Nanoparticle-based Planar Laser Scattering Research Articles

Experimental Investigation of a Roughness Element Wake on a Hypersonic Flat Plate

An experimental investigation was performed on the wake flow field of an isolated roughness element of a flat plate at Mach 6 by employing the nanoparticle-based planar laser scattering (NPLS) approach. The three-dimensional features and causes of the flow field structure were scrutinized by transient flow field images of roughness elements on various planes. The time-resolved NPLS technique was implemented to examine the time evolution characteristics of the wake flow field of roughness elements. In the following, the process of dynamic evolution of large-scale vortex structures in the wake flow field was methodically assessed. Additionally, the influences of roughness element heights on the wake vortex structure were evaluated and the obtained results were compared.

Generation and Propagation Characteristics of an Auto-Ignition Flame Kernel Caused by the Oblique Shock in a Supersonic Flow Regime

The auto-ignition caused by oblique shocks was investigated experimentally in a supersonic flow regime, with the incoming flow at a Mach number of 2.5. The transient characteristics of the auto-ignition caused by shock evolvements were recorded with a schlieren photography system, and the initial flame kernel generation and subsequent propagation were recorded using a high-speed camera. The fuel mixing characteristics were captured using NPLS (nanoparticle-based planar laser scattering method). This work aimed to reveal the flame spread mechanism in a supersonic flow regime. The effects of airflow total temperature, fuel injection pressure, and cavity length in the process of auto-ignition and on the auto-ignitable boundary were investigated and analyzed. From this work, it was found that the initial occurrence of auto-ignition is first induced by oblique shocks and then propagated upstream to the recirculation region, to establish a sustained flame. The auto-ignition performance can be improved by increasing the injection pressure and airflow total temperature. In addition, a cavity with a long length has benefits in controlling the flame spread from the induced state to a sustained state. The low-speed recirculating region created in the cavity is beneficial for the flame spread, which has the function of flame-holding and prevents the flame from being blown away.

Effects of AoA on tangential three-dimensional film cooling of a slender cone

A supersonic annular supersonic nozzle was used to protect a slender cone from aerodynamic heating at angle of attack (AoA). Experiments were carried out in a Mach 6 wind tunnel. Temperature-sensitive paint (TSP) was used to measure surface temperature change of the cone body; nanoparticle-based planar laser scattering (NPLS) was used to visualize the flow field; oil flow was utilized to monitor surface flow of the cone. Results showed that cooling film can obviously decrease surface heating load in present cases; the coolant layer was able cover a relatively long distance along the model's surface than most other cooling solutions. AoA caused obvious transverse flow on the surface, and it produced much difference heating load between different sides of the cone body. Although the cooling film can mitigate the transverse flow, the windward side was still the hardest part to cool.

Experimental Study of the Unstart/Restart Process of a Two-Dimensional Supersonic Inlet Induced by Backpressure

Unstart/restart phenomena induced by backpressure in a general inlet with a freestream of M = 2.7 are investigated in an in-draft supersonic quiet wind tunnel. The boundary layers are turbulent on the forebody while are laminar on the lip wall, which could mimick real flight conditions. The high-speed Schlieren imaging system and the nanoparticle-based planar laser scattering (NPLS) method are used to visualize the inlet flowfield. The inlet wall pressure is measured by high-frequency pressure transducers. The backpressure is reproduced by downstream transverse jets other than mechanical throttlers, which is more suitable to mimic backpressure caused by combustion. The high spatio-temporal resolution full-view images of inlet flow features during the complete unstart/restart process are captured, which are seldom seen before. The formation and disappearance process of massive boundary layer separation at the entrance of the unstarted inlet is observed. The backpressure transmits upstream through the shock wave/boundary layer interaction (SWBLI) regions. The shock structures change the angles and merge upstream to balance the pressure rise. The Mach shock reflection configuration is observed in both unstart/restart process, accompanied by the boundary layer separation extending to the leading edge. The experiment also revealed notable hysteresis in the unstart/restart process.

