Small Angles Of Attack Research Articles

Comparative Study of Different Turbulence Models for Cavitational Flows around NACA0012 Hydrofoil

The present work focuses on the comparison of the numerical simulation of sheet/cloud cavitation with the Reynolds Average Navier-Stokes and Large Eddy Simulation(RANS and LES) methods around NACA0012 hydrofoil in water flow. Three kinds of turbulence models—SST k-ω, modified SST k-ω, and Smagorinsky’s model—were used in this paper. The unstable sheet cavity and periodic shedding of the sheet/cloud cavitation were predicted, and the simulation results, namelycavitation shape, shedding frequency, and the lift and the drag coefficients of those three turbulence models, were analyzed and compared with each other. The numerical results above were basically in accordance with experimental ones. It was found that the modified SST k-ω and Smagorinsky turbulence models performed better in the aspects of cavitation shape, shedding frequency, and capturing the unsteady cavitation vortex cluster in the developing and shedding period of the cavitation at the cavitation number σ = 0.8. At a small angle of attack, the modified SST k-ω model was more accurate and practical than the other two models. However, at a large angle of attack, the Smagorinsky model of the LES method was able to give specific information in the cavitation flow field, which RANS method could not give. Further study showed that the vortex structure of the wing is the main cause of cavitation shedding.

Effect of Aerospike on Unsteady Transonic Flow over a Blunt Body

Blunt-nosed launch vehicles featuring large nose-cone angles experience high levels of pressure fluctuations over the payload region at transonic Mach numbers due to shock-wave/boundary-layer interactions. The main cause for this phenomenon appears to be flow instability associated with separated flow and formation of a vortex pair. Interactions between the induced velocity of the vortex pair and oncoming mean flow cause oscillations of the -shock system and high levels of fluctuating pressures. An aerospike causes flow separation at the nose and reattachment of the shear layer downstream, energizing the boundary layer. Consequently, flow separation and vortex formation are prevented: shock oscillations are stabilized. Dramatic reductions in the pressure fluctuations, around 95–35% in the low-frequency range, are observed along the payload region at small angles of attack. The observations are based on wind-tunnel tests involving unsteady pressure measurements, surface-flow patterns, and high-speed shadowgraph recordings on a blunt nose cone with various cone angles.

The Impact of Icing on the Airfoil on the Lift-Drag Characteristics and Maneuverability Characteristics

Icing has now become an important factor endangering flight safety. This paper takes the icing data of the NACA 23012 airfoil as an example, establishes an icing influence model for real-time simulation based on icing time and aircraft angle of attack, and analyzes the influence of different icing geometry on aircraft characteristics. The two-dimensional interpolation method is used to improve the model of the aircraft’s stall area, which is mainly divided into the correction of the lift-drag coefficient linear area and the stall area and the correction of the aircraft stability derivative and the control derivative. The aerodynamic equation of the airplane after icing is established, and the modal analysis of the airplane under different icing conditions is completed through the linearization of the flight equation. The closed-loop simulation system of the altitude holding mode and roll attitude holding mode is used to calculate and analyze the flight quality changes of the aircraft after the wing surface is frozen. The analysis results show that, under icing conditions, in the range of small angles of attack, icing has no obvious influence on the aircraft mode. As the degree of icing increases, the throttle skewness and the negative deflection angle of the airplane’s level flight requirements continue to increase. The case of icing flight in altitude hold and roll hold modes shows that flying in the autopilot mode under severe icing conditions is very dangerous and is prone to cause the aircraft to stall.

