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The performance of an industrial air amplifier is assessed through experimental and numerical characterization, with a focus on examining the influence of various operating conditions (isolated, “blowing,” and “suction” modes) and direct geometric scaling of the device within the specified range of the injection gap (δ) and the inlet pressure characteristic values. The findings underscore the presence of a linear trend of the entrained mass flow and a nonlinear decay of the amplification factor, both with notable sensitivity to the gap width. Numerical RANS simulations validate the experimental data, characterize the asymmetric flow downstream from the device, and facilitate the exploration of more complex scenarios. In this regard, scaling the device’s dimensions reveals an optimal aspect ratio between the minimum diameter (Dm) and δ to maximize the entrained mass flow. This research provides valuable insights into the behavior of air amplifiers, offering guidance for their design and application across various industrial contexts.

There are many important aerodynamic phenomena on vertical axis wind turbines. Particularly relevant is flow separation that causes structural fatigue and adversely affects self-starting characteristics of VAWTs at lower wind velocities. Passive flow-control devices on VAWTs rotor blades have been commonly studied to mitigate these adverse characteristics. While there are many studies on this topic, the dynamic stall characteristics of VAWTs rotor blades equipped with Gurney flaps (GFs) and vortex generators (VGs) are not completely clear. It is accordingly the goal of the present study to investigate the aerodynamic performance of the NACA 0021 airfoils in the angle of attack range from 0º to 360º. The experiments were performed in a wind tunnel using rotor blade equipped with VG and GF devices. Experimental results encompass aerodynamic force and moment coefficients of the NACA 0021 airfoil equipped with VGs and GFs. These devices proved to improve the aerodynamic characteristics of the NACA 0021 airfoil and to reduce the adverse effects of flow separation.

Growth in application fields of unmanned aerial vehicles (UAVs) and an increase in their total number are followed by higher and higher expectations imposed on improvements in UAV propulsion and energy management systems. Most commercial vertical takeoff and landing (VTOL) UAVs employ a constant pitch propeller that forces a mission execution tradeoff in the majority of cases. An alternative solution, presented here, consists of the use of a variable pitch propeller. The paper summarizes experimental measurements of the propulsion system equipped with an innovative variable pitch rotor. The investigations incorporated characteristics of the rotor for no wind conditions and a new approach to optimize pitch settings in hover flight as a function of UAV weight and energy consumption. As UAV battery capacity is always limited, efficient energy management is the only way to increase UAV mission performance. The study shows that use of a variable pitch propeller can increase the maximal takeoff weight of the aircraft and improve power efficiency in hover, especially if load varies for different missions. The maximal thrust measured was 31% higher with respect to the original blade settings. The coefficient of thrust during hover showed an increase of 2.6% up to 7.5% for various pitch angles with respect to the original fixed propeller.

This paper introduces a generic model for the study of aerodynamic behaviour relevant to fifth-generation high-performance aircraft. The model design is presented, outlining simplifications made to retain the key features of modern high-performance vehicles while ensuring a manufacturable geometry. Subsonic wind tunnel tests were performed with force and moment balance measurements used to develop a database of experimental validation data for the platform at a freestream velocity of 20 m/s. Numerical simulations are also presented and validated by the experiments and further employed to ensure the vortex behaviour is consistent with contemporary high-performance platforms. A sensitivity study of the computational predictions from the turbulence modelling approach is also presented. This geometry is the first in a suite of representative aircraft geometries (the Sydney Standard Aerodynamic Models), in which all geometries, computational models, and experimental data are made openly available to the research community (accessible via this link: https://zenodo.org/communities/ssam_gen5/) to serve as validation test cases and promote best practices in aerodynamic modelling.

This paper describes the experimental study of surface pressure over a supercritical airfoil which was oscillated in pure pitching, pure plunging and combined pitch-plunge motions at the Reynolds number of 8.76*105. While the surface pressure distribution is of significant importance in stability and performance of an airfoil, not sufficient information is available on the pressure distribution in dynamic stall. The experiments were conducted in a closed-loop wind tunnel utilizing pressure transducers array. The motions were designed to maintain constant reduced frequency, Strouhal number and phase difference. Three different regions were assumed to represent the pressure distribution over the airfoil. The results showed that LEV formed on the upper surface manifested different behavior. In the attached flow region the LEV grew and shrunk over the upper surface but in the light stall region the LEV spilled on the airfoil while a small partial LEV remained at the leading edge. In the deep stall region the LEV spilled entirely and the flow was fully separated. The formation of Laminar Separation Bubbles and suction peaks were also reported in low angles of attack. Besides, the pitching moment Damping Factor was studied to determine the level of airfoil stall flutter stability. For lower amplitudes of pitching motion the airfoil seemed to be stable except where deep stall occurred. However for high amplitudes the airfoil had a tendency to enter the stall flutter. Nevertheless, forcing the airfoil to undergo a combined motion improved the stability condition in all cases.

