Previous research on firefighting aircraft which employ dumping water from their lower surface to extinguish flames has shown that there are added lift benefits during dumping maneuvers which do not directly correlate to the exchange in water mass. In this case study, the effects of liquid spray formations on two dimensional aerodynamics are investigated by use of a numerical approach. Studies include chord-wise variation of liquid-jets and variation of spray momentum ratios. Comparisons are also made to previous, single-phase, jet-flap numerical experiments and show favorable agreement. Findings show that increases in lift coefficient and decreases in drag coefficient are observed with the presences of liquid jets located on the lower surface.

Many different active flow control methods are used to manipulate the flow field about aerodynamic surfaces in order to obtain the most desirable aerodynamic performance. Among these techniques, boundary layer suction is one of the most effective techniques used to improve aerodynamic performance of the airfoil. In this study, the configuration of a pure suction jet actuator is optimized over an oscillating NACA0012 airfoil at the Reynolds number of 1.35 × 10 5 to control the dynamic stall behavior. The airfoil was pitched around the quarter-chord location with a sinusoidal motion and the angle of attack was varied between −5 and 25 degrees. Genetic algorithm was implemented as the optimization method. However, since large number of numerical simulations were required for this purpose, an artificial neural network was employed for training a function between the control parameters and the airfoil aerodynamic coefficients. Aerodynamic performance defined as lift-to-drag ratio was chosen as the objective function of the optimization. Location, velocity amplitude, opening length and jet incidence angle were the control parameters of this optimization. It was shown that when the velocity amplitude and opening length were maximum, the airfoil reached its highest performance. Moreover, the aerodynamic characteristics of the airfoil were remarkably improved when the jet incident angle approached to 90 degrees. Placing the suction jet actuator in the range between 3 to 6 percent of the airfoil chord, was found to have the greatest effect on improving the aerodynamic performance. For the optimum configuration, the airfoil separation. It was shown that when the velocity amplitude and opening length were maximum, the airfoil reached its highest performance. Moreover, the aerodynamic characteristics of the airfoil were peaked in the range between 90 to 120 degrees, with 107 having the best performance in our database.

Use of active flow control techniques has become important in flow separation control. Continuous blowing jet is one of the most effective methods that can be used to improve aerodynamic performance of an airfoil. In the present work, different operational parameters of a continuous blowing jet were optimized to improve the aerodynamic performance of an oscillating NACA0012 airfoil. The airfoil underwent a sinusoidal motion about its quarter-chord between −5 and 25 degrees at the Reynolds number of 1.35 × 10 5 . Unsteady Navier-Stokes equations were solved with k-ω SST turbulence model. Due to the time-consuming nature of large number of numerical simulations required during the optimization process, two neural networks were employed to reduce the number of simulations required. The optimization was carried out with the use of a genetic algorithm. The objective function was defined as the lift-to-drag ratio. In these networks, the relationship between the jet operational characteristics and the aerodynamic coefficients were trained. The jet operational parameters that were considered in this study, included jet location (at 1–60 percent of chord length), jet-opening length (0.05 to 0.3 percent of chord length), blowing jet velocity magnitude (0 to 5U ∞ ), and blowing jet incident angle (0 to 180 degrees). Obtained results indicated that jet-opening length and blowing velocity magnitude have a greater effect on the aerodynamic performance when reached their upper values. Concerning the jet location, it was observed that the best jet location was about 2 to 5 percent of the chord Jet angle (θ) was found to results in the best performance when oriented at range 55 to 70 angle. Results indicated a significant improvement of the aerodynamic performance at the optimum blowing jet configuration.

The generation of lift is a fundamental problem in aerodynamics and in general in fluid mechanics. The explanations on how lift is generated are often very incomplete or even not correct. Perhaps the most popular explanation of lift is the one with the Bernoulli equation and with the longer path over an airfoil as compared to the path below the airfoil, assuming the flow arrives at the same time at the trailing edge on both paths. This is an intuitive assumption, but no equation is derived from this assumption. In some explanations the Bernoulli equation is also complemented with Newton’s laws of motion. In other explanations Newton’s law is said to be the only explanation. Other explanations mention the Venturi suction effect to explain the generation of lift. In books of aerodynamics and on the homepage of well-known research institutes the explanations are, although better and partially correct, still very often incomplete. In this contribution the generation of lift is explained in a scientific way based on the conservation principles of mass, momentum and energy and how they have to be applied to close the system of equations in order to explain the generation of lift. The most common incomplete or incorrect explanations of lift are also analysed and it is explained why they are incomplete or wrong. In this work the generation of lift is explained based on the conservation equations. It is shown how and when they apply to the problem of lift generation and how the system of equations has to be closed.

