of single vehicle and vehicle platoon are compared and verified with the experimental results to ensure the reliability of this study. The research results provide theoretical reference and guidance for the planning of intelligent transportation.

This paper aims to compare three closed non-planar wing configurations with a reference conventional wing-plus-horizontal tail aircraft, considering structural aspects, weights and aerodynamic characteristics, as well as operational issues, such as cruise performance.

A vortex lattice code is used and coupled with an in-house code for structural beam calculation subroutine to evaluate the configurations as a function of the four main parameters identified in the study.

The study concludes that the non-planar wing configurations have better performances than a conventional aircraft. Moreover, the joined-wing configuration seems to be better than the others, including the box-wing configuration, achieving an increase of 17 per cent in the range for maximum payload compared to the reference aircraft and a 3 per cent reduction of maximum take-off weight.

In the study, characteristic tools for a conceptual design are used, and, thus, absolute results should be considered with caution. Nonetheless, as all the cases are studied in the same way, there is a good precision in comparative or relative results.

The work shows that the non-planar wing configurations can be used as an alternative to the conventional aircraft to meet the objectives for the future of the aviation industry.

Non-planar wing configurations are able to reduce fuel consumption. Their use could lead to reductions in pollutant emissions and the impact on the environment of commercial aviation.

This study considers aerodynamic and structural aspects at the same time, as well as several non-planar wing configurations, making possible to obtain a more realistic comparison between them.

Boundary layer ingestion (BLI) is one of the probable noteworthy features of distributed propulsion configuration (DPC). Because of BLI, strong coupling effects are generated between the aerodynamics and propulsion system of aircraft, leading to the specific lift and drag aerodynamic characteristics. This paper aims to propose a model-based comprehensive analysis method to investigate this unique aerodynamic.

To investigate this unique aerodynamics, a model-based comprehensive analysis method is proposed. This method uses a detailed mathematical model of the distributed propulsion system to provide the essential boundary conditions and guarantee the accuracy of calculation results. Then a synthetic three-dimensional computational model is developed to analyze the effects of BLI on the lift and drag aerodynamic characteristics.

Subsequently, detailed computational analyses are conducted at different flight states, and the regularities under various flight altitudes and velocities are revealed. Computational results demonstrate that BLI can improve the lift to drag ratio evidently and enable a great performance potentiality.

The general analysis method and useful regularities have reference value to DPC aircraft and other similar aircrafts.

This paper proposed a DPS model-based comprehensive analysis method of BLI benefit on aerodynamics for DPC aircraft, and the unique aerodynamics of this new configuration under various flight altitudes and velocities was revealed.

The purpose of this paper is to optimize the downrange for hypersonic boost‐glide (HBG) missile under near‐real condition, and to validate the suitability of proposed wall cooling materials.

The trajectory optimization problem is characterized by a boost phase followed by a glide phase. A multi‐phase trajectory optimization tool is adopted to optimize the downrange. The associated optimal control problem has been solved by selecting a direct shooting method. The dynamics has been transcribed to a set of nonlinear constraints and the arising nonlinear programming problem has been solved through a sequential quadratic programming solver. An aerothermodynamics analysis method is introduced to calculate the aerodynamic heating at nose, leading edge, and ventral centerline regions.

HBG missile is suitable for long‐range attack, and the optimal trajectory solved is a novel boost‐glide‐skip trajectory, i.e. boost firstly, glide secondly, and skip at last. The proposed wall materials are valid.

This paper provides further study on the methods of trajectory design and aerothermodynamics analysis for HBG missile.

The purpose of this paper is to present a novel approach based on the differential search (DS) algorithm integrated with the adaptive network-based fuzzy inference system (ANFIS) for unmanned aerial vehicle (UAV) winglet design.

The winglet design of UAV, which was produced at Faculty of Aeronautics and Astronautics in Erciyes University, was redesigned using artificial intelligence techniques. This approach proposed for winglet redesign is based on the integration of ANFIS into the DS algorithm. For this purpose, the cant angle (c), the twist angle (t) and taper ratio (λ) of winglet are selected as input parameters; the maximum value of lift/drag ratio (E_max) is selected as the output parameter for ANFIS. For the selected input and output parameters, the optimum ANFIS parameters are determined by the DS algorithm. Then the objective function based on optimum ANFIS structure is integrated with the DS algorithm. With this integration, the input parameters for the E_max value are obtained by the DS algorithm. That is, the DS algorithm is used to obtain both the optimization of the ANFIS structure and the necessary parameters for the winglet design. Thus, the UAV was reshaped and the maximum value of lift/drag ratio was calculated based on new design.

Considerable improvements on the max E are obtained through winglet redesign on morphing UAVs with artificial intelligence techniques.

It takes a long time to obtain the optimum E_max value by the computational fluid dynamics method.

Using artificial intelligence techniques saves time and reduces cost in maximizing E_max value. The simulation results showed that satisfactory E_max values were obtained, and an optimum winglet design was achieved. Thus, the presented method based on ANFIS and DS algorithm is faster and simpler.

The application of artificial intelligence methods could be used in designing more efficient aircrafts.

The study provides a new and efficient method that saves time and reduces cost in redesigning UAV winglets.