Non-zero Angle Of Attack Research Articles

100 kHz PLEET velocimetry in a Mach-6 Ludwieg tube.

Picosecond laser electronic-excitation tagging (PLEET) was demonstrated in a Mach-6 Ludwieg tube at a repetition rate of 100 kHz using a 1064 nm, 100 ps burst-mode laser. The system performance of high-speed velocimetry in unseeded air and nitrogen Mach-6 flows at a static pressure in the range of 5-20 torr were evaluated. Based on time-resolved freestream flow measurements and computational fluid dynamics (CFD) calculations, we concluded that the measurement uncertainty of 100 kHz PLEET measurement for Mach 6 freestream flow condition is ∼1%. The measured velocity profiles with a cone-model agreed well with the CFD computations upstream and downstream of the shockwave; downstream of the shockwave the discrepancy between the CFD and experimental measurement could be attributed to a slight nonzero angle of attack (AoA) or flow unsteadiness. Our results show the potential of utilizing 100 kHz PLEET velocimetry for understanding real-time dynamics of turbulent hypersonic flows and provide the capability of collecting sufficient data across fewer tests in large hypersonic ground test facilities.

Development and testing of non-viscid solver based on UST3D programming code

Non-viscous versions of UST3D code elaborated in Laboratory of Radiative Gas Dynamics of Ishlinsky Institute for Problems in Mechanics were developed to testify the numerical method validity. Two numerical cases were examined to determine solution discrepancy between non-viscous and original algorithmes. The first case consisted in computational modeling of a flow over blunted cylinder-flare configuration under zero and non-zero angles of attack. The second case represented hypersonic flow over standard HB-2 ballistic model under angle of attack. Agreement of numerical results obtained with non-viscous solver and the original code was observed. These results approve the possibility of UST3D application for non-viscous problems.

Transport dynamics of an ellipsoidal particle in free molecular gas flow regime

Translational and rotational dynamics of a transporting particle in the free molecular gas flow regime are investigated using an in-house, multi-species, three dimensional Direct Simulation Monte Carlo (DSMC) solver. The DSMC algorithm is modified to study free molecular gas flows and validated against the analytical results for the dynamics of a spherical particle. A particular focus of this work is on estimating the effects of particle size, shape, and orientation on the dynamics of an ellipsoidal particle in free molecular gas and comparing them with that of a spherical particle. Properties such as particle speed, temperature, drag, and heat transfer coefficients are considered for comparison. The effect of particle shape on the aforesaid properties is qualitatively discussed and quantified through a comprehensive analysis taking care of lift, pitching moment, particle rotation, and the associated resistive torque. The relaxation of an ellipsoidal particle to surrounding gas conditions is simulated at zero and non-zero angles of attack to demonstrate the effect of particle orientation. Furthermore, the particle size effect on its translational and rotational dynamics is discussed. Finally, the trajectory of a spherical particle is compared with that of an ellipsoidal particle at different eccentricities.

Numerical investigations on the sabots discard process of an APFSDS at different angles of attack

Sabot asymmetric discard after the projectile being launched from the muzzle at various angles of attack and 4Ma is investigated. This is implemented by the coupling computational fluid dynamics (CFD) and six degrees-of-freedom exterior ballistic code through the unstructured dynamic mesh and user defined function. The flow field characteristics during sabot discard process and the trajectory parameters of all three sabots have been obtained. In addition, the aerodynamic coefficients of the projectile are also obtained. The numerical results show that the asymmetric discard of sabot is more obvious along with the increasing angle of attack. Moreover, the aerodynamic forces of projectile have a larger change and the pressure distribution of its surface is more asymmetric and complex. This means the aerodynamic interference at a non-zero angle of attack contributes more significantly to shooting dispersion and flight stability than that at zero angle of attack and the influence increases as the angle of attack increases.