The UAV industry is booming with investments in research and development on improving UAV systems. Current UAV machines are developed according to the quadcopter design which consists of a rotary propulsion system providing lift needed for flight. This design has some flaws; namely safety concerns and noise/vibration production both of which come stem from the rotary propulsion system. As such, a novel propulsion system using slip stream air passed through high performance slot jets is proposed and an analysis of the fluid characteristics is presented in this report. The test section for the experiment is developed using 3D printed ABS plastic airfoils modified with internal cavities where pressurized air is introduced and then expelled through slot jets on the pressure side of the airfoils. Entrainment processes develop in the system through high momentum fluid introduction into a sedentary secondary fluid. Entrainment is governed by pressure gradients and turbulent mixing and so turbulent quantities that measure these processes are extracted and analyzed according to the independent variable’s effects on these quantities. Pitot probe testing extracted one dimensional fluid information and PIV analysis is used to characterize the two-dimensional flow aspects. High slot jet velocities are seen to develop flows dominated by convection pushing momentum mixing downstream reducing the mixing in the system while low slot jet speeds exhibit higher mass fluxes and thrust development. Confinement spacing is seen to cause a decrease in flow velocity and thrust as the spacing is decreased for high speed runs. The most constricted cross sectional runs showed high momentum mixing and developed combined self-similar flow through higher boundary layer interactions and pressures, but this also hurts thrust development by minimizing secondary flows. The Angle of Attack of the assembly proved to be the most important variable. Outward angling showed the influence of coanda effects but also demonstrated the highest bulk fluid flow with turbulence driven momentum mixing. Inward angling created combined fluid flow downstream with high momentum mixing upstream driven by pressure. Minimal mixing is seen when the airfoils are not angled, and high recirculation zones occur along the boundaries. The optimal setup is seen when the airfoils are angled outwards where the highest thrust and bulk fluid movement is developed driven by the turbulent mixing induced by the increasing cross sectional area of the system.

Large Eddy Simulation for active flow control (AFC) by employing synthetic and continuous blowing is done to investigate their effects on resizing separation. The flow around an SD7003 airfoil at Reynolds number of 60,000 and angles of attack of 13° is considered where a widespread separation occurs at post stall. In this work, the Dynamic Smagorinsky model is used as to calculate the turbulent viscosity.

Computational Fluid Dynamics (CFD) models have been regarded as inevitable tools in the design of various aerodynamics applications like formula 1 cars, wind turbines, airfoil design, and many others. Such models rely on applying advanced numerical methods to solve the set of Navier-Stokes equations and other auxiliary models to quantify the aerodynamics performance, such as lift and drag in airfoil applications. However, currently adopted advanced mathematical models that are used to predict airfoil performance, though are essential tools in the advancement of aviation industry, incur large computational costs. As a primary step, an approximate model of an airfoil is built for the sake of reducing computational cost while preserving accuracy. The approximate model is based on creating a simplified algebraic model which is suitable then to perform aerodynamic design optimization rather than solving the highly non-linear set of continuity and momentum equations. In this work, the non-intrusive polynomial chaos expansion (NIPC) technique is employed on a NACA-2412 airfoil. The main objective of this research is to investigate the performance of the NIPC model with different orders on solution accuracy. A methodology is designed by first developing a full CFD model of the NACA-2412 airfoil. An experimental setup of the airfoil was then prepared with the variations of lift and drag coefficients under different operative conditions were reported. The experimental data were used to validate the numerical results generated by the CFD simulations. The surrogate model was then constructed via the NIPC technique. The performance of the surrogate model was then benchmarked with CFD results under different flow and operating conditions. Analysis was also performed to investigate the effect of polynomial order on the solution accuracy. Results showed that for a given data set size, increasing the order of the polynomial would deteriorate the solution accuracy. At a certain order k, the analytical solution starts diverging especially for values of angle of attack close to −60 and 60°.

