Non-zero Angle Of Attack Research Articles

Influence of flow interactions on the aerodynamic characteristics of tandem self-propelled flapping wings with a nonzero angle of attack

It is common in nature for birds or insects to fly in flocks. This study sought to understand the interaction mechanism between complex flows and the aerodynamic characteristics of flocks of flying organisms by employing the lattice-Boltzmann method to investigate tandem self-propelled flapping wings with an angle of attack of 10°. The effects of the initial heaving phase, the initial spacing between the fore and hind wings, and the phase difference between the heaving motions of the fore and hind wings were investigated. It was found that when the fore and hind wings flap in phase, the initial heaving phase and initial spacing can influence the final locomotive state of the tandem system, resulting in three modes: stable flight, collision, and separation. When the tandem system eventually achieves stable flight, only one equilibrium state is observed. In this equilibrium state, the trailing-edge vortex generated by the fore wing reattaches to the lower surface of the hind wing, resulting in 82.3% lower lift efficiency but 19.9% higher propulsive efficiency when compared to a single wing. When the fore and hind wings flap nearly out of phase, the tandem system has better lift characteristics while maintaining good propulsive performance. These findings improve the understanding of the principles of lift and thrust generation in flock flight and will help guide the design of bionic micro air vehicles.

Experimental investigation of wing-propeller aerodynamic interaction in eVTOL configurations

The present study is aimed at the experimental investigation of the effects of wing-propeller aerodynamic interaction in a boom-mounted configuration typically used in eVTOL aircraft. The investigation was focused on the repercussions on wing and propeller performance coming from the variation of angle of attack, advance ratio and blades sense of rotation. Moreover, particular attention was devoted to the evaluation of the effect of the propeller's longitudinal offset. Results of a comprehensive wind tunnel campaign performed on a propeller-mounted wing scaled model confirmed the advantageous effects of an inboard-up rotating propeller on lift generation and drag reduction, while outlined the opposite sensitivity of wing and propeller performances when exposed to a non-zero angle of attack. Propeller's longitudinal offset instead led to slight alterations on wing and propeller aerodynamic performance, while a noteworthy sensitivity on the mounting setup was perceived by propeller aerodynamic performance. Moreover, PIV measurements allowed to evaluate quantitatively the effects of the investigated parameters on propeller slipstream behaviour and blade tip vortices pattern.

Mixed convection flow over a horizontal plate and the horizontal wake far downstream

The laminar mixed convection flow over a heated or cooled horizontal plate of finite length is a surprisingly intriguing problem. The hydrostatic pressure jump at the trailing edge and across the wake is to be compensated by induced circulation in the outer potential flow, similar to a plate at a non-zero angle of attack. The mixed convection flows over semi-infinite and finite horizontal plates, respectively, are thus substantially different. As the hydrostatic pressure jump remains finite along the infinitely extended wake, the induced outer potential flow around the plate would grow beyond bounds in a domain extending to infinity. Thus, boundary conditions must be imposed at the boundaries of a finite domain. However, the numerical solution of the Navier-Stokes equations remains a challenging problem due to the lack of appropriate outflow conditions. The traditional outflow condition cannot comply with the finite hydrostatic pressure jump across the wake. Thus, an asymptotic expansion for the wake far downstream from the plate is performed. Remarkably, the perturbation of the flow does not decay with increasing distance from the plate as in a classical wake. The asymptotic solution is implemented as an outflow boundary condition for the numerical solution of the Navier-Stokes equations. The results are compared with a boundary-layer solution obtained in previous work for the limiting case of vanishing Prandtl number.

Experimental Investigation of Airfoil Instability Tonal Noise Reduction Using Structured Porous Trailing Edges

Reducing the tonal noise from airfoil instabilities has attracted significant interest from the aeronautical community in the past few years. The aim of this paper is to investigate the effect of structured porous trailing edges on the tonal noise reduction performance of a symmetrical NACA 0012 airfoil. Detailed parametric testing was performed in an open-jet wind tunnel between the baseline solid trailing edge and seventeen structured porous trailing edges with different sub-millimeter-scale pores. The experimental results demonstrate that structured porous trailing edges can reduce the noticeable tonal noise of the symmetrical NACA 0012 airfoil. Moreover, the design parameters for the structured porous edges have slightly different impacts on the tonal noise reduction performance between a zero angle of attack (α = 0°) and a non-zero angle of attack (α = 10°): better airfoil tonal noise reduction is due to the porous parameters of small pore coverage, small-to-moderate chordwise spacing, and moderate spanwise spacing at α = 0°. On the other hand, the optimal combination of the structured porous edge at α = 10° is the configuration with larger pore coverage, smaller chordwise spacing, and spanwise spacing.

