Measurement of shock structure and mixing enhancement via passive control in a Mach 1.86 square jet

Mixing enhancement in a subsonic-supersonic shear layer with a cavity splitter plate

Better mixing in the near-field region of jets with their surrounding fluid is of high interest for several industrial applications. Passive control that involves jet geometry modifications as compared to the traditional circular design is used in the present work. An analysis of the entrainment mechanism in the near jet-exit field is proposed for innovative hemispherical nozzles (circular and six-lobed). High-speed Time-Resolved Particle Image Velocimetry (TR-PIV) measurements are used to experimentally characterize the entrainment mechanism in these jets. The distributions of mean entrainment rates, shear layer growth, and momentum flux are investigated along the longitudinal direction within the near-field region of both circular and lobed hemispherical jets. Significant entrainment enhancement is found using the hemispherical geometry as a passive-control method. By comparing both investigated hemispherical nozzle geometries, it has been demonstrated that the lobed nozzle provides higher mixing rates compared to the circular jet. This enhancement in mixing can be attributed to the stronger streamwise vortex structures generated by the lobed nozzle geometry, which promote increased entrainment of the surrounding fluid.

A review on enhanced mixing methods in supersonic mixing layer flows

This experimental study examined the effects of differing levels of passive fuel-air premixing on flame structures and their associated NO x and CO emissions. Four alternative fuel injector geometries were explored, three of which had lobed shapes. Prior experimental studies of two of these lobed injector flowfields focused on nonreactive mixing characteristics (Smith, L. L., Majamaki, A. J., Lam, I. T., Delabroy, O., Karagozian, A. R., Marble, F. E., and Smith, O. I., Mixing Enhancement in a Lobed Injector, Physics of Fluids, Vol. 9, No. 3, 1997, pp. 667-678) and emissions measurements in the absence of air confinement [Mitchell, M. G., Smith, L. L., Karagozian, A. R., and Smith, O.I., Burner Emissions Associated with Lobed and Non-Lobed Fuel Injectors, Twenty-Seventh Symposium (International) on Combustion, The Combustion Inst., Pittsburgh, PA, 1998, pp. 1825-1831]. The present studies examined the effects of confinement of the crossflow to reduce the local equivalence ratio as well as the effects of altering the geometry and position of the flameholders to further influence passive fuel-air premixing. NO x and CO emissions as well as flame photographs and planar laser-induced fluorescence imaging of seeded acetone were used to characterize injector performance and fuel and flame evolution. It was found that, with significant air confinement, forcing a more intimate mixing between fuel and air before ignition and flameholding, both NO x and CO emissions could be simultaneously reduced under the same operating conditions via this passive flow control technique.

M IXING enhancement in jet flow has long been the topic of extensive research due to its important applications in aircraft propulsion and combustion. It also has military importance when mixing enhancement is used to suppress the infrared red signature for the fighter jet exhausts. Many passive or active approaches or techniques to enhance jet mixing have been developed and explored, and some of these techniques had also been reviewed and summarized by Gutmark and Grinstein [1]. In passive control strategies, the tabs or vortex generators in particular are simple but effective techniques in flow control. Bradbury and Khadem [2] were the pioneers to study the effect of solid tabs on jet flow. They found that the nozzle boundary-layer thickness, turbulence level, and convergent ratio did not have very strong influence on the jet development. On the other hand, to insert small rectangular tabs into the jet flow on the nozzle perimeter would induce some profound effects on the jet development. Since then, many researchers have been working on this technique. Notable examples include Samimy et al. [3], Reeder and Samimy [4], Zaman et al. [5], Yu and Koh [6], Paoli et al. [7] and many others. Zaman [8] conducted a systematic test on a series of nozzles with various shapes or orifices and found that the spreading of most asymmetric jets was not much different from that of a round jet, but the biggest increase in jet spreading was observed with the tabs. However, the penalty of introducing the tabs is also found to be significant. The thrust lost due to the tabs varies from 4.1 to 23.7%when theflowblockage (proportional to the facing area of tabs) increases from 1.1 to 14.1%. Tominimize the thrust lost caused by tabs but yet maintain its effectiveness in mixing enhancement is important to practical applications. Researchers have been searching for new techniques with better effects on mixing but minimum penalty on thrust. This Note presents results using the socalled air-tab technique. The air-tab technique is achieved by injecting a small amount of air (less than 1% of the volume flow rate of the primary jet) into the plume of the primary jet at choked speed and at certain attacking angles (45 and 90 deg with respect to the primary jet direction). The air-tab technique shows minimum or no impact on thrust, but it would be able to provide significant effect as the solid tab on the mixing enhancement. A similar idea had been used in air and fuel mixing: for example, the experiments conducted byMilanovic and Zaman [9,10]. The description of the experimental setup is presented subsequently. It will be followed by the results and discussion. The Note ends with brief concluding remarks.