Study of buzz phenomenon using visualization of external shock structure

An experimental study was carried out on an axisymmetric supersonic inlet with external compression in order to investigate the buzz phenomenon at different angles of attack and mass flow rates. The model was equipped with accurate and high-frequency pressure sensors, and the tests were conducted at Mach numbers varying from 1.8 to 2.5, for various angles of attack. Shadowgraph visualization technique, together with a high-speed camera, was used to provide the visual description of the shock structure in front of the inlet and to study the characteristics of buzz. Furthermore, pressure distribution over the spike surface was measured using several pressure sensors. Frequency of the buzz and shock displacement were measured by inspection of visualization pictures in each test. The obtained data from shadowgraphs were compared with those obtained from pressure measurements, and good agreement was found between them. The results revealed that for a moderate value of mass flow rate, the frequency of shock oscillation decreases as Mach number increases. Further, by increasing angle of attack, the shock displacement of oscillation will increase. At non-zero angles of attack, the displacement and frequency of shock motion show different behaviors on the leeward and windward sides of the body.

The method for calculating the traction resistance of a flat disk

Scuffler with flat discs are widely used in areas, which are subject to wind erosion of soils, to cover moisture on the stubble fields. Flat discs do not wrap the layers they remove, but only loosen them, shifting them aside. It is preferable to use such scrapers during presowing tillage and on steam, since in this case there is less sputtering of the soil, and its lower (moist) layers are not carried to the surface. Therefore, many industrially manufactured scufflers are equipped with flat discs, and the task of constructing a mathematical model for the interaction of such discs with soil is topical. Many mathematical models of the disc-soil interaction, taking into account the distribution of soil reaction forces on the disk, have been proposed, but all these models were constructed assuming that the disk is moving in its plane. Therefore, they are not applicable to the disk of a lakeshield moving with a non-zero angle of attack in the soil. The purpose of this work is to construct a mathematical model that allows calculating the traction resistance of a disk moving with a given angle of attack. The basic assumptions made in the construction of this model: the speed of the translational movement of the disk and its depth are constant; friction in the disk bearing can be neglected; the soil pressure on the lateral surface of the disc and the force of the soil reactions per unit length of its blade can be replaced by their mean values. An explicit expression is obtained for the traction resistance of a disk of a scuffler moving with an angle of attack, depending on its relative depth, the angle of attack, and the empirical coefficients. The adequacy of the constructed mathematical model is confirmed by comparison with known experimental data. It makes it possible to significantly reduce the number of full-scale experiments on the study of the interaction of flat discs of a scuffler with the soil, replacing them with computational ones, and reasonably carry out power and strength calculations of such discs. In addition, it can be used to optimize the parameters of the flat discs of the scuffler.

Improved k-ω-γ transition model by introducing the local effects of nose bluntness for hypersonic heat transfer

After considering the local effects of nose bluntness on the transition, a modification for the timescale of the second mode in the three-equation k-ω-γ transition model has been proposed to improve the hypersonic transition past cones. The ratio of local unit Reynolds number over freestream unit Reynolds number (Relocal/Re∞) is introduced to delay and reduce the development of second mode instability, according to the results of linear stability theory. The hypersonic transitions past a cone with three nose radii at the angle of attack of 0° were simulated to calibrate the new model and its constants. Then, the improved model is extended to predict the hypersonic transitions past the cones with larger radius and at non-zero angles of attack. Despite the original model can reflect the transition trend, the improved model performs much better than the original one when comparing with the available measurements.

Effect of sound-absorbing coatings on the disturbance evolution in a flow of a mixture of vibrationally excited gases

The flow around a solid plate and a plate with a sound-absorbing coating at a non-zero angle of attack in a hypersonic flow (M∞=8.44) of a mixture of vibrationally excited carbon dioxide and nitrogen is considered. Numerical simulations are performed by solving two-dimensional unsteady Navier–Stokes equations with a two-temperature model of relaxing flows. The vibrational energy as a function of time is defined by the Landau–Teller equation. A skeleton model, which is a set of square elements arranged in a staggered order, is used for simulating the porous coating made of foamed nickel with a porosity coefficient of 95%. The distance between the elements is equal to the pore diameter of the real sound-absorbing material. Data on the evolution of disturbances on the solid plate and on the plate with the sound-absorbing coating are presented for various angles of attack and CO2 concentrations in the mixture. The experimental and calculated data on pressure fluctuations on the plate surfaces are found to be in good agreement. The effects of various parameters of the sound-absorbing coating (depth, length, and location at the flat plate) are considered. It is shown that the sound-absorbing coating significantly reduces the intensity of pressure fluctuations on the plate surface as compared to the solid surface (up to 50% depending on the length and location of the sound-absorbing coating).