Scale effect of gas injection into a supersonic crossflow

This study focuses on the scale effect of transverse gas fuel injected into a Ma=2.95 supersonic flow. NPLS (Nanoparticle-based Planar Laser Scattering) and oil flow techniques were employed to obtain flow structures and jet penetration, while the RANS (Reynolds-averaged Navier-Stokes) model was utilized to investigate fuel mixing characteristics. In order to get the scale effect law, the diameters of jet orifice were chosen as D=2 mm and D=4 mm, and the jet-to-crossflow dynamic pressure flux ratio (J) were set as 2.3, 5.5, and 7.7. The jet penetration boundary could be extracted from NPLS images. By analyzing plenty of NPLS images, a corrected fitting formula was given, which could reflect the scale effect law of jet penetration. The oil flow experiments compared the shapes and the size of the V-shape recirculation zone of different scales. Results showed that the size of the recirculation zone amplifies proportionally with the diameter of the orifice growth. The RANS numerical simulations were conducted to study the mixing characteristics, and the detailed fuel mass fraction concentration and mixing efficiency analysis were given. It is concluded that the turbulent boundary layer plays an important role in the scale effect of the jet penetration and mixing characteristics, especially for the low jet-to-crossflow dynamic pressure flux ratio (J).

Effects of sweep angles on turbulent separation behaviors induced by blunt fin

An experimental study was conducted on turbulent separation behaviors induced by blunt fins with different sweep angles at Mach number 6.0. The Nano-particle based Planar Laser Scattering technique (NPLS) was applied to visualize the flowfield, complemented by pressure tests. Sweep angles of the fins were 10°, 20°,…,60°, with the same leading edge diameter of 10 mm. Fine structures of the interference flowfield induced by blunt fins have been obtained, including the shock systems and vortexes. It was found that the features and shapes of the detached shock depended on sweep angle. When sweep angle < 50°, the detached shock appeared as the form of trailing shock, and the supersonic jet with its reflection could be observed. The detached shock would be curved for the 50° and 60° fins and became a transmitted shock. The Scale-Invariant Feature Transform (SIFT) was successfully applied to obtain the velocity field from NPLS images, and the extent of the separated region was found to decrease with increasing sweep angle. No separation appeared as sweep angle > 30°. Two peak values were detected on the centerline pressure distribution. The first peak did not rely on sweep angle, while the second peak value decreased with increasing sweep angle.

Experimental investigation of supersonic turbulent flow over cylinders with various heights

The unsteady flow structures and pressure fluctuations of the cylinder-induced shock wave/boundary layer interactions (SBLIs) were investigated at Mach 3.4 and 3.8. The cylinders were mounted on the test section floor with a turbulent boundary layer, and six cylinders with various heights have been tested. The nanoparticle-based planar laser scattering (NPLS) flow visualization tests were performed, and the different features of shock systems as well as the "breath" effects of the separation bubble, have been validated by NPLS images. Dynamic pressure transducers were used to characterize the pressure fluctuations, and a remarkable spectra change was noticed around the cylinder foot. With increasing height, the PSD curve of x/D = − 1 ahead of the cylinder shaped from one-peak pattern to two-peaks pattern, while the second frequency peak showed a typical frequency of f ≈32 kHz. It was inferred that the first peak was more influenced by the free end of the cylinder and the detached shock. The second peak pattern of f ≈32 kHz was presumed to be affected by the vortex breakdown near the cylinder foot through analysis of flow visualization. Low frequency around 0 ~ 3 kHz was found to have an obvious rise at x/D = − 2, close to the typical frequency of SBLIs.

Turbulent boundary layer subjected to a sonic transverse jet in a supersonic flow

Particle Image Velocimetry (PIV) technology, oil flow visualization and Nanoparticle-based Planar Laser Scattering (NPLS) technology are used to investigate the turbulent boundary layer (BL) of a sonic jet injected into a supersonic flow at Ma=2.95. Instantaneous and averaged NPLS images display detailed structures of the bow shock, the horseshoe vortex, separation zones and the collision shock. Root Mean Square (RMS) values of flow quantities calculated from NPLS images and PIV data reveal that turbulent fluctuations are suppressed downstream of the horseshoe separation zone in the jet lateral region and a 33%-35% reduction of the density fluctuation magnitude across the boundary layer is observed, while it is enhanced in the jet wake by 34%-50%. Further away from the wall, the influences of the λ shock system, the expansion process, and the reflected shock on turbulence intensity and velocities of the crossflow BL is weakened. A three-dimensional schematic diagram of the interaction between the sonic jet and the turbulent BL is presented, explaining the formation of the turbulence decay zone in the jet lateral region.