In this thesis, the flow physics in low-compression scramjets, in particular inlet-fueled radical farming scramjets, are investigated using high fidelity numerical simulations. The two-dimensional inlet-fueled radical farming scramjet engine employed by McGuire provides the basis for this thesis. It is used to numerically investigate the flow physics that govern the mixing, ignition and combustion processes. The focus lies in particular on the effect that unsteady flow features and turbulent structures have on the scramjet performance. Therefore, a high fidelity numerical method, wall-modeled large-eddy simulation WMLES, is employed to resolve those effects. For comparison, the Reynolds-averaged Navier-Stokes RANS approach, which is commonly used for large scale scramjet simulations, is employed as well. Furthermore, the scramjet simulations performed here include finite rate chemistry, using the modified JetSurf 2.0 model, and thermal non-equilibrium effects, using the two-temperature model. This thesis represents a contribution to the knowledge of scramjet processes with particular new understanding of: the flow physics governing the enhanced mixing process in inlet-fueled scramjets; the combustion regimes present in radical farming scramjets; the effects of thermal non-equilibrium on scramjet flow structures and combustion performance. This thesis presents a detailed investigation of governing flow physics relevant to inlet injection. Experimental campaigns have shown that inlet-injection significantly improves the scramjet performance, a result of its mixing enhancing capabilities. In the past, researchers argued that increased mixing lengths on the inlet cause the mixing enhancement. More recently, it was shown numerically that fuel plume/flow structure interactions at the entrance of the combustor are responsible for the changing mixing process. However, no study has been conducted that explores the details of those interactions and their sensitivity to other scramjet related flow features. Therefore, a detailed investigation is performed that analyses the governing flow physical processes that are responsible for mixing enhancement. It will be shown that for maximized mixing enhancement, flow structures within the engine have to interact with the inlet injected fuel plume in such a way that the plume splits in half as it is compressed towards the combustor wall, thus increasing the effective mixing area and with it the rate of mixing. Utilizing such an approach, it is shown that the fuel-air mixing rate can be increased by a factor of up to five as the flow passes through the structures at the combustor entrance. Furthermore, the ignition and combustion process within the scramjet engine is investigated in detail. It was found that radicals produced around the fuel jets aid the ignition process. The gas mixture ignites initially at the combustor entrance near the sidewall. Further downstream, with the second shock impingement at the lower combustor wall the gas mixture ignites around the unburned fuel plumes near the scramjet center. As the combustion process proceeds, it becomes mixing limited towards the back of the combustor, where mixing and combustion efficiencies for this specific configuration reach values of 71\% and 61\%, respectively, at the combustor exit. Identifying combustion regimes present in radical farming engines provides valuable information to the scientific community as it will aid the development of turbulence-chemistry interaction models, which are necessary to accurately model combustion processes in turbulent flows. It will be shown that a wide spread of combustion regimes is relevant for this type of engine. In fact, it will be shown that future turbulence-chemistry interaction models should have the capability to accurately represent partially-premixed and non-premixed gas mixtures, whose combustion regimes range from distributed reaction zones to thin reaction sheets. The effect of thermal non-equilibrium on flow structures and the combustion process is analyzed as well, which is of particular interest for shock tunnel testing as the nozzle outflow is in a state of thermal non-equilibrium. It will be shown that the thermal state of the scramjet inflow has only a weak influence on shock structures in the engine, while thermal non-equilibrium modeling within the scramjet has a larger effect. Temperature distributions in hot regions that develop near the scramjet wall are, however, strongly influenced by thermal non-equilibrium effects, in particular for the un-fueled engine. For hydrogen fueled scramjets thermal non-equilibrium effects become negligible downstream of the combustor entrance, as the relaxation process between air and hydrogen is drastically enhanced compared to un-fueled simulations. Nevertheless, thermal non-equilibrium affects the radical production around the jet plumes on the inlet, which influences the ignition process further downstream. The insights provided by this thesis increases the understanding of flow physics in radical farming engines significantly, which allows us to improve future scramjet designs. Furthermore, the relevance of certain flow phenomena is now better understood, thus providing more information to assess the numerical modeling of such phenomena.

Turbulence management can be accomplished via manipulation of large-scale coherent structures (CS) by controlling their generation, growth and interactions. By turbulence management, we mean either enhancement or suppression of turbulence and related phenomena. Since accumulated evidence so far suggests that turbulent shear flows are characterized by CS, it is evident that turbulence management can be achieved in all turbulent shear flows. Obviously, such turbulence management can yield profound potential benefits in controlling turbulence phenomena, e.g., enhancement of mixing, heat transfer and chemical reaction (including combustion), and reduction of drag and aerodynamic noise.