Numerical analysis on the effect of angle of attack on evaluating radio-frequency blackout in atmospheric reentry

A three-dimensional numerical simulation model that considers the effect of the angle of attack was developed to evaluate plasma flows around reentry vehicles. In this simulation model, thermochemical nonequilibrium of flowfields is considered by using a four-temperature model for high-accuracy simulations. Numerical simulations were performed for the orbital reentry experiment of the Japan Aerospace Exploration Agency, and the results were compared with experimental data to validate the simulation model. A comparison of measured and predicted results showed good agreement. Moreover, to evaluate the effect of the angle of attack, we performed numerical simulations around the Atmospheric Reentry Demonstrator of the European Space Agency by using an axisymmetric model and a three-dimensional model. Although there were no differences in the flowfields in the shock layer between the results of the axisymmetric and the three-dimensional models, the formation of the electron number density, which is an important parameter in evaluating radio-frequency blackout, was greatly changed in the wake region when a non-zero angle of attack was considered. Additionally, the number of altitudes at which radio-frequency blackout was predicted in the numerical simulations declined when using the three-dimensional model for considering the angle of attack.

Aerodynamic characteristics of multiple flapping wing configurations

Natural flyers and man-made MAVs generally use multiple flapping wing configurations. To understand the aerodynamic performance, three different flapping configurations: single wing, tandem wings, and biplane wings are numerically simulated by a URANS solver coupled with an overset grid method. Moreover, effect of kinematics including oscillating frequency, angle of attack and wing to tail distance are detailed investigated. Results show that the wing-tail interaction significantly benefits the thrust generation when the wings are tandem arranged. Additionally, the tandem arrangement is the most efficiency configuration when applied with high frequencies. Biplane wings model has the most inefficiency propulsive performance, nevertheless it can provide an extensive aerodynamic force. With the increasing AOA, biplane has the largest critical angle from thrust to drag. Wing-tail interaction becomes weaker when the tail is mounted further from the flapping wings. The present of the tail in tandem model bring more benefits compared with the tail in biplane model. The tail in biplane model is only functional for flight control when applied with a non-zero angle of attack.

Computational fluid dynamics (CFD) simulation of a newly designed passive particle sampler

In this work a series of computational fluid dynamics (CFD) simulations were performed to predict the deposition of particles on a newly designed passive dry deposition (Pas-DD) sampler. The sampler uses a parallel plate design and a conventional polyurethane foam (PUF) disk as the deposition surface. The deposition of particles with sizes between 0.5 and 10 μm was investigated for two different geometries of the Pas-DD sampler for different wind speeds and various angles of attack. To evaluate the mean flow field, the k-ɛ turbulence model was used and turbulent fluctuating velocities were generated using the discrete random walk (DRW) model. The CFD software ANSYS-FLUENT was used for performing the numerical simulations. It was found that the deposition velocity increased with particle size or wind speed. The modeled deposition velocities were in general agreement with the experimental measurements and they increased when flow entered the sampler with a non-zero angle of attack. The particle-size dependent deposition velocity was also dependent on the geometry of the leading edge of the sampler; deposition velocities were more dependent on particle size and wind speeds for the sampler without the bend in the leading edge of the deposition plate, compared to a flat plate design. Foam roughness was also found to have a small impact on particle deposition.