The flow around, and in the wake of, pitching airfoils has received renewed interest due to its potential for thrust production at low Reynolds numbers. Past work has centered on the flow fields generated by symmetric pitching of the airfoil. Studies investigating the effects of asymmetric motion are more limited. This work focuses on the wake patterns developed due to asymmetric pitching. Particle Image Velocimetry (PIV) is used to quantify the flow field around a NACA0012 airfoil undergoing small amplitude, high frequency asymmetric pitching. The airfoil is pitched about the quarter chord point with an amplitude of ±4° at reduced frequencies of k = 2.6–5.8 at a Re c = 12000. Pitching symmetries of 50/50, 40/60 and 30/70 are studied, where the symmetry is defined by the fraction of the cycle spent in the pitch down versus pitch up motion. The data show that for the 50/50 (symmetric) motions two alternating sign vortices, with equivalent strength, are formed as expected. The asymmetric cases show that a single vortex is formed during the “fast” portion of the pitching motion. Multiple vortices are formed during the “slow” portion of the pitching motion. The number of secondary vortices and the downstream evolution of the vortices depends on the symmetry value. In some cases they remain isolated but orbit other vortical structures, while in other cases they pair with other vortical structures, and finally when the reduced frequency and asymmetry values are high enough the vortex array shows interaction between cycles.

The effect of placing a stationary airfoil downstream of a pitching airfoil in a tandem configuration was investigated using computational fluid dynamics. A finite-volume URANS solver was employed to simulate the flow field. An optimization study was performed at the low Reynolds number of 30,000 for SD7003 airfoils using a genetic algorithm. Three artificial neural networks were coupled with the genetic algorithm to reduce the computational cost. The optimization process was used to find the optimum design parameters in order to maximize lift to drag ratio of airfoils together in addition to the thrust coefficient of the hindfoil. It was found that a significant improvement of 174% and 15% could be achieved for the thrust coefficient of static hindfoil and L/D of the combined airfoils respectively in tandem configuration over single airfoils at optimum values of design parameters.

Active control of flow dynamics is an ever-increasing field of study. This paper presents details on a new experimental apparatus for the testing of dynamic fluid-structure interaction and the control thereof. Specifically, the experimental apparatus involves an open return wind tunnel with an actuated rigid airfoil. The airfoil is replaceable, but this initial evaluation uses a flat plate as a basic geometry airfoil. The pitch of the airfoil is actively controlled in real-time via an angle sensor, microcontroller, and brushed DC electric motor. For a given PID controller, sine sweeps are performed to identify the closed-loop frequency response at varying wind speeds. The wind creates a destabilizing feedback torque on the airfoil which increases with wind speed and results in a change in the closed-loop frequency response. A model of the system is augmented with a mathematical model for the wind dynamics. The assumed form wind model is manually tuned to match the closed-loop frequency response for all wind speeds. Finally, steady-state tests are conducted in which the setpoint for angle of attack is incrementally changed and the frequency spectrum is found for the angle sensor, in the control loop, and a hotwire behind the airfoil which is not in the control loop. The results from the angle sensor show the system holds steady, with slight variation at low frequencies for high angles of attack. The results from the hotwire show a distinct sheading frequency at low angles of attack, which is replaced by chaotic low frequency flow at high angles of attack characteristic of flow separation.

In this study, Reynolds-averaged Navier-Stokes simulation (RANS) of the uniform blowing and suction (UB/US) control on a Clark-Y airfoil was performed aiming at improving an airfoil performance by friction drag reduction. First, the control effect when only the uniform blowing control or uniform suction control is applied on the airfoil surface was investigated by changing the control locations. The blowing or suction velocity was 0.14% of the free-stream velocity and the blowing/suction area was set at four different locations from the leading edge to the trailing edge on both the upper and lower surfaces. The Reynolds number based on the chord length is 1.5 × 10 6 . The angle of attack is set to 0°. It was found that friction drag is decreased/increased by single UB/US control. It was also found that the lift-to-drag ratio improved with UB on the lower surface or US on the upper surface, and decreased with UB on the upper surface or US on the lower surface. In the combined control of UB and US, the blowing and suction velocity was 0.14% or 0.26% of the free-stream velocity and the locations of blowing/suction and flow conditions were the same as those in the cases with either UB or US. It seemed that the lift-to-drag ratio was improved by the combined control of UB on the lower surface and US on the upper surface. In particular, the lift-to-drag ratio was most improved by US on the lower rear surface and UB on the upper rear surface.