DESIGN OF DELTA WING AUTONOMOUS UNDERWATER GLIDERS FOR LONG TERM SUSTAINABLE AND STEALTH RECONNAISSANCE USING NUMERICAL STUDIES

Autonomous Underwater Gliders (AUGs) are a unique source of collecting oceanographic data for varying depths with longer endurance than other types of unmanned underwater vehicles. The buoyancy-driven propulsion mechanism assists in diving deep into the ocean, collect data, resurface in saw-tooth and spiral motions. There have been various hull-forms introduced for this purpose starting from - Torpedo hull, XRay Flying-Wing Glider, etc., all of which have seen an improved understanding of the hydrodynamics associated with the hull-form of the gliders. The saw-tooth motion represents the glider's capacity to cover larger distances with a smaller angle of attack, thereby achieving an increased range for the same power included within the hull batteries. In this paper, a design study of delta wing hull-forms, which are proven to have better longitudinal motion characteristics, is proposed to arrive at a preliminary hull form that can be used for long term reconnaissance missions, remaining in stealth to assist submarines with valuable data. Delta wing hull forms are varied using different NACA sections, and numerical analysis of the obtained hull forms is performed to arrive at the hydrodynamics; namely, the drag, lift forces and coefficients for varying angles of attack, using commercial CFD software – StarCCM+. These results are then compared against each other to derive the preliminary hull form for further study on better longitudinal motion characteristics, thereby using less power for its motion underwater.

Insights on the unsteadiness and stall effects on the characteristics and responses of continuous wing-based systems

The effects of the aerodynamic loads on continuous wing-based systems can be represented using a quasi-steady or a fully unsteady reduced-order modeling. Frequently, the quasi-steady formulation is used to model the aerodynamic loads at small reduced frequency values. This approximation, unlike the unsteady formulation, does not account for any lag between the unsteady oscillations and its effect on the aerodynamic loads which might lead to over or under predicting some crucial information related to the system studied, particularly concerning the onset of flutter or type of instability. To this end, in this effort, a comparison of the effects of different aerodynamic formulations on a wing-based system is conducted. A linear analysis considering the influence of each noncirculatory term in both formulations is inspected individually, and the results generated from both aerodynamic formulations are compared while depicting the effects of the structural damping on the linear flutter speed and frequency. A nonlinear analysis is then performed to investigate the stall effect on the system’s response as well as the impacts of different nonlinearities including inertial and geometric. The system’s behavior is examined using higher modes in the Galerkin discretization to make sure that it is accurately modeled and that the number of modes selected does not fail to correctly portray its aeroelastic response. The results of this study show the importance of using the unsteady fluid–structure interaction with stall modeling for continuous wing systems even for small angles of attack and reduced frequency values.

The Numerical Simulation of the Vertical Water Entry Process of High Speed Projectile with Small Angle of Attack

Based on the numerical simulation method, the high speed vertical water entry process of projectile moving at 1000m/s with small angle of attack is studied. By analyzing the cavity configuration and hydrodynamic characteristics of the process, the influences of the small entry angle on the water entry stability process are obtained. The simulation results indicate that the projectile’s ballistic can achieve relative stable state when the water entry angle is less than 0.5°, otherwise the water entry cavity can be pierced by tail fin. The pierced size tend to increase with the increased water entry angle as well as the projectile’s ballistic become unstable. In addition, the rotation of the projectile owing to pitching moment has little effect on cavity configuration and water entry stability.

S4 - A demonstrator of HYPLANE, a single stage suborbital spaceplane and a hypersonic business jet

The mid-to-long term perspectives of suborbital hypersonic spaceplanes require the availability of ultra-fast, winged vehicles, characterized by low wing loading, streamlined fuselages, sharp nose and wing leading edges, able to manoeuvre along flight trajectories at small angles of attack. From another point of view, the cost associated with sub-orbital space access flight is strongly conditioned by the still small dimension and uncertainties of its market.Therefore, hypersonic technologies and vehicle design suitable for the suborbital market as well as for point-to-point fast transportation can facilitate the endeavour. A high-speed business-jet is widely recognized to be more technologically affordable than larger aircraft and even more feasible from an economic point of view. Along this wave, HYPLANE is proposed as a 6-pax hypersonic business jet designed for P2P stratospheric transportation and suborbital flight at reduced ticket price. It will be based on more readily state-of-art available technologies like titanium alloys for its structure which are suitable to manage the levels of heat produced on external surfaces at those speeds and turboramjet engines, instead of less proven scramjets required for faster aircrafts. No disruptive or currently low TRL technology is needed, rather enhancing already available innovations. Attention is given to maintaining high level of environmental protection, especially in terms of sonic boom effects on ground, global and local pollution and noise emissions. To flight validate the system and related technologies, a multi-mission Single Stage Suborbital Spaceplane (S4) demonstrator is proposed. In parallel, the evolution of the Campanian Grazzanise airport to become a spaceport is being supported in the frame of a Suborbital TEst Polygon (STEP) proposed to be established between Campania and Sardinia regions. It will make available the perfect operational scenario where to test and make use of the S4 demonstrator.