Tonal noise production from a wall-mounted finite airfoil

This study is concerned with the flow-induced noise of a smooth wall-mounted finite airfoil with flat ended tip and natural boundary layer transition. Far-field noise measurements have been taken at a single observer location and with a microphone array in the Virginia Tech Stability Wind Tunnel for a wall-mounted finite airfoil with aspect ratios of L/C=1–3, at a range of Reynolds numbers (ReC=7.9×105–1.6×106, based on chord) and geometric angles of attack (α=0–6°). At these Reynolds numbers, the wall-mounted finite airfoil produces a broadband noise contribution with a number of discrete equispaced tones at non-zero angles of attack. Spectral data are also presented for the noise produced due to three-dimensional vortex flow near the airfoil tip and wall junction to show the contributions of these flow features to airfoil noise generation. Tonal noise production is linked to the presence of a transitional flow state to the trailing edge and an accompanying region of mildly separated flow on the pressure surface. The separated flow region and tonal noise source location shift along the airfoil trailing edge towards the free-end region with increasing geometric angle of attack due to the influence of the tip flow field over the airfoil span. Tonal envelopes defining the operating conditions for tonal noise production from a wall-mounted finite airfoil are derived and show that the domain of tonal noise production differs significantly from that of a two-dimensional airfoil. Tonal noise production shifts to lower Reynolds numbers and higher geometric angles of attack as airfoil aspect ratio is reduced.

Regimes of tonal noise on an airfoil at moderate Reynolds number

Tonal noise generated by airfoils at low to moderate Reynolds number is relevant for applications in, for example, small-scale wind turbines, fans and unmanned aerial vehicles. Coherent and convected vortical structures scattering at the trailing edge from the pressure or suction sides of the airfoil have been identified to be responsible for such tonal noise generation. Controversy remains on the respective significance of pressure- and suction-side events, along with their interaction for tonal noise generation. The present study surveys the regimes of tonal noise generation for low to moderate chord-based Reynolds number between $\mathit{Re}_{c}=0.3\times 10^{5}$ and $2.3\times 10^{5}$ and effective angle of attack between $0^{\circ }$ and $6.3^{\circ }$ for the NACA 0012 airfoil profile. Extensive acoustic measurements with smooth surface and with transition to turbulence forced by boundary layer tripping are presented. Results show that, at non-zero angle of attack, tonal noise generation is dominated by suction-side events at low Reynolds number and by pressure-side events at high Reynolds number. At smaller angle of attack, interaction between events on the two sides becomes increasingly important. Particle image velocimetry measurements complete the information on the flow field structure in the source region around the trailing edge. The influences of both angle of attack and Reynolds number on tonal noise generation are explained by changes in the mean flow topology, namely the presence and location of reverse flow regions on the two sides. Data gathered from experimental and numerical studies in the literature are reviewed and interpreted in view of the different regimes.

Pressure Gradient Effects on Supersonic Transition over Axisymmetric Bodies at Incidence

Boundary-layer transition on axisymmetric bodies at a nonzero angle of attack in Mach 2 supersonic flow was investigated using experimental measurements and linear stability analysis. Transition over four axisymmetric bodies (namely, the Sears–Haack body, the semi-Sears–Haack body, the straight cone, and the flared cone) with different axial pressure gradients was measured in two different facilities with different unit Reynolds numbers. The semi-Sears–Haack body and flared cone were designed specifically to achieve a broader range of axial pressure distributions. Measurements revealed a dramatic effect of body shape on transition behavior near the leeward plane of symmetry. For a body shape with an adverse pressure gradient (that is, a flared cone), the measured transition patterns show an earlier transition location along the leeward symmetry plane in comparison with the neighboring azimuthal locations. For a nearly zero pressure gradient (that is, the straight cone), such leeward-first transition is observed only at the larger unit Reynolds number. Finally, transition occurs farther downstream along the leeward plane for the remaining two body shapes with a favorable pressure gradient. The observed transition patterns are partially consistent with the numerical predictions based on linear stability analysis.