Instability of an asymmetric supersonic planar wake

Supersonic planar wake has strong instability characteristics due to its complex flow component, such as wake/shear layer, attachment, and complex shock waves interferences. In this experimental study, an asymmetric supersonic planar wake, with Mach numbers of 1.37 and 2.39 on the upper and lower side, is investigated by Nanoparticle-based Planar Laser Scattering (NPLS) technique. The blending thickness of the re-development mixing layer and the position of the re-attachment point are analyzed. Results show that the flow phenomenon of the re-development mixing layer is unsteady. The flow state of the mixing layer can be small scale structures or large scale coherent structures. The distribution of re-attachment points in the polar map shows the spatial unsteadiness of near-wake. Besides, the streamwise instability is dominant. The streamwise span of the re-attachment points is consistent with the thickness of the plate. By correlation analysis, the flow state of the mixing layer is not sensitive to the position change of the re-attachment point.

Experimental study of parallel injections with different distances into a supersonic crossflow

The flow physics aspects of parallel sonic transverse gaseous injections in a Ma = 2.95 supersonic crossflow are experimentally investigated in this study. NPLS (Nanoparticle-based Planar Laser Scattering) and oil flow techniques are employed for qualitative comparison. Under the supersonic crossflow with jet-to-crossflow momentum flux ratio (J) of 7.7, the NPLS images reveal the shock structures, K–H vortex and slip lines on different spanwise planes of each of the experimental conditions, and the oil flow experiments illustrate the bow shocks, separation zones and other interactions in the near-wall region. Furthermore, the fractal dimensions of the jets mainstream region are given to analyze the turbulent intensity. The interaction mechanism of the flow field is completely different with the distances of the orifices growing. As the orifices are close together, the flow structures of three orifices, including bow shocks, separation zone, horseshoe vortex and jet mainstreams, almost merge into one. With the distances between jets growth, the bow shocks would merge into a normal shock and the separation zones interact with each other. When the distances between jets are large enough, the jet mainstreams are independent, while the bow shocks interact with each other forming a complex shock system. The interactions between jets promote the development of turbulence in general, while the interactions between too close jets would restrict the development of turbulence. In addition, the horseshoe vortex and the shedding vortex are analyzed.

Supersonic boundary layer transition induced by self-sustaining dual jets**Project supported by the National Natural Science Foundation of China (Grant Nos. 11602299, 11872374, and 51809271).

To promote high-speed boundary layer transition, this paper proposes an active self-sustaining dual jets (SDJ) actuator utilizing the energy of supersonic mainflow. Employing the nanoparticle-based planar laser scattering (NPLS), supersonic flat-plate boundary layer transition induced by SDJ is experimentally investigated in an Ma-2.95 low-turbulence wind tunnel. Streamwise and spanwise NPLS images are obtained to analyze fine flow structures of the whole transition process. The results reveal the transition control mechanisms that on the one hand, the jet-induced shear layer produces unstable Kelvin–Helmholtz instabilities in the wake flow, on the other hand, the jets also generates an adverse pressure gradient in the boundary layer and induce unstable streak structures, which gradually break down into turbulence downstream. The paper provides a new method for transition control of high-speed boundary layer, and have prospect both in theory and engineering application.