Aerodynamic and Aeroacoustic Effects of Different Transition Mechanisms on an Airfoil

The paper delves into the detailed sound generation and propagation mechanisms associated with tripping-induced flow perturbations and trailing-edge scattering using wall-resolved large-eddy simulations. Two distinct boundary-layer tripping techniques, namely a geometrically resolved stair strip and an artificially modeled trip using suction and blowing, are investigated. To facilitate comparison, the natural boundary-layer transition is also simulated as a baseline scenario. The analysis takes into account a Reynolds number of 4×105, a Mach number of 0.058, and a nonzero angle of attack of 6.25° over a NACA 0012 airfoil configuration. The mechanisms for sound generation and propagation related to trailing-edge noise remain consistent across the three transition scenarios. However, boundary-layer tripping notably leads to intricate, scenario-specific noise generation: there is an interaction between the laminar separation bubble and tripping for the stair strip case, whereas laminar boundary-layer instability is evident for the suction and blowing scenario. Aerodynamic flowfields involving acoustic noise sources, their propagating natures near the wall, and far-field acoustics are cross-examined in detail. The comprehensive analysis of observed phenomena provides valuable insights for understanding nonlinear flow and acoustic interactions relevant to airfoil noise and designing new types of trips under adverse pressure gradient flows.

Airfoil response to periodic vertical and longitudinal gusts

Gust response has consistently been a concern in engineering. Critical theories have been proposed in the past to predict the unsteady lift response of an airfoil experiencing vertical gusts by Atassi, and longitudinal gusts by Greenberg. However, their applicability for an airfoil with non-zero angles of attack still needs clarification. Thus, force measurements are conducted to examine these theories’ validity and quasi-steady corrections are applied to compensate potential disparities between the idealised and real flow conditions. Velocity measurements are performed to scrutinise the effect of gusts on the flow around the airfoil, and subsequently to reveal the underlying mechanism governing the airfoil's response to gust-induced perturbations. In the study, two pitching vanes are arranged upstream to generate periodic vertical and longitudinal gusts, whereas a downstream airfoil with angles of attack of 0–12° is subjected to two gust types. It is found that Greenberg's theory demonstrates superior predictive capability in pre-stall regimes, with the potential for its effectiveness to be expanded to post-stall regimes through theoretical refinements. In contrast, Atassi's theory exhibits significant deviations from experimental outcomes across the measured angles of attack. Nevertheless, a modified version of the theory aligns better with experimental results at small angles of attack, whereas substantial discrepancies persist as the angle of attack increases. In the pre-stall regime, the aerodynamic response of the airfoil to vertical gusts displays a linear correlation with the flow angle near the leading edge. In the post-stall regime, the vertical gust induces dynamic stall of the airfoil. The flow angle has an essential effect on the lift coefficient but it alone is inadequate to dictate the trend of the lift coefficient. The vorticity statistics show that negative vortex circulation strongly correlates with the lift coefficient. Thus, further correction of the theory or a new vortex model can be expected to predict the lift variation.

Heat flux predictions for hypersonic flows with an overset near body solver on an adaptive block-structured Cartesian off-body grid

The manual volume mesh generation process for body-fitted CFD codes can be a time-intensive process, particularly for complex flight vehicles. The work shown here presents a hybrid body-fitted/Cartesian overset CFD solver capable of automatic volume mesh generation capabilities and accurate surface heating predictions. The solver uses a body-conformal near body solver for capturing boundary layer effects and an off-body Cartesian solver with adaptive mesh refinement for shock and wake tracking within the CHAMPS CFD code. The formulated near body Cartesian solver is the first of its kind to be applied within the context of low and high enthalpy, hypersonic viscous flows. The developed solver is validated on a Mach 9.47 Mars Science Lander and a Mach 6 Orion Crew Exploration Vehicle in a low-enthalpy flow at various non-zero angles of attack followed by a Mach 25.2 AS-202 capsule at a 17.8-degree angle of attack for Earth re-entry. Validation is shown based on grid convergence studies as well as good agreement to available numerical and experimental data in consideration of surface heating and surface pressure predictions.

