Dual Vortex Structure Research Articles

Comparison of fluid structure downstream of an evaporative flameholder with subsonic uniform and subsonic-supersonic inflows

Subsonic uniform and supersonic mixing flows are widely present in the combustion chamber of a multi-mode combined scramjet engine, where the significant pressure, velocity, and temperature gradients of the two impose severe limitations on flame propagation. The vortex structures downstream of an evaporative flameholder were experimentally measured under subsonic-supersonic mixing inflow and subsonic uniform inflow. Results indicate that, under subsonic-supersonic mixing inflow conditions, an asymmetric vortex structure in the recirculation zone is formed due to the presence of a subsonic-supersonic shear layer, which exerts compressive or stretching effects on the recirculation zone downstream of the flameholder, in contrast to the symmetric dual-vortex structure observed under subsonic uniform inflow conditions. The asymmetric recirculation zone displays fluid being entrained from one vortex to another, with the degree and direction of entrainment dependent on the supersonic expansion state. Additionally, a simplified flow field structure is proposed for comparing the differences between the subsonic-supersonic mixing inflow and subsonic uniform inflow, along with the introduction of five regions and two stages to describe the flow field structure downstream of the flameholder under subsonic-supersonic mixing inflow. Analysis of the flow characteristics downstream of the flameholder under subsonic-supersonic mixing inflow and subsonic uniform inflow conditions can offer design insights for the combustion scheme of aerospace-integrated propulsion systems.

Mathematical modelling of couple stress fluid flow around a semi‐permeable sphere enclosing a solid core

AbstractThe focus of this article is to study theoretically the steady‐axisymmetric creeping flow dynamics of a couple stress fluid external to a semi‐permeable sphere containing a solid core. This problem is motivated by emulsion hydrodynamics in chemical engineering where rheological behaviour often arises in addition to porous media effects. The non‐Newtonian Stokes’ couple stress fluid model features couple stresses and body couples that are absent in the classical Navier–Stokes viscous model. It provides a robust framework for simulating emulsions, complex suspensions and other liquids which possess microstructure. The physical regime is delineated into two zones – the interior of the semi‐permeable zone (region II) which engulfs the solid core and the external couple stress fluid zone (region I). The model is formulated using a spherical polar coordinate system in terms of the stream function ψ. The Brinkman‐extended Darcy model is deployed for the porous medium hydrodynamics and isotropic permeability is considered. Analytical expressions are derived for dimensionless pressure, tangential stress and the couple stress components using the method of separation of variables and Gegenbauer functions of the first kind. The integration constants are evaluated with appropriate boundary conditions on the inner and outer boundary of the semi‐permeable zone with the aid of Mathematica symbolic software. Solutions for the drag force exerted by the couple stress fluid on the semi‐permeable sphere and volumetric flow rate are also derived with corresponding expressions for the drag coefficient and non‐dimensional volumetric flow rate. The influence of permeability (k), separation parameter (l), couple stress viscosity coefficient (η), couple stress inverse length dependent parameter (λ = √(μ/η)) and couple stress viscosity ratio (τ = η/η/) on all key variables is studied graphically. Additionally, streamline contours are computed for a range of parameters including inverse permeability parameter (α = 1/k). The computations show that increasing couple stress inverse length dependent parameter (λ) greatly reduces the dimensionless volumetric flow rate in particular at high values of separation parameter (l). Flow rate is, however, markedly enhanced with permeability (k) and also couple stress viscosity parameter (η). In the presence of the solid core, a much greater drag coefficient is observed at all values of couple stress inverse length dependent parameter (λ) relative to the case without a solid core. A significant distortion in streamlines is computed with increasing separation parameter (l) with a dual vortex structure emanating. With greater couple stress viscosity parameter (τ), no tangible modification is observed in the streamline contours. The present work generalizes the earlier study of Krishnan and Shukla to consider a semi‐permeable (porous medium) outer sphere and furthermore presents detailed streamline visualizations of the creeping flow for a range of emerging parameters of relevance to non‐Newtonian chemical engineering processes including emulsion droplet dynamics in porous materials.