Characteristics of mixing enhancement achieved using a pulsed plasma synthetic jet in a supersonic flow

Supersonic mixing layers exist extensively in supersonic engineering applications. The rapid mixing of fuel and oxidant at short distances is of great importance, but makes it difficult to develop efficient propulsion systems. The plasma synthetic jet (PSJ) is regarded as a promising high-speed flow control technique. The characteristics of mixing enhancement achieved using a pulsed PSJ were investigated via experiments. Results showed that the PSJ is an effective method for mixing enhancement. Nanoparticle-based planar laser scattering (NPLS) was used to obtain flow structures in three directions. The velocity fields near the PSJ actuator orifice were measured by particle image velocimetry (PIV). Indexes of the fractal dimension and mixing layer thickness were applied to estimate the effect of the PSJ actuator on the supersonic mixing layers. The large-scale vortex structures induced by the pulsed PSJ in the supersonic mixing layers were successfully captured by NPLS. The effect of the PSJ on the supersonic mixing layers was remarkable. The mixing layer thickness under perturbation was larger than that under no perturbation in the downstream. The distribution of the fractal dimension suggests that perturbation of the PSJ cannot improve the fractal dimension values of the fully developed supersonic mixing layers.

Flowfield structures of pylon-aided fuel injection into a supersonic crossflow

Flowfield structures of pylon-aided fuel injection of a sonic jet into a supersonic crossflow with Ma = 2.95 have been investigated by Nanoparticle-based Planar Laser Scattering (NPLS) technique and oil flow technique, separately. Two experimental cases have been done for jet-to-crossflow momentum flux ratios (J) of 7.7 and 20.6, respectively, and the results are compared with those obtained by normal injection of a single jet from a flat plate without a pylon. Experimental results indicate that the flow structures of pylon-aided injection cases are very similar to those found in the normal injection cases in the downstream region of the jet. Upstream of the jet orifice, the detached shock generated at the leading edge of the pylon is identified by inserting with the separation shock. For lower injection pressure case there is always larger influences of the pylon in the flow structures in terms of jet plums' position and the shock intensity. Besides, the mechanism of mixing enhancement resulting from the pylon is also elaborated in the present work.

Mixing enhancement for supersonic mixing layer by using plasma synthetic jet

Mixing enhancement for supersonic mixing layer is of great importance for developing scramjet engine. The growth rate of supersonic mixing layer is smaller than that of subsonic mixing layer. As the compressibility increases, the mixing enhancement becomes more difficult. Plasma synthetic jet is regarded as a promising flow control technology. The plasma synthetic jet generator can produce high energy jet. This generator has no moving parts and does not need additional gas source. It is the first time that plasma synthetic jet has been used to enhance the mixing in supersonic mixing layers. The influence of plasma synthetic jet on the supersonic mixing layer is investigated experimentally and numerically. The experiments are conducted in the low noise supersonic mixing layer wind tunnel. The Mach number of upper stream and lower stream are 1.37 and 2.39 respectively. The convective Mach number of this wind tunnel is 0.32. The plasma synthetic jet actuators are installed in the splitter plate. The distance between the jet hole and the splitter plate end is 15 mm. The nanoparticle-based planar laser scattering (NPLS), particle image velocimetry (PIV) and schlieren are used to obtain the response of the supersonic mixing layer to single pulse plasma synthetic jet perturbation. The NPLS successfully captures the large-scaled vortex structures induced by the plasma synthetic jet in the supersonic mixing layers. The effect of plasma synthetic jet is remarkable. The schlieren images show the process of the perturbation. An oblique shock wave is generated when the jet is ejected. The PIV is employed to obtain the influence of plasma synthetic jet on the velocity field. The &lt;i&gt;y&lt;/i&gt;-velocity standard deviation increases due to the perturbation. The actuators’ mixing enhancement effects and actuators’ performances at three locations are investigated by two-dimensional numerical simulation. The three actuators are located on the upper, bottom and end surface of splitter plate respectively. The numerical simulation results show that the mixing layer thickness is increased by the plasma synthetic jet perturbation. There are two mechanisms of perturbations while actuators are located at different positions. The actuators installed on the upper and bottom surface of splitter plate influence the mixing layer through perturbing the upper and lower stream respectively. The actuator installed at the end of splitter plate affects the mixing layer directly. The response time of supersonic mixing layers to the perturbation of the actuator installed at the end of splitter plate is shorter than those of the others. The performance of each actuator is sensitive to the location.