Experimental and Numerical Analysis of Incompressible Flow over an Iced Airfoil

Determining the aerodynamic characteristics of iced airfoil is an important step in aircraft design. The goal of this work is to study experimentally and numerically an iced airfoil to assess the aerodynamic penalties associated with presence of ice on the airfoil surface. Three iced shapes were tested on NACA 0012 straight wing at zero and non-zero angles of attack, at Reynolds No. equal to (3.36*105). The 2-D steady state continuity and momentum equations have been solved utilizing finite volume method to analyze the turbulent flow over a clean and iced airfoil. The results show that the ice shapes affected the aerodynamic characteristics due to the change in airfoil shape. The experimental results show that the horn iced airfoil consumes more power than the other shapes of ice, its value was (44.4W). The horn iced shape has the worst effect on the airfoil than the other shapes. The present results are compared with previously reported results; it is found there is a very good agreement between them. A comparison between the experimental and computational results of the presented work were pursing the same behavior.

The aerodynamic analysis of iced and clean NACA 4410 and 0012 airfoils

AbstractThe aims of the work is to develop a numerical iced airfoil; quarter round forward ice was tested on NACA 4410 and NACA 0012 airfoils at zero and non-zero angles of attack, and Reynolds numbers equal to (2∙105, 2.4∙105) based on airfoil chord and Mach number 0.04. The two-dimensional steady-state momentum equations with the continuity equation have been solved applying a finite volume method to examine the turbulent flow over a clean and iced airfoil. The cambered airfoil NACA 4410 spends less power than the unsymmetrical airfoil at the same angle of attack. The reported numerical results demonstrated that for airfoil NACA 4410, the drag was increased by 40%, and the lift was reduced by 22%. However, for airfoil NACA 0012, the drag was increased by 43%, and the lift was decreased by 21%.

Mixed-Compression Supersonic Intake and Engine–Airframe Integration

Hypersonic flight is gaining increasing attention by aerospace companies interested in designing and developing reusable aircraftlike vehicles (spaceplanes). Advanced concepts include a combination of space and aviation approaches able to fly few minutes in space for sightseeing or fast enough to enable point-to-point transportation. In this framework, a concept of hypersonic systems for space travel and high-speed transportation is presented, with a particular focus on the intake and engine–airframe configurations. The adopted intake is designed to ensure an efficient functioning, not only in nominal conditions but along the supersonic ascent trajectory thanks to a movable inner spike. A nonzero angle of attack resulted in a little degradation of the performances in terms of efficiency, spillage, and flow distortion. In the second part, the interaction between the engine and airframe is also investigated. An initial configuration, characterized by the propulsion system located on the fuselage side near the wing, is characterized by a strong interference responsible for an off-design functioning with a high percentage of air spillage. An alternative solution with engine–wing integration exhibits a completely different behavior with interference minimized, resulting in a design intake functioning without air spillage and a considerable drag reduction.

Trailing edge noise reduction using bio-inspired finlets

Inspired by the silent flight of the owl, the current work focuses on the reduction of trailing edge noise using 2D finlets. Overset-Large Eddy Simulations are carried out of two configurations, (i) the baseline configuration consisting of a NACA0012 airfoil and (ii) an identical configuration, however, now with finlets installed near the trailing edge. The considered chord based Reynolds number Rec equals 4.2×105 for a non-zero angle of attack α of 4°. The resulting turbulence statistics in the boundary layer and the power spectral density (PSD) spectrum of the surface pressure fluctuations show good agreement with the experimental data. Detailed analysis of the unsteady flow data revealed various noise reduction mechanisms for the here considered case. Firstly, finlets provide a ‘lifting-up’ effect for the most energetic eddies in the boundary layer, which prolongs downstream of the finlets. As a result, the edge scattering is significantly weakened. Secondly, the finlets mounted on the airfoil provide an additional wetted area to the flow which in-turn has a dissipating effect on the surface pressure fluctuations, thereby again weakening the edge scattering phenomena. A reduction in the flow velocity is observed for the flow exiting the finlets channel and towards the trailing edge, which also has a favourable effect in reducing the trailing-edge noise. Finally, a marginal breakdown of the span-wise coherent length scale for frequencies between 2.5 kHz to 9 kHz is also observed.

