Pseudo-shock Wave Research Articles

Effects of isolator length on pseudo-shock wave in a rectangular duct

Effects of isolator length on pseudo-shock wave in a rectangular duct

Research on combustion characteristics of the strut-equipped variable geometry scramjet combustor at Mach 2 condition

In this paper, the unstart boundary of the strut-equipped variable geometry combustor was experimentally verified at the Mach 2 condition. Limited by the flow field structure at a low Mach number, the working boundary of the supersonic combustor had certain requirements on the fuel equivalent ratio and combustor expansion ratio. By simulation analyses and static regulations, the optimization of flow field structure and the successful ignition of the strut-equipped supersonic combustor were achieved. To adapt the fuel equivalent ratio with the expansion ratio for performance enhancement, experimental investigations were conducted with Mach 2 condition, where the fuel equivalent ratio ranges from 0.6 to 1.0, and the combustor expansion ratio ranges from 2.0 to 3.43. The conclusions were drawn from qualitative summaries and quantitative analyses. Under the static flow field and a low expansion ratio condition, the starting boundary of the combustor fuel equivalent ratio was restricted to the equivalent ratio interval from 0.6 to 0.8, which can be broadened by the ignition effect. In the process of equivalent ratio upgrading, the high-level heat release intensity required larger expansion ratios for heat exchange, meanwhile, the increased pseudo-shock wave length improved the unstart risk. A stepwise performance enhancement at the 0.8 equivalence ratio was achieved by dynamic regulation, and the flame boundary integration conjecture based on the previous theoretical analyses was proposed for this phenomenon. The performance distribution pattern of the combustor was obtained from the experiments, which showed that the high-performance range is mostly concentrated around the 0.8 equivalence ratio.

Effect of angle of attack on the performance of the supersonic intake for High Mach Integrated Control Experiment (HIMICO)

Experimental, numerical and theoretical investigations are carried out to grasp the effect of angle of attack (AoA) on the performance of the supersonic intake for High-Mach Integrated Control Experiment, HIMICO. The results of theoretical analysis using simple shock wave (TFD) suggests that mass capture ratio (MCR) and total pressure recovery (TPR) are increased with positive angle of attack. In fact, comparing the result of 2.5 deg AoA with that of −2.5 deg AoA gained by CFD, MCR and TPR become 21pt and 5pt larger when the nozzle throat height (HNT) is 12.5 mm. The main factors of these phenomena are the increased projected area of the intake entrance and the change of the oblique shock system. On the other hand, TFD incorrectly estimates HNT and TPR in critical state because the pseudo-shock wave is generated instead of the terminal normal shock wave considered in TFD. So the characteristics of the pseudo-shock wave are investigated by CFD. As a result, it was found that the total pressure loss due to the pseudo-shock wave is larger than that of the normal shock wave, and that the total pressure loss becomes smaller with positive AoA due to the decrease in the Mach number at the intake throat. Using the CFD results, we created a model to estimate TPR and HNT in the critical state from the relational equation between the intake throat Mach number and the total pressure loss due to the pseudo-shock wave.

The characterization of pseudo-shock wave in various ducts with constant area

The characterization of pseudo-shock wave in various ducts with constant area

Effects of back pressure fluctuation on pseudo-shock waves in a rectangular duct

Pseudo-shock waves in a scramjet isolator are worth studying. These waves are generally subjected to back pressure fluctuation induced by combustion. The characteristics and unsteadiness mechanism of pseudo-shock waves are investigated using a numerical method. A second nozzle with a movable pintle is utilized to simulate the back pressure fluctuation of a combustion chamber in the numerical model. A 2D numerical method with the kω-shear stress transport turbulence model is adopted to investigate the behavior of pseudo-shock waves. Validation of the numerical model is performed by comparing its results with existing experimental results. The accuracy of the numerical model is proven. The present study aims to investigate the unsteady characteristics of pseudo-shock waves in a scramjet isolator. As indicated by the results, the amplitude of pseudo-shock wave motion can be reduced by increasing the exciting frequency of back pressure fluctuation. The influence mechanism of exciting frequency is the phase difference caused by time delay.