Experimental Study on Noise Reduction Characteristics of Slanting Serrated Trailing Edge Blades

In order to suppress the narrowband tone noise generated by the blunt part at the root of non-flat plate serrated trailing edge, a novel slanting root serrated trailing edge is adopted on NACA65019 blade. The experimental investigation on the aeroacoustic characteristics of the blade with slanting serrated trailing edge is carried out in the full anechoic wind tunnel laboratory. Reynolds number (Re) based on the chord is 2 × 105 and attack angle is 0° to 12° in the experiment. The result of experiment shows that the blade with traditional serrated trailing edge will reduce noise in high frequency range at small angle of attack, but it will produce narrow band peaks at specific frequencies, leading to an increase in overall noise level. The blade with slanting serrated trailing edge can maintain the noise reduction effect in broadband frequency range, while significantly weakening narrow band peaks. At the large attack angle, the narrow-band peaks of the traditional blunt trailing edge disappear and the noise reduction effect is obvious, while the novel slanting serrated trailing edge has more obvious noise reduction effect. The results of this paper can pro-vide a new idea for the bionic design configuration of blades.

Aerodynamic Interaction Effects Between Propellers in Typical eVTOL Vehicle Configurations

Many electric vertical takeoff and landing concepts are characterized by nontraditional vehicle layouts with distributed propellers. Two propeller interaction types were distinguished in this Paper, which investigates how propeller interaction in side-by-side and one-after-another configuration affects performance, in terms of thrust, power, in-plane forces, and out-of-plane moments, and how those performance effects depend on axial and lateral propeller spacing. A wind-tunnel experiment was performed with two propeller units, one instrumented with a force/torque sensor and the other introducing the aerodynamic interaction. Total pressure and planar particle-image velocimetry measurements were taken to investigate slipstream characteristics. A strong dependency of interaction effects on the geometric layout was found. For side-by-side interaction characteristic of vertical takeoff and transition, interaction effects varied from weak at small angle of attack to strong at larger angles. A drop in rear propeller thrust of up to 30% was found at constant advance ratio. Keeping thrust constant resulted in power penalties up to 13% for the two propellers combined. For one-after-another interaction, characteristic of cruise, a maximum reduction of thrust of up to 80% was observed. Thrust compensation led to power penalties up to 30% for the rear propeller alone. An extended blade element momentum model captured most interaction effects with sufficient accuracy.

Aerodynamic performance enhancement of co-flow jet airfoil with simple high-lift device

The present study performed a numerical investigation to explore the performance enhancement of a co-flow jet (CFJ) airfoil with simple high-lift device configuration, with a specific goal to examine the feasibility and capability of the proposed configuration for low-speed take-off and landing. Computations have been accomplished by an in-house-programmed Reynolds-averaged Navier-Stokes solver enclosed by k-ω shear stress transport turbulence model. Three crucial geometric parameters, viz., injection slot location, suction slot location and its angle were selected for the sake of revealing their effects on aerodynamic lift, drag, power consumption and equivalent lift-to-drag ratio. Results show that using simple high-lift devices on CFJ airfoil can significantly augment the aerodynamic associated lift and efficiency which evidences the feasibility of CFJ for short take-off and landing with small angle of attack. The injection and suction slot locations are more influential with respect to the aerodynamic performance of CFJ airfoil compared with the suction slot angle. The injection location is preferable to be located in the downstream of the pressure suction peak on leading edge to reduce the power expenditure of the pumping system for a relative higher equivalent lift-to-drag ratio. Another concluded criterion is that the suction slot should be oriented on the trailing edge flap for achieving more aerodynamic gain, meanwhile, carefully selecting this location is crucial in determining the aerodynamic enhancement of CFJ airfoil with deflected flaps.