Thermal Protection Mechanism of a Novel Adjustable Non-Ablative Thermal Protection System for Hypersonic Vehicles

In order to improve the thermal protection performance of the active thermal protection system (TPS) based on the spike and jet, an adjustable non-ablative thermal protection system, of which the spike can be rotated in the direction of the free stream, is proposed in this paper. The thermal protection mechanism and the optimal installation angle are analyzed by adopting the numerical method. The results show that the angle of attack has great influence on the peak heat flux of hypersonic vehicles, the dangerous point is on the windward side of the vehicles at the non-zero angle of attack. With the increase in angle of attack, the heat flux of the windward side of the vehicles rises rapidly, leading to the decrease in the global thermal protection efficiency. The adjustable non-ablative TPS in this paper greatly reduces the aeroheating of the windward side through the installation angle between the spike and nose cone, thus improving the global thermal protection efficiency. The optimal installation angle can be obtained by numerical or experimental methods in engineering design, and the difference between the angle of attack and the optimal installation angle is about 2.4° for the proposed model. Therefore, the installation angle can be automatically adjusted based on the angle of attack to achieve the highest thermal protection efficiency.

Effect of Mach number on the aeroacoustic feedback loop generating airfoil tonal noise

Airfoil tonal noise emission at low-to-moderate Reynolds number and flow conditions featuring a laminar separation bubble close to the trailing edge is often related to an aeroacoustic feedback mechanism and, therefore, the Mach number is a primary parameter for the flow field and noise generation. This study experimentally explores the effect of the Mach number on airfoil tonal noise generation in the nominally incompressible flow regime. Using airfoil profiles of different chord lengths, the Mach number is varied for a constant Reynolds number. Acoustic and flow field measurements for a range of combinations of Reynolds and Mach numbers were conducted. At zero incidence, the tonal noise regime in the Reynolds number domain is found to be sensitive to the Mach number. At non-zero angle of attack (2°), the noise generation is found to be dominated by vortex shedding over a separation bubble on the pressure side. The details of the separation bubble and shedding process depend on the Mach number. The frequencies of the dominant tones and the frequency intervals between tones increase with the Mach number. Moreover, the measured frequency interval can be collapsed using a relation based on the aeroacoustic feedback loop model. The relation is rewritten to separate the effects of the Reynolds and Mach numbers. As a result, the dependence on the Mach number is identified and tested. In contrast, the tonal noise level shows a more complex dependence on details of the laminar separation bubble and the vortex shedding process.

Aerodynamic Forces on a Bluff Cylinder in Sinusoidal Streamwise Winds with Different Angles of Attack

In the present study, multiple-fan active control wind tunnel tests are conducted to investigate the aerodynamic forces on a 5:1 rectangular cylinder in sinusoidal streamwise winds with different angles of attack (AoA). The effects of the frequency, amplitude, and AoA of the sinusoidal flow on the statistical parameters, spectral characteristics, and spanwise distributions of drag, lift, and moment coefficients are analyzed. Results show that each force has two components: the one induced by the approaching velocity oscillation and the one induced by wake vortex-shedding—this is quite different from that in the smooth flows, where the fluctuating forces are totally due to wake vortex-shedding. For each force, changes of the two components and their relationship with the frequency, amplitude, and AoA are presented. The drag fluctuations are generally dominated by the approaching velocity oscillations, whereas the lift and moment are more sensitive to wake vortex-shedding. Therefore, the drag force has better spanwise correlations than the lift and moment forces. Meanwhile, at a non-zero AoA, the inflow amplitude has different effects on the vortex-shedding-induced component as that at a zero AoA. The differences of spanwise distributions between the sinusoidal flow cases and the smooth flow cases are analyzed.