Peculiarities of low-Reynolds-number supersonic flows in long microchannel

The characteristics of low-Reynolds-number supersonic flows in a long microchannel having a rectangular cross section are investigated computationally. The channel is composed of a Laval nozzle and a straight duct. The design Mach number of the nozzle is 2.0 and the Reynolds number calculated at the nozzle exit is 3100. The length of the straight duct is changed from 2 to 18 h, where h is the duct height. In the computations, the Navier–Stokes equations are numerically solved. The computational code is validated using the experimental data measured by the laser-induced fluorescence (LIF) technique. The computational results demonstrate that neither a normal shock wave nor a pseudo-shock wave, which corresponds to the starting shock wave in a supersonic wind tunnel, appears in microchannel flows. Namely, a low-Reynolds-number supersonic flow is created in a channel without the starting shock wave passing along the duct, although it has been believed that a supersonic internal flow should have been formed through the starting shock wave. In addition, it is found that the microchannel flow changes gradually its supersonic state with the channel length under an underexpanded condition, although a starting shock wave for high-Reynolds-number flows suddenly appears in a channel just as its length exceeds a certain specific length. Such unexpected phenomena originate from the peculiarity that the low-Reynolds-number flows can expand (accelerate) along a straight duct at supersonic speeds, although the high-Reynolds-number flows cannot.

Thermal Effects of Liquid/Supercritical Carbon Dioxide Arising From Fluid Expansion in Fracturing

Summary Liquid/supercritical carbon dioxide (L/SC-CO2) fracturing is an emerging technology for shale gas development because it can effectively overcome problems related to clay swelling and water scarcity. Recent applications show that L/SC-CO2 fracturing can induce variations in temperature. Understanding of this phenomenon is rudimentary and needs to be carefully addressed to improve the understanding of CO2 thermodynamic behavior, and thus helps to optimize CO2 fracturing in the field. In this paper, we develop a numerical model to assess the impact of thermal effect on fracture initiation during CO2 fracturing. The model couples fluid flow and heat transfer in the fracture, and is verified by a peer-reviewed solution and observation in laboratory experiments. The velocity, pressure, and temperature are calculated at various time to demonstrate the thermodynamic behavior during fracture initiation. A pseudo shock wave is observed, associated with a compression wave and an expansion wave, which finally leads to an increase in temperature in the new fracture and a decrease in temperature in the initial fracture. The thermal stress is derived to investigate the difference between hydraulic fracturing and CO2 fracturing. The results show that thermal stress, resulting from CO2 fracturing initiation, is comparable to the rock strength, which will help induce microfractures, and thus promote the fracture complexity. The formation pressure after CO2 fracturing is also calculated to evaluate the pressure-buildup potential. This work highlights the importance of CO2 expansion during and after fracturing. It is one of the unique features that differs from hydraulic fracturing. For field-design recommendations, to enhance the thermal effect of CO2 fracturing, it is a good strategy to pump CO2 at high pressure and low temperature into the reservoirs with high Young's modulus, low Poisson's ratio, low permeability, and high geothermal temperature (or large depth). This paper does not address the dynamics of fracture propagation under the influence of thermal effect. Rather, it intends to demonstrate the potential of the thermal effect of CO2 fluid in assisting the fracture propagation, and the importance of incorporating the compressibility of CO2 into fracture modeling and operation design. Failing to account for this thermal effect might underestimate the fracture complexity and stimulated reservoir volume.

CFD study on inlet unstart prediction of dual-mode scramjet

Dual-mode scramjet has become a viewpoint in the research fields of propulsion systems of aeronautics and astronautics. The key operating process of the dual-mode scramjet is from a started situation to an unstarted situation. The operating process composes of three phases. The first phase is when the pseudo shock wave is in the combustor. The second phase is when the pseudo shock wave is in the isolator. The last phase is when inlet unstart happens. In this study, two-dimensional numerical model has been made to replicate the process before inlet unstart. The location of pseudo shock wave has been determined in according to the location of the plug. As the throat area of the plug is decreasing, pseudo shock wave goes upstream. The prediction of inlet unstart has been focused in terms of the throat area ratio. The transition point of throat area ratio just before inlet unstart has been indicated. The transition points of throat area ratio in steady simulation and unsteady simulation are different. This is because the transition information needs time to spread from the plug to upstream location. The difference could be used for predicting the inlet unstart.