Investigations on flow and growth properties of a cavity-actuated supersonic planar mixing layer downstream a thick splitter plate

Mixing enhancement is vital to combustion efficiency in a rocket-based combined cycle engine. Presently, we propose a cavity as an actuator and investigate the cavity-actuated supersonic mixing layer experimentally using particle image velocimetry (PIV) and nanoparticle-based planar laser scattering (NPLS). Large eddy simulation (LES) is conducted to obtain the detailed flow structures and pressure spectra of the supersonic mixing layer. Results indicate that the vortex evolution of the supersonic mixing layer is locked strongly to the excitation frequencies from the self-sustaining oscillation of supersonic cavity flow in both the near and far flowfields. Multiple coherent structures of different sizes coexist in the initial flowfields, and then only large ones survive downstream. Pressure spectra can reflect the vortex evolution entirely. The third mode of the pressure spectrum gradually becomes dominant due to having the highest energy content. Curvature shocklets are observed in this case at the low convective Mach number of 0.22. The growth rate in the linear process is slightly larger than that of the case without a cavity while the velocity thickness is 3.34 times as large as that. The normalized growth rate sees a massive increase and is 8.95 times as large as the typical result of a free supersonic shear layer.

Impact of pure favorable pressure gradient on a supersonic flat-plate turbulent boundary layer

By employing the particle image velocimetry and Nanoparticle-based Planar Laser Scattering method, the impact of streamwise favorable pressure gradient on the mean and turbulent characteristics of a Mach 2.95 turbulent boundary layer is experimentally investigated. Through a careful arrangement of the experiment, the possible influence of streamline convex curvature is minimized to have a flat-plate streamwise favorable-pressure-gradient boundary layer. While the log law is found to be well preserved at all streamwise positions, the wake region is weakened by the pressure gradient. Different from the boundary layers formed over the convex wall and over the sudden expansion ramp, both principal strain rate and bulk dilatation in the near-wall region are found to be barely changed along the streamwise direction in the flat-plate favorable-pressure-gradient boundary layer. Because of this, the impact of the favorable pressure gradient on the near-wall turbulent fluctuation is also found to be insignificant, which differs from the supersonic convex boundary layer where the near-wall turbulence is found to be greatly suppressed.

Structures of near-wall wakes subjected to a sonic jet in a supersonic crossflow

Nanoparticle-based Planar Laser Scattering (NPLS) technology and Oil flow visualization are employed to investigate the wake flow of a sonic jet injected into a supersonic crossflow at Ma = 2.95. Experiments are run for seven jet-to-crossflow momentum flux ratios (J) of 2.3, 5.5, 7.7, 11.2, 16.0, 20.6, and 28.9 of a sonic jet injected into a supersonic crossflow at Ma = 2.95. Experimental results suggest that the near-wall wake zone could be divided into three regions: the V-shape separation region behind the jet, followed by the reattachment region and the mixing and recovery zone further downstream. Lower jet-to-crossflow momentum flux ratio cases always have shorter distance from the orifice to the interaction position of the V-shape collision shock and the reflected shock. The angle between the main separation lines is found to be independent of J. Correlations for predicting the separation length and width are proposed based on the experiments of different momentum flux ratios.

Shock wave structures in the wake of sonic transverse jet into a supersonic crossflow

This study experimentally and numerically investigated the flow physics of a sonic transverse gaseous jet under a Ma = 2.95 supersonic crossflow. NPLS (Nanoparticle-based Planar Laser Scattering) and oil flow techniques were combined to achieve experimental visualization, while RANS (Reynolds-averaged Navier-Stokes) model was used to carry out the numerical simulation. Under the supersonic crossflow with jet-to-crossflow momentum flux ratio (J) of 7.7, the typical structures, including a bow shock, a barrel shock, horseshoe vortex, and separation zones, were clearly observed by the NPLS techniques. V-shape separation bubbles and V-shape collision shocks were identified based on the oil-flow results. Detailed flow fields around the V-shape separation bubble were then revealed by numerical calculations. It is shown that the collision shock induced V-shape separation bubbles can further affect the Mach disk, the reflected shock and the barrel shock. The reflected shock deflects the collision shock toward the two flanks. The collision shock intersects with the barrel shock, influencing the shape of the Mach disk. In addition, the interaction between the reflected shock and the wall leads to the velocity slip in the near-wall region.