Friction drag reduction effect of a passive blowing on a Clark-Y airfoil is investigated. The passive blowing is conducted by the pressure difference on a wing surface between suction and blowing regions. The suction and the blowing regions are respectively set around the leading edge and the rear part of the upper surface. The Reynolds number based on the chord length is 0.65 × 10 6 and 1.54 × 10 6 . The angle of attack is set to 0° and 6°. The mean velocity profiles on the blowing region and the downstream are shifted away from the wall by the passive blowing. This behavior qualitatively suggests local reduction of skin friction on the wing surface. As a result of the quantitative assessment, which takes into account the effects of pressure gradient and the roughness of the wall, the local friction drag reduction effect of passive blowing is estimated to reach 4%–23%.

A rotorcraft main-rotor blade experiences a broad range of the Mach number in high-speed forward flights or rapid maneuvering conditions. Near the boundary of the flight envelope, the rotor blade often encounters severe dynamic stall which limits the overall aerodynamic performance. Of interest here is effects of the Mach number on the dynamic stall. A rotor airfoil, VR-12, is computationally investigated for both static and dynamic stall conditions with varying the Mach number from 0.2 to 0.4. For the small enough Mach number 0.2, both static and dynamic stall are significantly influenced by flow separation due to adverse pressure gradient. For the highest Mach number 0.4, compressibility near the leading edge is no longer negligible, forming a shock there. The shock-induced separation occurs near the leading edge, which dominates both the static and dynamic stall at the high Mach number. At the intermediate Mach number 0.3, it is observed that both adverse pressure gradient and the shock affect the stall. Current computations are conducted with the unsteady Reynolds-averaged Navier Stokes approach with the Spalart-Allmaras model. Numerical results are compared to relevant wind-tunnel test data.

The effects of applying a coupled unsteady blowing and suction combination over SD7003 airfoil at Reynolds number of 60,000 at an angle of attack of 13°, where a large separation on the suction side of the airfoil existed, was considered to investigate active flow control (AFC) mechanism. URANS equations were employed to solve the flow field and k–ω SST was used as the turbulence model. The unsteady blowing and suction were implemented at an angle to the surface crossing the boundary layer (CBL). The influence of location and frequency of the blowing/suction jets were examined.

Flow behaviors of a downstream object can be affected significantly by an upstream object in close proximity. This concept is used for flow control in this study to maximize the lift/drag ratio on a NACA0012 airfoil. A cylinder with cross-flow translational motion is placed upstream of the airfoil. Numerical simulations are carried out with an immersed-boundary method to solve the incompressible, viscous flow at the Reynolds number of 2000. Control parameters that influence the dynamics of flow around the airfoil are systematically investigated, including the oscillating frequency and amplitude of the upstream cylinder, the distances between the cylinder and the airfoil, and the diameter of the cylinder. To obtain sample data properly and efficiently for carrying out the neural network study, the idea of the orthogonal test method is used to set the control parameters in the numerical simulation. The combination of the back-propagation neural network algorithm and the genetic algorithm is applied to find the optimal value of the lift/drag ratio and the corresponding control parameters. The results show that when the cylinder oscillating frequency increases, the ratio increases until negative coefficients occur; when the distance between the cylinder and the airfoil increases or the amplitude of oscillating cylinder increases, the ratio decreases first and then increases; and when the cylinder diameter increases, the ratio increases. Compared to the reference case, the optimized lift/drag ratio increases 178%.

In this contribution, an assessment of compressible Reynolds Averaged Navier Stokes equations (RANS) and Large Eddy Simulation (LES) is presented using transonic organic vapor flow past a NACA4412 airfoil as a case study. The NACA4412 represents a canonical geometry, which, in case of air, has been well investigated numerically and experimentally. The results of the real gas simulations are compared with those of air simulations. For the real gas, the organic vapor Novec 649 ® is chosen as a representative fluid. The thermodynamic behavior of Novec 649 ® is modeled with the Peng-Robinson equation of state. Different inlet Mach numbers are applied, namely, a sub-critical, the critical, and a super-critical Mach number. It turns out, that the critical Mach number of the NACA4412 airfoil increases when Novec 649 ® is used as working fluid. Furthermore, it is shown that real gas flow simulations cause additional difficulties for the computational fluid dynamics (CFD) analysis. Although the speed of sound of Novec 649 ® is lower than the speed of sound of air, a finer grid resolution is required for the real gas simulations due to its high density. Based on an extensive simulation study, an assessment of different numerical modelling strategies and methods is given.