Inlet static pressure ratio effect on vortex structure downstream of the flameholder in subsonic–supersonic mixing flow

The flow field characteristics downstream of the evaporative flameholder in the subsonic–supersonic mixing flow were experimentally investigated. The study focused on examining the effects of different inlet static pressure ratios characterized by supersonic and subsonic flow parameters. The results indicated that the increase in the static pressure ratio enhanced the fragmentation of the multiple vortices downstream of the flameholder located in the subsonic mainstream. It also exacerbated the asymmetry in the recirculation zone downstream of the flameholder and strengthened the tendency of the fluid to flow from one vortex to another. The regions with higher vorticity were mainly concentrated in the subsonic–supersonic shear layer between the subsonic and supersonic mainstream and the subsonic–subsonic shear layer region downstream of the flameholder. Furthermore, an increase in the static pressure ratio widened the range of peak distribution while reducing the magnitude of the peaks. The recirculation zone downstream of the flameholder exhibited four distinct changes in the vortex structure as the static pressure ratio increased from 1.07 to 1.96. These typical changes in the vortex structure observed are as follows: asymmetric dual-vortex structure, single vortex structure (away from the supersonic mainstream region), asymmetric dual-vortex structure, and single vortex structure (near the supersonic mainstream region).

Effect of non-uniform inlet profile on the combustion performance of an afterburner with bluff body

The quenching problem caused by the inlet distortion of the combustor during emergency attitude adjustment of an aircraft has not been adequately solved by a lack of experimental evidence in non-uniform inflow. Six non-uniform inlet velocity profiles with different velocity peak positions and values were built by perforated plates to explore the combustion performance of a bluff body. The results indicate that the velocity profile with a peak and off-center can produce the asymmetry of the dual-vortex structure downstream of the bluff body. The significant phenomenon is that the fluid flows from one vortex to another. More asymmetrical fluid-structure is facilitated by larger peak values or further off-center positions. Then, the asymmetrical flow field leads to non-uniform fuel distribution, which results in an asymmetrical flame and outlet temperature profile. Interestingly, the non-uniformity of outlet temperature increases first and then decreases with the rise of velocity peak position, while it just increases with the growth of the velocity peak value. Compare to uniform inflow, the non-uniform inlet velocity profile will destroy the ignition performance and combustion efficiency in most cases. However, the partial fuel accumulation may occur in the asymmetric fluid structure, which can widen the lean blowout limit. Overall, the flame stability limits and combustion efficiency under the non-uniform inflow in this study are beneficial to the design and application of combustor.

Effect of non-uniform inlet velocity profile on flow field characteristics of a bluff body

To explore the influence of the non-uniform inlet velocity profile of the bluff body on flame stabilization, PIV measurements are carried out to investigate the characteristics of the flow field at 300 K and 0.101 MPa with seven velocity profiles, including four-velocity peaks, h/H, and three non-uniformities, δv. The experimental results indicate that the recirculation is approximately symmetric with the non-interfering dual-vortex structure under the condition of uniform inlet flow, whereas it is asymmetric and the fluid flows from one vortex core into another one when the velocity peak rises. The tendency of the dual-vortex interaction becomes more pronounced with the increase of the non-uniformity and the velocity peak. Besides, the length of the recirculation zone increases with the increase of h/H, while the largest width occurs at h/H = 0 and 4/6. At the same time, the distance of one core decreases first and then increases, while the other one is opposite to that. The width of the recirculation zone along the x-axis corresponding to δv = 30% is the same as that at δv = 0, and its value is higher than that at δv = 40% and 50%.