Experimental investigation of boundary layer transition on rotating cones in axial flow in 0° and 35° angle of attack

In this paper, experimental results of using hot wire anemometer and smoke visualization are presented. The results obtained on the hot wire anemometer for critical Reynolds number and transitional Reynolds number are compared with previous results. Excellent agreement is found for the transitional Reynolds number. The results for the transitional Reynolds number are also compared with previous linear stability results. The results of the smoke visualization clearly show the crossflow vortices which arise in the transition process from a laminar to a turbulent flow. A non-zero angle of attack is also considered. We compare our results with linear stability theory put forward by Garrett and Peake (Eur J Mech B Fluid 26:344–353, 2007). Also the visualization and hot wire anemometer results have been compared graphically. The goal of this paper is to check reliability of using hot wire anemometer and smoke visualization in transition problems and check reliability of linear stability theory for this case and compare our results with some trusty experimental works.

Single velocity-component modeling of leading edge turbulence interaction noise.

A computational aeroacoustics approach is used to predict leading edge turbulence interaction noise for real airfoils. One-component (transverse), two-component (transverse and streamwise), and three-component (transverse, streamwise, and spanwise) synthesized turbulence disturbances are modeled instead of harmonic transverse gusts, to which previous computational studies of leading edge noise have often been confined. The effects of the inclusion of streamwise and spanwise disturbances on the noise are assessed. It is shown that accurate noise predictions can be made by modeling only transverse disturbances which reduces the computational expense of simulations. The accuracy of using only transverse disturbances is assessed for symmetric and cambered airfoils, and also for airfoils at non-zero angle of attack.

In-Plane Forces Prediction and Analysis in High-Speed Conditions on a Contra-Rotating Open Rotor

Due to the growing interest from engine and aircraft manufacturers for contra-rotating open rotors (CROR), much effort is presently devoted to the development of reliable computational fluid dynamics (CFD) methodologies for the prediction of performance, aerodynamic loads, and acoustics. Forces transverse to the rotation axis of the propellers, commonly called in-plane forces (or sometimes 1P forces), are a major concern for the structural sizing of the aircraft and for vibrations. In-plane forces impact strongly the stability and the balancing of the aircraft and, consequently, the horizontal tail plane (HTP) and the vertical tail plane (VTP) sizing. Also, in-plane forces can initiate a flutter phenomenon on the blades or on the whole engine system. Finally, these forces are unsteady and may lead to vibrations on the whole aircraft, which may degrade the comfort of the passengers and lead to structural fatigue. These forces can be predicted by numerical methods and wind tunnel measurements. However, a reliable estimation of in-plane forces requires validated prediction approaches. To reach this objective, comparisons between several numerical methods and wind tunnel data campaigns are necessary. The primary objective of the paper is to provide a physical analysis of the aerodynamics of in-plane forces for a CROR in high speed at nonzero angle of attack using unsteady simulations. Confidence in the numerical results is built through a code-to-code comparison, which is a first step in the verification process of in-plane forces prediction. Thus, two computational processes for unsteady Reynolds-averaged Navier–Stokes (URANS) simulations of an isolated open rotor at nonzero angle of attack are compared: computational strategy, open rotor meshing, aerodynamic results (rotor forces, blades thrust, and pressure distributions). In a second step, the paper focuses on the understanding of the key aerodynamic mechanisms behind the physics of in-plane forces. For the front rotor, two effects are predominant: the first is due to the orientation of the freestream velocity, and the second is due to the distribution of the induced velocity. For the rear rotor, the freestream velocity effect is reduced but is still dominant. The swirl generated by the front rotor also plays a major role in the modulus and the direction of the in-plane force. Finally, aerodynamic interactions are found to have a minor effect.