Investigation of large-scale structures in supersonic planar non-reactive shear layer with a base flow region

Large-scale structures in supersonic planar non-reactive shear layer with a base flow region are experimentally investigated. By utilizing the recently developed nanoparticle-based planar laser scattering (NPLS) method, the fine transient streamwise and spanwise flow fields are acquired, and the large-scale structures in the reattachment region, initial turbulent shear region, and developing turbulent shear region are analyzed. The predominant instability is convective for the two separated shear layer, whereas the instability in the base flow zone is still absolute according to the linear instability analysis. From the initial to developing turbulent shear layer region, the predominant instability transited gradually from absolute to convective instability, and the corresponding scales of the large-scale structures turn from about half to whole of the shear layer thickness. The entire streamwise mean flow topology and the streamwise and transverse velocity fluctuations are compared with different pressure ratios and with incompressible blunt trailing-edge shear layer. Linear instability analysis shows further evidence to the instability evolution of the supersonic shear layer with a base flow along the streamwise. It implied that instability induced by unmatched pressure in the wake zone can increase the absolute instability and improve the receptivity to inflow perturbations. The reattachment points are approximated with the intersection points of the reattachment shock pairs, which can be obtained from the transient NPLS images. The distribution of the reattachment points suggests that: the principal axis of the reattachment point distribution is dominated by the high-speed flow, and unmatched pressure increases the perturbation in the wake zone, result in a high absolute instability.

Fine flow structure and mixing characteristic in supersonic flow induced by a lobed mixer

In a supersonic suction type of mixing layer wind tunnel, by employing nanoparticle-based planar laser scattering (NPLS) method, contrast experiments are carried out with the emphasis on the fine flow structures of planar mixing layer and the mixing layer induced by triangular lobed mixer. The normal-shock equation, isentropic equation and sound speed relationship are utilized to calculate the flow parameters. The calculated Mach numbers are 1.98 and 2.84 for upper and lower airstreams respectively with a convective Mach number of 0.2. The NPLS images clearly shows the Kelvin-Helmholtz vortices, streamwise vortices, shock waves and the pairing processes of large-scale vortex structures. The unsteady properties of development and evolution for large-scale vortices are obtained by contrasting the NPLS images at different times. Also, it has been demonstrated by the present experimental investigation that in supersonic mixing layer with low convective Mach number, the small shock waves are still existing. These small shock waves that occur have negative effects on the mixing process. It is because the convection flow process of upper and lower airstreams is non-isentropic, causing the total pressure to lose. Based on the NPLS results, flow structures and mixing characteristics are analyzed quantitatively by using fractal and intermittency theory. The results show that the mixing efficiency increases obviously with the introducing of large-scale streamwise vortices. The nibbling of vortex clusters induced by large-scale streamwise vortices obviously increases the interface area of mixing. Meanwhile, compared with planar mixing layer, larger spanwise structures roll up in triangular lobed mixing layer, leading to more entrainment of upper and lower airstreams. In the present investigation of supersonic planar mixing layer, the value of fractal dimension of fully turbulent region is stable at 1.55-1.6. Whereas the value of fractal dimension for triangular lobed mixing layer reaches 1.88 at the flow field far away downstream, which breaks through the value of fully developed turbulence for planar mixing layer. Besides, in triangular lobed mixing layer, the shear action between streamwise vortices and spanwise structures plays a leading role in promoting mixing. The mixing flow shows the property of apparent crushability and three-dimensional behavior, which plays a positive role in promoting mixing at a scalar level. The analysis of intermittency indicates that the interaction between streamwise and spanwise vortices dominates the mixing characteristics, and due to the entrainment of streamwise vortices, the mixing region induced by triangular lobed mixer becomes larger, and more fluids are engulfed into the mixing region to complete the mixing process.