Numerical study of guide vane effects on reacting flow characteristics in a trapped vortex combustor

ABSTRACTTrapped Vortex Combustor (TVC) is a simple and promising concept for flame stabilizations in aero-engines. In this concept, cavity trapped vortices are used to establish pilot flames for robust combustion performance. The main objective of this study is to numerically investigate the effects of guide vanes on the performance of a TVC installed in a small ramjet. Three dimensional steady and unsteady simulations are performed by solving the Reynolds-Averaged Navier-Stokes (RANS) equations with the commercial flow solver ANSYS FLUENT v16.1. The shear stress transport (SST) model is selected for the turbulence closure, and the eddy dissipation model (EDM) is used for the combustion modeling of gaseous propane (C3H8) fuel. In the current work, four cases are studied: (1) TVC without vanes, (2) TVC with a square-tip vane, (3) TVC with a sharp-tip vane, and (4) TVC with a sharp-tip vane and additional cavity fuel injection. Numerical results reveal that the introduction of guide vanes significantly changes the cavity vortex structure. Unlike the previous single cavity vortex, a counter-rotating dual-vortex structure is formed inside the cavity when the vane is implemented: one vortex is generated by the air jet guided by the vane and the other is the wake vortex downstream the vane. Because of the additional air introduced into the cavity by the guide vane, a fuel-lean mixture is obtained inside the cavity. This indicates fuel can be further injected inside to make full use of the air. Combustion results show that compared with the TVC without vanes, case 4 increases the combustion efficiency by 14% (to 99.3%) at the exit of combustor. Meanwhile, the combination of the sharp tip vane and cavity fuel injection only provides around 5% increase in aerodynamic drag and 3.9% in overall total pressure loss.

Analysis of cooling performance and combustion flow in advanced vortex combustor with guide vane

Numerical simulation was performed to study the flow characteristics of advanced vortex combustor (AVC) with guide vanes and slots for aircraft engine. The results show that structural parameters of guide vane have significant effects on AVC, when a/B=0.4, b/H=0.4, c/L=0.1, it can obtain desirable dual-vortex structure, acceptable pressure loss and better combustion efficiency. AVC with guide vane has better performance than AVC without guide vane. Five slots air film cooling are brought in, the effect of flow injection angle (θ=30°, 60° and 90°) and injection ratio (R=1.2, 1.5, 2.0) on cooling performance of AVC are analyzed. The results show that there are 5 change regions for each cooling efficiency under different injection conditions. The variation law of the first region is different from the four region behind. The first slot air mainly mixes with the main flow and gets involved in the combustion while the four region behind can play a good role in cooling efficiency. Meanwhile, the effects of θ and R on flow field, temperature field, total pressure loss, combustion efficiency, outlet temperature distribution factor (OTDF) and the average cooling efficiency are presented. Finally, the film cooling conclusions are explained by field synergy theory, and results show that it can optimize the wall cooling effectiveness by field synergy theory.

Experimental study of flow field distribution over a generic cranked double delta wing

The flow fields over a generic cranked double delta wing were investigated. Pressure and velocity distributions were obtained using a Pitot tube and a hot wire anemometer. Two different leading edge shapes, namely “sharp” and “round”, were applied to the wing. The wing had two sweep angles of 55° and 30°. The experiments were conducted in a closed circuit wind tunnel at velocity 20m/s and angles of attack of 5°–20° with the step of 5°. The Reynolds number of the model was about 2×105 according to the root chord. A dual vortex structure was formed above the wing surface. A pressure drop occurred at the vortex core and the root mean square of the measured velocity increased at the core of the vortices, reflecting the instability of the flow in that region. The magnitude of power spectral density increased strongly in spanwise direction and had the maximum value at the vortex core. By increasing the angle of attack, the pressure drop increased and the vortices became wider; the vortices moved inboard along the wing, and away from the surface; the flow separation was initiated from the outer portion of the wing and developed to its inner part. The vortices of the wing of the sharp leading edge were stronger than those of the round one.