On high-frequency noise scattering by aerofoils in flow

AbstractA theoretical model is developed for the sound scattered when a sound wave is incident on a cambered aerofoil at non-zero angle of attack. The model is based on the linearization of the Euler equations about a steady subsonic flow, and is an adaptation of previous work which considered incident vortical disturbances. Only high-frequency sound waves are considered. The aerofoil thickness, camber and angle of attack are restricted such that the steady flow past the aerofoil is a small perturbation to a uniform flow. The singular perturbation analysis identifies asymptotic regions around the aerofoil; local ‘inner’ regions, which scale on the incident wavelength, at the leading and trailing edges of the aerofoil; Fresnel regions emanating from the leading and trailing edges of the aerofoil due to the coalescence of singularities and points of stationary phase; a wake transition region downstream of the aerofoil leading and trailing edge; and an outer region far from the aerofoil and wake. An acoustic boundary layer on the aerofoil surface and within the transition region accounts for the effects of curvature. The final result is a uniformly-valid solution for the far-field sound; the effects of angle of attack, camber and thickness are investigated.

Evaluation of Assumed-Transition-Point Criterion in Context of Reynolds-Averaged Simulations Around Lighter-Than-Air Vehicles

In this work, the performance of the assumed-transition-point criterion, in conjunction with Reynolds-averaged Navier–Stokes simulations is evaluated over the hull of a lighter-than-air vehicle. The evaluation is performed by comparing the Reynolds-averaged Navier–Stokes computations of drag, lift, and moment coefficients against those measured in experiments where 1) transition is forced to occur (forced-transition case) and 2) transition is allowed to occur naturally (free-transition case). It was found that by using the assumed-transition-point criterion Reynolds-averaged Navier–Stokes computations satisfactorily recover the drag coefficient both at zero angle of attack and nonzero angles of attack up to five degrees when compared against the forced-transition case. However, irrespective of the location of the assumed transition point, Reynolds-averaged Navier–Stokes fails to recover the experimentally observed values of the force and moment coefficients for the free-transition case. At this point, it is not clear if this failure is attributable to the assumed-transition-point criterion or to the Reynolds-averaged Navier–Stokes framework itself. A similar systematic evaluation with more scale resolving computational fluid dynamics methods (large eddy simulation or bridging methods) may be helpful in ascertaining the reason of this failure.

Nadal’s Formula and High Speed Rail Derailments

Several railroad vehicle derailment criteria are based on the (L/LVV) ratio where L is the lateral force and V is the vertical force acting on the wheelset. Derailment is assumed to occur if this ratio exceeds a certain limit. The (L/LVV) ratio has its roots in Nadal’s formula which was introduced more than a century ago. When a planar analysis that corresponds to zero angle of attack is used, Nadal’s formula can be derived using a geometric approach, while a Nadal-like formula can be derived using a kinetic approach. In the geometric approach, a coordinate transformation is used to define the normal and friction forces in another coordinate system. In this case, the lateral and vertical forces are interpreted as the components of a vector that defines the normal and friction forces in another coordinate system without consideration of other external forces that are applied to the wheelset. Because the geometric approach does not account for other forces, it should not be used as the basis for derailment studies. In the kinetic approach, on the other hand, the L and V forces are interpreted as the resultant of the forces excluding the normal and friction forces at the contact point. That is, in both approaches discussed in this paper, the L and V forces cannot be interpreted as the resultant forces acting on the wheelset. The condition for the wheel climb using the kinetic approach is obtained and examined. It is shown that formulas obtained in this paper are based on assumptions that do not capture the gyroscopic moments, and therefore, these formulas should not be used as the basis for general high speed rail derailment criteria. It is also shown that in the case of zero angle of attack and using single degree of freedom planar kinetic model assumptions, an increase in the lateral force L can reduce the tendency for wheel climb, while reducing L can increase the wheel climb risk. Furthermore, the single degree of freedom assumptions used to obtain the wheel climb formula presented in this paper do not allow for wheel lift, and as a consequence, wheel lift derailment scenarios that can be the result of large moment should not be investigated using Nadal’s formula or one of its derivatives that employ the same assumptions. Furthermore, since the analysis presented in this paper assumes zero angle of attack, the conclusions obtained in this investigation do not apply to the wheel climb scenarios with nonzero angle of attack. It is also important to point out that this paper is not intended as a discussion on how Nadal’s formula is interpreted by researchers and engineers; instead, the paper is mainly focused on examining the roots of this formula and the problems that can arise from the assumptions used in its derivation.