Comparison of Flap and Bank-Angle Steering for Mars Entry

Previous Mars entry, descent, and landing systems have used bank-angle steering during the hypersonic phase of entry to control range. An alternate solution for hypersonic steering is a set of independently articulated aerodynamic flaps. Deflecting these flaps results in trim at nonzero angles of attack and sideslip angles, enabling the vehicle to generate lift in arbitrary directions. This study compares current state-of-the-art bank-angle steering to flap-based steering across several metrics, including aerodynamic control authority, range capability, and response time. The effective bank angle and effective lift-to-drag ratio are developed to compare the control authority between flap and bank-angle steering systems. The shape of the nonaxisymmetric control authority depends on the flap placement, area, and number of flaps; however, flap-based steering systems likely require a mechanism for roll or bank in order to have a uniform control authority. Aerodynamic control authority is then mapped to a range capability. Flap steering vehicles exhibit range capabilities with markedly different shapes than bank-angle steering vehicles with equivalent maximum hypersonic lift-to-drag ratios. Flap steering vehicles are found to be able to rotate directly between angular positions and have both a more uniform and faster response time than bank-angle steering vehicles.

Aerodynamic performance of a self-propelled airfoil with a non-zero angle of attack

In the natural world, numerous flying creatures generate both thrust and lift by flapping their wings. Aerodynamic mechanisms of forward flight with flapping wings have received much attention from researchers. However, the majority of previous studies have simplified the forward-flight motion of flapping wings to be uniform, and there has been no detailed evaluation of the validity of this simplification. Motivated by this, aerodynamic characteristics of a self-propelled flapping wing with a non-zero angle of attack were investigated. The results showed that the asymmetric leading-edge vortex produced in the wing's upstroke and downstroke leads to transient thrust, driving the self-propelled wing to move with variable forward velocities. Compared to the uniform forward-flight cases, significant losses in lift and severe changes in the flow field were observed in self-propelled flapping wings. In addition, the changes in the aerodynamic performance—including the forward propulsion speed, lift, and power efficiency—of the self-propelled flapping wing with changes in various dimensionless parameters were also investigated. The heaving amplitude was shown to have significant effects on lift and propulsion speed of the self-propelled flapping wing, while the effects of ratio between the airfoil density and fluid density as well as the Reynolds number, were relatively small. In most conditions, when the Strouhal number was in the range 0.2–0.4, the self-propelled flapping wing performed well in terms of both lift generation and propulsive efficiency. These research results suggest that it is necessary to consider the fluctuating forward speed in aerodynamic modeling of propulsive flapping wings.

Numerical study on the supersonic near-wake flow structures of a cylinder at an angle of attack

The supersonic near-wake flow experiment conducted by Dutton on a cylinder with 10 degrees angle of attack was studied numerically with Delayed Detached Eddy Simulation (DDES) method. The wind tunnel test section were included in the geometry model to examine if its omission led to the unsuccessful prediction of several important flow features in previous studies. Flow structures were compared to experiment measurements, Gaitonde's k-ε model results and our previous Monotonically Integrated Large Eddy Simulation (MILES) results. It was revealed that with present geometry model and numerical method, good agreements with wind tunnel experiment result were obtained, including the wave structures, density fields and velocity fields. In particular, the inclusion of wind tunnel test section wall in geometry model reproduces the finite area nozzle exit flow and the expansion wave behind the recirculation zone in the actual experiment setup, which lead to correct prediction of the position of maximum streamwise velocity point. The present simulation also reproduced the uniform pressure profile and value observed in the experiment and the deviation between computation and experiment was within 10%. The flow field data obtained was further analyzed to investigate the vortex surface shape and recirculation region shape in near wake flow with non-zero angle of attack and significant differences were found compared to the zero angle of attack case. The vortex surfaces show a non-axisymmetric slitting movement in the near wake region, which gives an explanation to the non-circular density contour pattern on streamwise flow field slices. The recirculation region shape was identified with zero isosurface of streamwise velocity and was found to be of complex three-dimensional shape with maximum length beside the symmetry plane and a local minimum at windward side of the symmetry plane, which is significantly different from that of zero angle of attack case. The observation that the length of the recirculation zone was greatly reduced on the symmetry plane at non-zero angle of attack in previous study is due to the aforementioned local minimum on the symmetry plane. The reduction of recirculation length is much less if we consider the three-dimensional shape of the region.