Despite the exceptional aerodynamic performance of delta wing-body configurations, the phenomenon of vortex bursting and its influencing factors remain inadequately understood. This research investigates the effects of Mach number, Reynolds number, and body upwash on the angle of attack and location of vortex bursting over a delta wing-body planform within subsonic and transonic flight regimes. Experimental and computational methodologies were employed. Aerodynamic forces and moments were measured using a six-component internal balance in the ASSEF Transonic Wind Tunnel (ATWT), while surface static pressure measurements were acquired with electronically scanned pressure (ESP) modules. The experimental results aligned with previous studies, validating the experimental setup. Numerical simulations were corroborated by experimental surface pressure, force, and moment data. The study initially examines the burst angle of attack, revealing a rise from 12° to 13° as the Mach number increases from 0.3 to 0.9. However, at Mach 0.95, vortex bursting did not occur up to an angle of attack of 15°, indicating a threshold in the transonic regime beyond which vortex bursting is absent at moderate angles of attack. Furthermore, a decrease in Reynolds number leads to a lower burst angle of attack, and the presence of a body significantly reduces this angle compared to wing-only configurations. The study also explores the burst position, observing a nonlinear shift with increasing Mach number. A threshold of 14° for the angle of attack was identified, where further increases in Reynolds number alter the burst position both longitudinally and laterally. The newly identified thresholds provide valuable insights for optimizing the design and improving the understanding of vortex bursting in similar configurations.

Recent developments in electrical Vertical Take-off and Landing (eVTOL) vehicles show the need for a better understanding of transient aero-mechanical propeller loads for non-axial inflow conditions. The variety of vehicle configurations conceptualized with different propellers in terms of blade geometry, number of blades, and their general integration concept results in aerodynamic loads on the propellers which are different from those on conventional fixed-wing aircraft propellers or helicopter rotors. Such varying aerodynamic loads have to be considered in the vehicle design as a whole and also in the detailed design of their respective electric propulsion systems. Therefore, an experimental approach is conducted on two different propeller blade geometries and a varying number of blades with the objective to explore the characteristics at non-axial inflow conditions. Experimental data are compared with calculated results of a low-fidelity Blade Element Momentum Theory (BEMT) approach. Average thrust and side force coefficients are shown to increase with inflow angle, and this trend is captured by the implemented numerical method. Measured thrust and in-plane forces are shown to oscillate at the blade passing frequency and its harmonics, with higher amplitudes at higher angles of inflow or lower number of blades.

This paper describes some insights on applicability of a Selective Laser Melting and Direct Metal Laser Sintering technology-manufactured turbine blade models for aerodynamic tests in a wind tunnel. The principal idea behind this research was to assess the possibilities of using ‘raw’ DLMS printed turbine blade models for gas-flow experiments. The actual blade, manufactured using the DLMS technology, is assessed in terms of surface quality (roughness), geometrical shape and size (outline), quality of counterbores and quality of small diameter holes. The results are evaluated for the experimental aerodynamics standpoint. This field of application imposes requirements that have not yet been described in the literature. The experimental outcomes prove the surface quality does not suffice to conduct quantitative experiments. The holes that are necessary for pressure measurements in wind tunnel experiments cannot be reduced below 1 mm in diameter. The dimensional discrepancies are on the level beyond acceptable. Additionally, the problem of ‘reversed tolerance’, with the material building up and distorting the design, is visible in elements printed with the DLMS technology. The results indicate the necessity of post-machining of the printed elements prior their experimental usage, as their features in the ‘as fabricated’ state significantly disturb the flow conditions.

Considered are interactive relationships between a normal shock wave and the downstream shock wave leg of the associated lambda foot, as well as between a normal shock wave and time-varying static pressure as measured along the bottom surface of the test section. Such relationships are investigated as they vary with two different magnitudes of inlet unsteady Mach wave intensity and are characterized using shadowgraph flow visualization data, as well as power spectral density, magnitude-squared coherence, and time lag data. Employed for the investigation is a specialty test section with an inlet Mach number of 1.54, as utilized within a transonic/supersonic wind tunnel. The resulting data provide evidence of distinct interactions over a wide range of frequencies between the normal shock wave and the downstream shock wave leg of the lambda foot for low inlet unsteady Mach wave intensity. Note that these are not present in the same form and over the same ranges of frequency with high inlet unsteady Mach wave intensity. These differences are partially due to the location where flow events originate. The most significant sources of flow unsteadiness within the present investigation are mostly associated with the normal and oblique shock waves (with low inlet unsteady Mach wave intensity), and mostly with inlet flow disturbances from unsteady Mach waves (with high inlet unsteady Mach wave intensity). The present experimental results additionally evidence important connections between the normal shock wave and unsteady flow events within lower portions of the lambda foot, especially near the adjacent boundary layer separation region.

The flow structure and unsteady evolution characteristics of dual impinging jets represent a flow problem of significant engineering importance in the aerospace field. Currently, there is a lack of systematic research on the unsteady characteristics and the underlying mechanisms of flow structure evolution in dual impinging jets across different velocity regimes. This study investigates a dual impinging jet configuration with a nozzle pressure ratio ranging from 1.52 to 2.77, an impingement spacing of 5d (where d is the nozzle exit diameter), and an inter-nozzle spacing of 10.42d. By employing Particle Image Velocimetry and Proper Orthogonal Decomposition, the evolution of the flow field structure from subsonic to supersonic conditions is systematically analyzed. The results demonstrate that the fountain motion is composed of an anti-symmetric oscillatory mode, a symmetric breathing mode, and an intermittent transport mode. The upper confinement plate obstructs the fountain motion to some extent, inducing unsteady oscillation modes. An increase in jet velocity enhances the upwash momentum of the fountain and raises the characteristic frequencies of its dynamic structures. This research elucidates the influence of jet velocity variation on the flow field structure, providing a theoretical basis for formulating flow control strategies in related engineering applications.