Uncertainties will make aircraft deviate from the designed condition, resulting in the decrease in aerodynamic performance and even destruction. This paper presents a fast nonlinear interval analysis method considering geometric uncertainties. DFFD method is used to parameterize the airfoil shape, and the Kriging model for aerodynamic force and uncertainty variables is optimized by PSO algorithm to find the upper and lower bounds of the objective interval. The effects of geometric uncertainties on NACA0012 airfoil are analyzed using the above method. And then, a robust optimization design method is established based on the interval analysis method. FFD method is used to produce the deterministic design variables and the order relation of interval number is employed to transform the uncertain optimization to deterministic multi-objective optimization which is solved by MOPSO based on Pareto entropy. The robust optimization design is implemented for the symmetrical airfoil with the drag objective under geometric uncertainties and thickness constraint, and the results are compared with the deterministic optimization to validate the effectiveness of the developed method.

Nowadays, mini unmanned aerial vehicles (MUAVs) and micro air vehicles (MAVs) are not only beneficially used as aviation models but also as modern drones for military missions and other civilian applications. Hence, research and development of propulsion sources for MUAVs and MAVs dynamically increase with a future trend of high performance, but low energy consumption. Certainly, using micro and ultra-small-size gas turbine is a good option for the propulsion source. To achieve ideal flight of MUAVs and MAVs powered by micro and ultra-small-size gas turbines under this trend, understanding of flow phenomena at wide ranges of Reynolds number is essential. This research presents a 2D numerical study of characteristics of laminar flow separation and the trailing-edge vortex on a turbine vane at extra-low Reynolds numbers (Res) i.e. Re = 1800 and 3600, and three rotational angles (α) i.e. α = 0°, 15° and 30° using immersed boundary method (IBM). With this method, the problem of incompressible flow is addressed by a sharp interface IBM. Numerical results indicate that IBM can characterize phenomena of laminar separation flow, which usually happens on the turbine airfoil when laminar boundary layer cannot overcome adverse pressure gradients and viscous effects. To our current knowledge, this may be the first research to study flow behavior at such low Res for gas turbine vanes using IBM. Even though it is now not common to operate micro and ultra-small-size gas turbines under these conditions, it is important to know how aerodynamic performance may be if micro and ultra-small-size gas turbines need to run under such conditions in the near future.

To provide an engine room of mobile hydraulic vehicle with an effective cooling for the combination set of a radiator, a charge air cooler, and an oil cooler, a 500mm-diameter, axial fan is designed to have a 8,800 m 3 /hr at a resistance of 20mm Aq static pressure with a sound power level less than 86 Lw A . The design parameters of sweep angle amplitude, wavelength of sweep angle change, airfoil type, and stagger angle are examined in terms of fan performance and its sound power generation. The surface curvatures generated by the sinusoidal sweep angle variation in the radial direction are proved to result in quite different flow patterns, thereby different types of specific sound power characteristics at the same flowrate. The acoustic noise sources are examined and discussed by using an acoustic imaging technique.

In recent years, the bladeless fan that does not have visible impellers have been widely applied in household appliances. Since the customers are particularly sensitive to noise and the strength of wind generated by the fan, the aerodynamic and acoustic performances of the fan need to be accurately characterized in the design stage. In this study, computational fluid dynamic (CFD) and computational aeroacoustics (CAA) are applied to investigate the performances of different designs of a bladeless fan model. The influence of four parameters, namely the airfoil selection for cross-section of the wind channel, the slit width, the height of cross-section and the location of the slit, is investigated. The results indicate the streamwise air velocity increases significantly by narrowing the outlet, but the noise level increases simultaneously. In addition, the generated noise increases while the height of fan cross-section increases, and a 4mm height of the cross section is optimal for aerodynamic performance. When the slit is closer to the location of maximum thickness, the performances of the bladeless fan increases. Moreover, the performance is not changed significantly by changing the cross-sectional profile. Finally, the optimal geometric parameters are identified to guide the future design of the bladeless fan.