The aerodynamic characteristics and its estimation method of finned missile experiencing spin-deformation coupling motion

The coupling of rolling motion and bending deformation of a finned missile can significantly affect its lateral force and yawing moment, thus threatening flight stability and maneuverability. By solving the unsteady Reynolds-averaged Navier-Stokes (URANS) equations with a dual-time stepping method and a previously proposed coupling motion mode, flow over the Air Force Modified Basic Finner Missile (AFF) experiencing spin and spin-deformation coupling motion were simulated, and the aerodynamic characteristics were obtained. The relationship between the variation tendency of aerodynamic characteristics and different motion modes was analyzed. On this basis, an aerodynamic model previously proposed by Murphy was revised and a modified model suitable for the aerodynamic prediction of spin-deformation coupling missiles was developed and verified. The results indicate that as angle of attack (AoA) increases, the total time-averaged lateral force induced by spin-deformation coupling motion alters direction and increases by an order of magnitude compared with that induced by spin, and the time-averaged body lateral force becomes dominant. Combining the three aerodynamic components, including that of rigid spinning missile at non-zero AoA, that produced by bending deformation and spin-deformation coupling effect at zero AoA, and that account for the aerodynamic interference effect at non-zero AoA, the accuracy of the modified aerodynamic model can be greatly improved. At forebody cross sections, the results of the modified aerodynamic model and the computational fluid dynamic (CFD) method match well. While at aftbody cross sections, the transient results have certain deviation, while the relative differences for the time-averaged normal force and lateral force are within 10% and 20%, respectively.

Fish-like swimming in oblique flows: A numerical investigation

Although being a common situation for fish and underwater vehicles, limited research has been reported for a fish-like body swimming in oblique incoming flows. The present work employs the immersed boundary-lattice Boltzmann method to numerically study the dynamics of this problem. A more comprehensive analysis method, which utilizes the effective force parameters that consolidate the angle of attack (AOA) into the traditional force parameters, is proposed for investigating the dynamics of cases with non-zero AOA. Several distinct features in terms of the dynamic performance, the periodic features, and the flow field characteristics are presented. For the dynamic performance, undulation amplitude plays a more dominant role than the undulation form in thrust generation. The periodic features revealed by the Fourier spectra present increasing complexity with larger AOA. Besides, three flow regimes are identified, one of which observes hierarchical vortex structures with different orders in time and space. The small-scale vortex street generated by undulation is “local” in both time and space, and it is embedded in the large scale vortex street generated by the body and the free-stream.

On the effects of anisotropic turbulence on leading edge noise

This paper presents an investigation on the broadband noise produced by the interaction of anisotropic turbulence with isolated aerofoils. Computational aeroacoustic simulations are performed using a synthetic turbulence method with a linearised Euler solver. A comprehensive and fundamental parameter study on leading edge noise is presented to assess the effects of aerofoil thickness, mean flow Mach number, and angle of attack in the presence of moderately anisotropic turbulence. To this end, the streamwise-to-transverse length scale ratio is varied from 0.33 to 3, which can be representative of the anisotropy in the fan wakes from aero-engines, grid-generated turbulence in open-jet wind tunnel experiments, and rotors ingesting turbulent boundary layers. Anisotropic turbulence presents a significant redistribution of the energy in the turbulence spectra in comparison with an isotropic baseline, which affects the resulting noise spectra. It has been found that the noise spectra are mainly influenced by the transverse velocity component for the examined anisotropic flows. The frequency at which the noise spectrum is at a maximum only relies on the mean flow speed and the axial length scale of the anisotropic turbulent flow. Noise reduction due to aerofoil thickness at high frequencies follows a linear trend, but the amount of noise reduction is sensitive to the anisotropy in the flow. It is shown that the contribution of the streamwise disturbances and length scale become increasingly important for thick aerofoils at high Mach numbers and for aerofoils at non-zero angle of attack.