Interaction of multiple vortices over a double delta wing

Interaction of strake and wing vortices over a 70°/50° double delta wing were studied experimentally in a wind tunnel using particle image velocimetry (PIV) measurements. The upstream effect of the wing vortex on the formation of the strake vortex was identified. A dual-vortex structure of the strake vortices was observed before the wing vortex developed. Further downstream, wing and strake vortices rotated around each other slowly initially, and then faster with downstream distance, at an increasing rate with increasing incidence. Prior to vortex breakdown, both wing and strake vortices were found meandering in relatively small regions. The correlation between the instantaneous locations of the vortices increases if the vortices become sufficiently close to each other. The proper orthogonal decomposition (POD) analysis of the instantaneous velocity fields suggested that, for both wing and strake vortices, the most energetic mode was displacement in the first helical mode. The most energetic mode reveals out-of-phase displacements when the vortices are close to each other.

Effect of structure parameters of the flow guide vane on cold flow characteristics in trapped vortex combustor

The cold flow characteristics are investigated to show the effect of the structural parameters of the flow guide vane on the trapped vortex combustor (TVC). The results show that the structural parameters have significant effects on the TVC. As a/H increases, the total pressure loss, the wall shear stress at the bottom of the cavity and the turbulent intensity in the main combustion zone increase. b/B does not have a significant effect on the cavity flow structure and the total pressure loss, and the wall shear stress at the bottom of the cavity increases as increases. There is no significant increase of the turbulent intensity with the increase of b/B. The increase of c/L has little effect on the total pressure loss, and it is not conducive to a stable combustion. As c/L increases, the wall shear stress at the bottom of the cavity decreases. When a/H = 0.4, b/B = 0.4, c/L = 0.1, a desirable dual-vortex structure is formed with an acceptable pressure loss to achieve a stable combustion. Moreover, to ascertain that the flame is stable for different values of Vma with the optimal structural parameters, the effect of Vma on the flow field is discussed. Results suggest that the dual-vortex structure has no relationship with the increase of Vma. Furthermore, an unsteady simulation is conducted to show the generation and the development of the dual-vortex.

Magnetic-field-sensing mechanism based on dual-vortex motion and magnetic noise

In this study, we report two novel field sensing mechanisms using elliptical permalloy single layer. Using micromagnetic modeling, dual-vortex structure is observed and stabilized in elliptical permalloy single layer by applying hard bias field (along the y-axis) and vertical axis field (perpendicular to plane). During the increasing or decreasing of the hard bias field within certain range, the dual vortices would move away from or approach to each other at a constant velocity, leading to a positive correlation between the hard bias field and the vortex gap. By exploring the magnetic noise properties of the elliptical permalloy single layer under various vortex gap, the vortex gap is found to be positively correlated with both the FMR (Ferromagnetic Resonance) peak positions and the integrated thermally excited mag-noise. Therefore, the combination of the dual-vortex motion and the magnetic noise properties make it possible to measure external field (along hard bias direction) through measuring the FMR peak positions or integrated thermally mag-noise. This FMR-peak-based field sensing mechanism and integrated-noise-based field sensing introduce a simple field sensor structure with expected highest sensitivity to 1.1%/Oe and field detectable range over 1000 Oe, which is promising for potential sensor applications.

Reynolds number and aspect ratio effects on the leading-edge vortex for rotating insect wing planforms

AbstractPrevious studies investigating the effect of aspect ratio ($\mathit{AR}$) for insect-like regimes have reported seemingly different trends in aerodynamic forces, however no detailed flow observations have been made. In this study, the effect of$\mathit{AR}$and Reynolds number on the flow structures over insect-like wings is explored using a numerical model of an altered fruit fly wing revolving at a constant angular velocity. Increasing the Reynolds number for an$\mathit{AR}$of 2.91 resulted in the development of a dual leading-edge vortex (LEV) structure, however increasing$\mathit{AR}$at a fixed Reynolds number generated the same flow structures. This result shows that the effects of Reynolds number and$\mathit{AR}$are linked. We present an alternative scaling using wing span as the characteristic length to decouple the effects of Reynolds number from those of$\mathit{AR}$. This results in a span-based Reynolds number, which can be used to independently describe the development of the LEV. Indeed, universal behaviour was found for various parameters using this scaling. The effect of$\mathit{AR}$on the vortex structures and aerodynamic forces was then assessed at different span-based Reynolds numbers. Scaling the flow using the wing span was found to apply when a strong spanwise velocity is present on the leeward side of the wing and therefore may prove to be useful for similar studies involving flapping or rotating wings at high angles of attack.