Effects of CFJ flow control on aerodynamic performance of symmetric NACA airfoils

This paper investigates the effects of co-flow jet on the aerodynamic performance of some symmetric blade sections of wind turbines. The Reynolds number of 1.3 × 105, angles of attack between 0° and 18° and momentum coefficients of 0.03, 0.05, and 0.08 are considered for four airfoils of NACA 0012, 0015, 0018, and 0021 as the blade sections. To numerically simulate the fluid flow, the Navier-Stokes equations are solved with a transitional turbulence model. The aerodynamic coefficients along with the lift to drag ratio are calculated and the results are compared with together for the baseline and co-flow jet geometries. The results of the baseline airfoils show that the NACA 0012 and NACA 0015 airfoils generate a larger lift at small angles of attack where the NACA 0018 and NACA 0021 airfoils produce a higher lift at larger angles of attack. Furthermore, the obtained results reveal that applying the co-flow jet has a positive effect on delaying the stall angle and increasing the lift coefficient. The aerodynamic coefficients are also more improved by increasing the thickness once the appropriate momentum coefficient is applied. The results also indicate that the momentum coefficient of 0.05 is sufficient for most of the cases.

Verification of the numerical analysis using a collocated grid system directly solving the quasi-inviscid incompressible flow around a symmetric airfoil

The present study describes the results of validating a high-quality computational scheme of kinetic energy conservation characteristics under a collocated grid system. In the present study, the pseudo-inviscid incompressible flow around the Kármán-Trefftz airfoil is analyzed. In this analysis, the kinetic energy conservation error is reduced by temporally evolving the governing equation using pressure increments in fractional steps. This sufficient conservation of kinetic energy is validated using an inviscid homogeneous isotropic fluctuation field. The velocity field obtained by this analysis is compared to that of the potential flow obtained using the Kármán-Trefftz airfoil. The numerical analysis is also validated using the lift coefficient profile as a function of a small angle of attack.

Dynamic Stall Under Combined Pitching and Surging

Dynamic stall often occurs under conditions of simultaneous unsteady pitching and surging (e.g., rotorcraft and wind turbines), but many models employ a dimensionless time base that implicitly assumes that surging is superimposed, in a quasi-steady manner, on dynamic pitching. An unsteady wind tunnel was used to examine this assumption, where a technique was developed to quantify the unsteady effects of surging on a pitching NACA 0018 airfoil. The technique involved performing multiple harmonic pitching experiments under nominally steady freestream conditions that bracketed a corresponding 50% surging amplitude (). By interpolating these data, unsteady-pitching/quasi-steady-surging data sets were constructed and compared with de facto synchronous pitch and surging experiments, thereby isolating the unsteady effects of surging on a pitching airfoil. Both large and small poststall maximum angles of attack ( and ) were considered at multiple pitch-surge phase differences. During deep dynamic stall (), with large-scale separation, surging was seen to have a secondary effect on the unsteady aerodynamics. However, at small poststall maximum angles of attack (), either light or deep dynamic stall behavior was observed depending upon the pitch-surge phase difference. This was attributed to Reynolds number history effects, exemplified by boundary-layer transition, and thus it can be referred to as “transitional” dynamic stall.