The current study provides numerical investigation into the use of “Hartmann whistle” as an effective passive flow control device by covering the major area between the nozzle exit and cavity inlet using a cylindrical shield. The passive control is accomplished by allowing the pulsating jet to exit through two small openings in the shield so that it can be utilized for various flow control applications such as mixing enhancement, drag reduction, noise mitigation. The current study numerically investigates the effect of partially covered cylindrical shield on the shock as well as regurgitant oscillation characteristics of a Hartmann whistle when the pulsating jet exits through the two small openings of the cylindrical shield. The relevant parameters that modify the flow/shock oscillations of the Hartmann whistle are the cavity standoff distance, nozzle pressure ratio, cavity length, cavity shield, etc. The studies were performed for various standoff distances values of 10, 20, and 30 mm to demonstrate the role of standoff distance in effective flow control. The modifications in the shock as well as regurgitant oscillation features of partially covered Hartmann whistles are systematically compared using transient velocity vectors, Mach number contours, etc. for various standoff distances. The velocity vectors indicate flow diversion features near the cavity mouth as well as inflow and outflow jet regurgitant phases. The Mach contours of partially shielded Hartmann whistles indicate shock structures, zones of flow deceleration and re-acceleration. It also clearly demonstrates that the resonant oscillations are primarily driven by jet regurgitance at smaller standoff distances, but at higher standoff distances they are primarily driven by the fluid column oscillations in the shock-cells, shield as well as in the cavity zones. Thus, the current study reveals that the standoff distance is a crucial parameter that controls the strength of shock, regurgitant as well as fluid column oscillations in a partially covered Hartmann whistle in order to achieve an effectual flow control.

The effect of internal grooves cut along the inner surface of the diverging portion of a Mach 1.8 convergingdiverging nozzle on the characteristics of an axisymmetric jet was investigated experimentally. Decay, growth, and noise suppression characteristics of supersonic jets from plain nozzle, nozzle with two semicircular grooves, and nozzle with two square grooves are presented. The grooves act as effective passive controls, resulting in significant enhancement of jet mixing. The shock cell structure from grooved nozzle is weaker than that of plain nozzle. Acoustic measurement was taken in the nozzle exit plane and in the far field. In the grooved plane grooves show a definite advantage in terms of jet noise attenuation. However, in the plane normal to the grooved plane they are not effective in screech suppression. Further, the present results authenticate that nozzle pressure ratio plays an important role in the case where an adverse pressure gradient exists near the nozzle exit. Nomenclature M = nozzle-exit Mach number Pa = ambient pressure Pc = jet centerline pitot pressure P0 = stagnation pressure in the settling chamber X = coordinate perpendicular to the nozzle-exit plane Y = coordinate parallel to the grooved plane Z = coordinate normal to the grooved plane I. Introduction T HE passive control scheme investigated in this study is based on the modification of the boundary layer, growing along the nozzle inner walls achieved through partial notches. Streamwise vortices generated by the notches cut at the nozzle exit have been demonstrated to be effective in jet noise reduction. 1 These vortices provide the necessary secondary instabilities that aid the faster amplification of the primary instabilities and hence the growth of the coherent structures. In effect, the streamwise vortices bring in threedimensionality to the otherwise, basically, two-dimensional spanwise organized vortical structures (coherent structures). Thus the evolution of the large-scale structures gets altered by the notches, which, in turn, alter the mixing and acoustic characteristics of the jet. In view of the effectiveness and simplicity of the notched axisymmetric nozzles, they were investigated in the present study. Similar studies have been conducted by many researchers. 1−3 Pannu and Johannesen 1 investigated underexpanded jets issuing from notched nozzles. The centerline pitot-pressure data indicated that the shock cell structure was modified and the jet decayed faster than the unnotched nozzle flow beyond the core region. They demonstrated that the dominant feature of the flow which determined the structure far downstream was the trailing vortices shed from the swept edges of the notches. They concluded that the notches were effective silencers, mainly because they caused the noise sources to be surrounded by a broad region of low-speed turbulent flow. Smith and Hughes 2 presented experimental results obtained for jets from notched nozzles in a coflowing freestream. The results showed

Geometric scale effect of the subsonic-supersonic shear layer based on a sinusoidal lobed splitter plate

: The subject research program sought to explore a novel method for achieving passive flow field control, with applications to mixing enhancement and noise reduction, through the interaction of the flow with flexible filaments. The filament was attached to a jet centerline where it was allowed to interact with and modify the large-scale structures in the flow. This flow control resulted in improved mixing, lower noise, and a more stable flow. The filament was shown to be particularly effective in suppressing screech in underexpanded supersonic jets. The study was successful in identifying an optimal filament configuration, which was used to obtain attenuation levels as high as 32dB. An investigation of temperature effects indicated that the filament performance was actually enhanced with moderate temperature increase. Sound field mappings revealed that the filament created dramatic changes in the sound field. Finally, flow field analysis of the supersonic jet revealed that the filament was successful in extensively modifying the structures in the exhaust plume, which can provide a physical explanation for the measured noise reduction. The researchers feel that this program was successful, as the results prove that the filament effectively attenuated jet noise in both supersonic and subsonic flows.