Effects of sinusoidal leading edge on delta wing performance and mechanism

Lift and drag characteristics of delta wings with low swept angle and various sinusoidal leading edges (SLE) were investigated in a wind tunnel. Three amplitudes and three wavelengths of SLE were tested. It is revealed that, in comparison with the baseline case, when the leading-edge amplitude A ≤5% C (root chord length of a delta wing), the stall of the delta wing can be delayed without penalty on the maximum lift coefficient; meanwhile, the lift-to-drag ratio was kept nearly unchanged. These are beneficial to aircraft maneuverability and agility. Surface oil and hydrogen-bubble flow visualization experiments were further conducted to provide a general view of the underlying flow mechanism of SLE on delta wings. It was found that, for the flow over delta wing with SLE, vortices were generated from every crest of SLE, in contrast to the dual leading-edge vortex structure generated from the apex of the base wing. At high angle of attack, the breakdown of those vortices originating from the crests of SLE may provide additional turbulent kinetic energy to the flow, resulting in the increase of the flow reattachment region on the leeward side, therefore the stall can be delayed.

Transformation of flow structure on a delta wing of moderate sweep angle during pitch-up maneuver

The transformation of flow structure on a delta wing of moderate sweep angle is investigated during pitch-up via quantitative imaging in crossflow planes, in relation to qualitative dye visualization. The comparisons are made with the corresponding stationary wing. The effects of different pitch-up maneuver rates and Reynolds numbers on evolution of flow patterns are also addressed. A technique of high-image-density particle image velocimetry is employed for the quantitative analysis. The transformation of the flow structure starts with intertwining of the dual vortex structure which is distinctively seen at a low angle of attack of 50° swept wing, and then continues with the disappearance of one of the vortices of the dual vortex structure, in particular, the vortex located closest to the leading-edge of the wing. At further stages, a single, large-scale, leading-edge vortex, and distinctive form of vortex breakdown is witnessed. This distinctive type of vortex breakdown turns into a more traditional form at later stages of the maneuver. During the final stage of the pitch-up maneuver, the onset of vortex breakdown moves toward the apex of the wing. The general trend of aforementioned flow transformation happens irrespective of the values of pitch rates and Reynolds numbers. When the foregoing patterns are compared with the corresponding patterns on the stationary wing, it is evident that a significant time lag of development of the flow patterns occurs, relative to the motion of the wing.

Numerical simulation of natural convection in a sessile liquid droplet

Heat transfer in a sessile liquid droplet was studied with numerical methods. A computer code was developed for solving the problem of convection in an axisymmetric hemispherical droplet and in a spherical layer as well. The problem of establishing an equilibrium state in a droplet was solved using several variables: temperature, stream function, and vorticity. Simulation was performed for droplets of water, ethyl alcohol, and model liquids. Variable parameters: intensity of heat transfer from droplet surface, Rayleigh and Marangoni dimensionless criteria, and the characteristic temperature difference. It was revealed that the curve of convective flow intensity versus heat transfer intensity at droplet surface has a maximum. A dual-vortex structure was obtained in a stationary hemispherical profile of liquid droplet for the case of close values for thermocapillary and thermogravitational forces. Either thermocapillary or thermogravitational vortex might be dominating phenomena in the flow structure.