Newtons Laws Explain How Frisbees Fly

Newtonian mechanics explain the physics of how frisbees fly. Specifically, a frisbee in horizontal flight with a positive angle of attack (AOA), will fly through a static mass of air each second (‘m/) that it accelerates to a velocity (’a’) downwards, to create a downward force (Force = ma), due to Newtons 2nd law of motion. The ‘equal & opposite’ upward force generated (Newtons 3rd law of motion), provides an upward force and vertical lift. Air goes down and the frisbee goes up. It’s that simple.
 There are two airflows: The underside of the frisbee pushes air down. While the curved topside of the frisbee pulls air downwards due to the Coanda effect. This explanation is different to other theories of lift currently available, such as those based on fluid mechanics (Bernoulli and Navier-Stokes) and vortices.
 The spin on a frisbee enhances the stability of flight due to the gyroscopic effects. In turn, stability of flight allows the frisbee to generate laminar (smooth) airflows, and thus better lift. The spin itself does not directly contribute towards lift. In addition, airflow vortices account for trick throws, where the frisbee appears to defy normal physics.
 So what? This provides new insight into lift and flight not currently found in accepted physics textbooks. The best performing frisbee will be one that maximises the Coanda effect on the topside of the disc; to increase the amount of air displaced down. Newtons laws can explain why a frisbee thrown flat (small AOA) generates better lift and will fly further than a frisbee thrown high, on a parabolic type of path. This explanation of lift can also be applied to airplane wings.

Asymptotic Solution of a Boundary Value Problem for a Spring\u2013Mass Model of Legged Locomotion

Running is the basic mode of fast locomotion for legged animals. One of the most successful mathematical descriptions of this gait is the so-called spring–mass model constructed upon an inverted elastic pendulum. In the description of the grounded phase of the step, an interesting boundary value problem arises where one has to determine the leg stiffness. In this paper, we find asymptotic expansions of the stiffness. These are conducted perturbatively: once with respect to small angles of attack, and once for large velocities. Our findings are in agreement with previous results and numerical simulations. In particular, we show that the leg stiffness is inversely proportional to the square of the attack angle for its small values, and proportional to the velocity for large speeds. We give exact asymptotic formulas to several orders and conclude the paper with a numerical verification.

Simulation of separated flow past an inclined and normal plates by a discrete vortex method

Unsteady separated flow past a normal or an inclined flat plate is simulated using the discrete vortex method. The plate is replaced by attached discrete vortices and the separated flow motion behind the plate is modeled by free discrete vortices moving with a local velocity. The intensities of attached vortices at each moment of time are determined from the condition of impermeability of the streamlined body surface, as well as the condition that the total intensity of the attached and free vortices is equal to zero. A formula is introduced for determining the pressure coefficient in the position of the attached vortices, and it is obtained from the pressure in an unsteady-state flow by Cauchy-Lagrange function. The forces exerted on the plate at each instant of time determined from the generalized Blasius theorem.The calculations have shown, that by choosing an integration step and a quantity of the plate breakdowns, it is possible to obtain a satisfactory convergence with the experimental results on the normal force coefficient. But this results in some discrepancy in the Strouhal number of vortex shedding, especially at the small angles of attack. Replacement of the plate by the attached vortices leads to a passage of the free vortices through it, which leads to a discontinuous change in its aerodynamic characteristics. To avoid the latter, it is necessary to introduce a condition, for example, to exclude a normal velocity component of the free vortex near the plate. The introduction of the initial disturbance affects to some extent a subsequent development of the vortices and, thus, the dependence of the normal force coefficient on a time period.

Wind tunnel and flight tests of glider flow separation control by microsecond dielectric barrier discharge

As a new kind of active flow control technology, plasma flow control has a bright future, for its simple structure, fast response, and wide frequency band. The wind tunnel and flight tests were conducted with microsecond dielectric barrier discharge on a glider. For the tests, the microsecond pulse power supply and remote control system were designed and built. In the wind tunnel test, the flow separation on the glider wing surface can be controlled effectively, and static pressure at the leading edge pressure is decreases by 177%. The flow control effects under different pulse frequencies are compared, and the optimal pulse frequency for actuation is found to be 100 Hz. A significant hysteresis effect was observed with microsecond dielectric barrier discharge at small angle of attack (α ≤ 18°), which means the flow control effect can last more than 300 s after turning off the plasma actuation. In the flight test, the maximum roll angle decreases by 7.0°, and the maximum aileron deflection angle decreases by 9.4° with plasma actuation at both sides of the wing, which means the glider becomes more stable with microsecond dielectric barrier discharge. With unilateral actuation, the rolling moment generated by the plasma actuation is larger than that produced by the ailerons with the angle of attack within 12.94° ≤ α ≤ 29.77°, which shows strong rolling control ability of microsecond dielectric barrier discharge. The wind tunnel and flight tests results verified the flow control effect of microsecond dielectric barrier discharge, and paved the way for the plasma flow control technology to practical applications.