*This paper presents experimental results on shock structure, flow separation, and mixing enhancement of ambient dry air in rectangular over-expanded supersonic nozzles. A total of 44 symmetrical or asymmetrical nozzles with three different aspect ratios, namely 2.78, 3.57, and 5.00 varying in expansion ratio between 1.00 and 2.10 and varying in nozzle exit angles between 0 and 4.8 degrees, were investigated. Most results presented in this study pertain to nozzles with an aspect ratio of 5.00. Diagnostic tools utilized in this research included primarily Schlieren photography to capture the instantaneous flow structure in the nozzle and the jet plume as well as nozzle reservoir pressure measurements. The study indicates that mixing enhancement in the plume of a jet can be correlated to the shock structure in the nozzle. The shock structure is determined by the nozzle geometry and flow conditions. Nozzles with the most promising mixing enhancement characteristics were symmetrical and had half-nozzle exit angles of at least 2.6 degrees while operated at nozzle pressure ratios between 1.4 and 2.5, depending on the aspect ratio. In all cases with severe plume instability, the Mach stem with asymmetric incident and reflective shocks, needed to be placed at an area greater or equal to 1.2 times the sonic area.

The passive control of a shock wave-boundary-layer interaction involves placing a porous surface beneath the interaction, allowing high pressure air from the flow downstream of the shock wave to recirculate through a plenum chamber into the low pressure flow upstream of the wave.The simple case of a normal shock wave at a Mach number of 1·4 interacting with the turbulent boundary layer on a flat wall is investigated both experimentally and numerically. The experimental investigation made use of holographic interferometry, while the computational section of the investigation made use of a Navier-Stokes code to derive pressure gradients, boundary-layer properties and total pressure losses in the interaction region. It is found that the structure of shock wave-boundary-layer interactions with passive control consists of a leading, oblique shock wave followed by a lambda foot. The oblique wave originates from the upstream end of the porous region, and its strength is determined by the magnitude of the local blowing velocities. The shape of the lambda foot depends on the position of the main shock relative to the control region, resembling an uncontrolled foot when the main shock wave is towards the downstream end of the porosity, but becoming increasingly large as the shock moves upstream and eventually merging with the leading, oblique shock to form a single, large, lambda structure.Improved forms of passive control are suggested based on the findings of this investigation, including the use of passive control systems which incorporate streamwise variations in the level of porosity.

This paper presents the numerical analysis of a convergent-divergent circular nozzle with the exit Mach number of 1.69 with and without passive control at the exit. The passive control method opted for this analysis was inward and outward ascending triangular protrusion. This paper explores the influence of the passive control geometry and its blockage area concerning the nozzle exit. The nozzle pressure ratio (NPR) used for carrying out the flow analysis were 3, 4.932, and 6. Two different inward and outward protrusions were used with a height of 1.5 mm and 3 mm. From the results, the potential core length of the protrusion 1.5 mm height was not much changed in the both outward and inward cases. But when the height of the protrusion was increased to 3 mm, there was a noticeable core length reduction at all NPR but with different cases. At the NPR of 6, the potential core length of the inward protrusions 3 mm was reduced by 44 % compared to the plain CD nozzle.

This research investigated a passive flow control technique to mitigate the adverse effects of shock wave–boundary layer interaction on a NACA 0012 airfoil. A perforated plate with a strategically positioned cavity beneath the shock wave anchoring spot was employed. Airfoils with perforated plates of varying orifice sizes (ranging from 0.5 to 1.2 mm) were constructed using various manufacturing techniques. Experimental analysis utilized an “Eiffel”-type open wind tunnel and a Z-type Schlieren system for flow visualization, along with static pressure measurements obtained from the bottom wall. Empirical observations were compared with steady 3D density-based numerical simulations conducted in Ansys FLUENT for comprehensive analysis and validation. The implementation of the perforated plate induced a significant alteration in shock structure, transforming it from a strong normal shock wave into a large lambda-type shock. The passive control case exhibited a 0.2% improvement in total pressure loss and attributed to the perforated plate’s capability to diminish the intensity of the shock wave anchored above. Significant fluctuations in shear stress were introduced by the perforated plate, with lower stress observed in the plate area due to flow detachment from cavity blowing. Balancing shock and viscous losses proved crucial for achieving a favorable outcome with this passive flow control method.