Experimental and Numerical Studies in a Compact Trapped Vortex Combustor: Stability Assessment and Augmentation

Fundamental studies on a compact trapped vortex combustor indicate that cavity injection strategies play a major role on flame stability. Detailed experiments indicate that blow-out occurs for a certain range of cavity air flow velocities. An unsteady RANS-based reacting flow simulation tool has been utilized to study the basic dynamics of cavity vortex for various flow conditions. The phenomenon of flame blow-out at certain intermediate cavity air velocities is explained on the basis of transition from a cavity-stabilized mode to an opposed flow stagnation mode. A novel strategy is proposed for achieving flame stability at all conditions. This involves using a flow guide vane in the path of the main flow to direct a portion of the main flow into the cavity. This seems to result in a desirable dual vortex structure, i.e., a small clockwise vortex behind the vane and large counterclockwise vortex in the cavity. Experimental results show stable flame at all flow conditions with the flow guide vane, and pressure drop is estimated to be within acceptable limits. Cold flow simulations show self-similar velocity profiles for a range of main inlet velocities, and high reverse velocity ratios (−0.3) are observed. Such a high-velocity ratio in the reverse flow shear layer profile leads to enhanced production of turbulence imperative to compact combustors. Reacting flow simulations show even higher reverse velocity ratios (above −0.7) due to flow acceleration. The flame is observed to be stable, even though minor shear layer oscillations are present in the form of vortex shedding. Self-similarity is also observed in reacting flow temperature profiles at combustor exit over the entire range of the mainstream velocity. This indicates that the present configuration holds a promise of delivering robust performance invariant of the flow operating conditions.

Vortical Flow Prediction Validation for an Unmanned Combat Air Vehicle Model

As part of the NATO Applied Vehicle Technology 161 technical group, a study of the aerodynamic behavior of the stability and control configuration wind-tunnel model is presented. Sharp and round leading-edge versions of the model are computed. A validation of Reynolds-averaged Navier–Stokes predictions obtained using two block structured codes are made. Static cases are analyzed and compared with wind-tunnel measurements. The vortical flow features are described in detail for a range of angles of attack. The predictions are in good agreement with the experiments at low angles of attack, whereas for higher angles of incidence (alpha > 15), some discrepancies are seen. A dual vortex structure is present in this region for both leading-edge configurations, resulting in highly nonlinear aerodynamic behavior.

Dual leading-edge vortex structure for flow over a simplified butterfly model

The dye visualization experiments show that a dual leading-edge vortex (LEV) structure exists on the suction side of a simplified butterfly model of Papilio ulysses at α = 8°−12°. Furthermore, the results of particle image velocimetry (PIV) measurement indicate that the axial velocity of the primary (outer) vortex core reaches the lower extreme value while a transition from a “wake-like” to a “jet-like” axial velocity profile occurs. The work reveals for the first time the existence of dual LEV structure on the butterfly-like forward-sweep wing configuration.

Effect of wing planform on leading-edge vortex structures

Flow visualization experiments are conducted in water tunnel for low aspect ratio cropped wings at low Reynolds number. The experimental results show that the model sweep angle Λ influences the formation and development of the leading-edge vortex. For wings with Λ =0°, the dominant flow structure is transverse vortex. When Λ⩾26°, the dual vortex structure can be observed at some angles of attack, and it is confirmed that the dual vortex is a special structure for flow over low aspect ratio wing at low Reynolds number. For Λ⩾56° wings, the dual vortex structure can be observed in a large range of attack angles. Moreover, in comparison with the outer vortex, the breakdown position of the primary vortex is delayed, and the larger the Λ, the later the breakdown location at the same angle of attack.

Experimental Investigations on Leading-Edge Vortex Structures for Flow over Non-Slender Delta Wings

The dye injection and hydrogen bubble visualization techniques are used to investigate the dual-vortex structure including its development, breakdown and the spatial location of vortex core over nonslender delta wings. It is concluded that the dual-vortex structure can be affected significantly by sweep angle and Reynolds number, and generated only at small angle of attack. The angle between the projection of outer vortex core on delta wing surface and the root chord line has nothing to do with the Reynolds Number and angle of attack, but has simple linear relation with the sweep angle of the model tested.