General-Purpose Object-Oriented Framework for Vorticity-Dominated Flow Simulation

This paper describes the development of a computational framework for aerodynamic simulations based on the unsteady vortex lattice method under the object-oriented programming paradigm. The effort aims to fulfill the fluid-dynamics portion of the fluid–structure interaction problem in the multiphysics field. The construction and implementation of an object-oriented model is introduced along with its analysis and design stages, focusing on the generality and versatility. The overall class organization is presented, and the main classes’ characteristics are discussed. The software validation indicates that the transient and the steady-state aerodynamics for the impulsively started motion of a body are accurately captured by the simulations. In addition, comparisons against experimental results show that rotors in hover flight are acceptably represented for small angles of attack of the blades, within the limitations of the flat-plate representation and the inviscid hypothesis of the model. Regarding applications, six different models are analyzed: 1) a single propeller rotor, 2) a small-scale quadcopter drone, 3) a flying-wing drone, 4) a joined-wing aircraft, 5) a drag-reduction system for cars, and, 6) a flapping-wing insect. In all cases, the results indicate that a well-tuned simulation based on vorticity-dominated flows is widely suitable for the aerodynamic study of several cases. All simulations expose the main advantages of a modular object-oriented program: the ability to change the case of analysis with minimal reprogramming, and the flexibility to link with different models in a co-simulation strategy.

Aerodynamic characteristics of morphing wing with flexible leading-edge

The morphing wing can improve the flight performance during different phases. However, research has been subject to limitations in aerodynamic characteristics of the morphing wing with a flexible leading-edge. The computational fluid dynamic method and dynamic mesh were used to simulate the continuous morphing of the flexible leading-edge. After comparing the steady aerodynamic characteristics of morphing and conventional wings, this study examined the unsteady aerodynamic characteristics of morphing wings with upward and downward deflections of the leading-edge at different frequencies. The numerical results show that for the steady aerodynamic, the leading-edge deflection mainly affects the stall characteristic. The downward deflection of the leading-edge increases the stall angle of attack and nose-down pitching moment. The results are opposite for the upward deflection. For the unsteady aerodynamic, at a small angle of attack, the transient lift coefficient of the upward deflection, growing with the increase of deflection frequency, is larger than that of the static case. The transient lift coefficient of the downward deflection, decreasing with the increase of deflection frequency, is smaller than that of the static case. However, at a large angle of attack, an opposite effect of deflection frequency on the transient lift coefficient was demonstrated. The transient lift coefficient is larger than that of the static case when the leading edge is in the nose-up stage, and lower than that of the static one in the nose-down stage.

Research on parameterization and optimization procedure of low-Reynolds-number airfoils based on genetic algorithm and Bezier curve

A systematic procedure of high-precision parameterization and multi-objective optimization for airfoils was proposed in this paper in order to improve the aerodynamic performance under full working conditions. The Bezier curve was applied in the airfoil parameterization with the coordinates of the 8 Bezier control points chosen to be the design variables. The control points were initialized by self-programming software and then optimized by direct search method to guarantee the fitting precision. The multi-objective optimization of low-Reynolds-number airfoil E387 was executed through NSGA II based on the result from XFOIL calculation, with the increment and variance of lift-drag ratio chosen to be the objective functions. Two representative solutions were selected from the Pareto front, and the performances of the new airfoils were promoted in different ways. The lift-drag ratio of the optimized airfoil with smaller variance is increased uniformly under different attack angles and maintains the similar characteristics with the original one. For the optimized airfoil with larger variance, the distribution of lift-drag ratio relative to the attack angle is altered obviously, and the performance under small attack angles and stall conditions are significantly ameliorated. The introduction of variance control in the objective function enables the differentiation of the optimized solutions for different requirement in engineering application.