Investigation of corner separation and suction control in constant area duct

To control shock-wave/boundary-layer interaction with suction, numerical simulations in a constant-area rectangular isolator with asymmetry nozzle are conducted to understand the complex flow phenomena. Results show that three-dimensional effect significantly weakens isolator resistance to backpressure, which indirectly emphasizes enormous role of viscous effect and geometry to the dynamic behavior of shock train or pseudo-shock wave. The large-scale separated flow only locates at the corner of the upper wall. The mechanism analysis indicates that the flow is accelerated more rapidly along the bottom wall, thereby providing more momentum at the nozzle exit. The displacement and momentum thicknesses and their increasing rate on the upper wall are all greater than that of the bottom wall, which means that upper corner flow is subjected easily to separation. The vortex structure of different cross sections in the pseudo-shock region is formed in different ways and the large-scale vortex still exists after the airflow mixing, which contributes to the deceleration and pressurization. Through the control effect analysis of suction slot, the suction slot in spanwise direction on the side wall can shorten effectively the length of shock train, which indicates that removing the low-momentum fluid near the corners can control significantly the shock-wave/boundary-layer interactions.

Optimization study on the isolator length of dual-mode scramjet

The dual-mode scramjet is a prime candidate for air-breathing propulsion engines given its broad speed range and high efficiency. The detailed structure of the pseudo-shock wave in the isolator is worth investigating. Therefore, the present study investigates the effect of isolator length on the pseudo-shock wave. We replicate the experimental work of the reference paper with a two-dimensional numerical model. Pressure distribution on the body surface is applied in the numerical validation. The numerical result is in good agreement with that of the existing experimental work. We simulate the thermal choking condition with a throttling device. As the throat area of throttling device decreases, the back pressure continues increasing and the pseudo-shock wave moves upstream. Avoiding unstart phenomenon is a key objective in the operation of dual-mode scramjet. Therefore, we investigate the critical condition shortly before unstart phenomenon occurs. The pseudo-shock wave structure is described by pressure distribution. Our main objective is to explore the effect of isolator length on the dual-mode scramjet. The optimal isolator length depends on the maximum back pressure and total pressure loss in the critical condition shortly before the unstart phenomenon occurs. The maximum back pressure in the critical condition is higher when the isolator length ranges from 8.7 to 20.7. The total pressure loss in the critical condition decreases at the beginning and then increases when the isolator length ranges from 8.7 to 20.7.

Mixing augmentation mechanism induced by the pseudo shock wave in transverse gaseous injection flow fields

The mixing process between the injectant and air is very crucial for the engineering implementation of the scramjet engine, and this is due to the very short resident time of fuel in supersonic flows. In the current study, the three-dimensional Reynolds-average Navier–Stokes (RANS) equations coupled with the two equation k-ω shear stress transport (SST) turbulence model have been employed to investigate the transverse injection flow field with the pseudo shock wave induced by the high back pressure, and the freestream Mach number is 3.75. At the same time, the influence of the back pressure on the flow field properties has been evaluated as well. The obtained results show that the pseudo shock wave induced by the back pressure plays an important role in the mixing enhancement between the injectant and air. When the back pressure ratio is larger than 5.0, the mixing efficiency increases with the increase of the back pressure ratio. However, when the back pressure ratio is 3.0, the near-field mixing process has been improved, and accordingly its mixing efficiency in this region is larger than the benchmark. This implies that the intense combustion downstream of the injector can enhance the mixing process between the injectant and air, and the mixing and combustion process can be enhanced mutually. When the pseudo shock wave has been pushed upstream of the wall orifice, more injectant has been brought into the separation zone upstream of the injector, and this is beneficial for the mixing process between the injectant and air.

LES를 이용한 Pseudo-Shock Waves의 가시화

The interaction between a normal shock wave and a boundary layer along a wall surface in internal compressible flows causes a very complicated flow. This interaction region containing shock train and mixing region is called as pseudo-shock waves. Pseudo-shock waves in the divergent part of a rectangular nozzle have been investigated by using large-eddy simulation (LES). LES studies have been done for the complex flow phenomena of three-dimensional pseudo-shock waves. The LES results have been validated against experimental wall-pressure measurements. The LES results are in good agreement with experimental results. Pseudo-shock length and corner separation have been studied in three-dimensional LES model. Comparison of centerline pressure measurement and 3D visualization measurement has been discussed for the corner separation position. It has been concluded that the pseudo-shock length should be measured by using 3D visualization measurement.

Large Eddy Simulation of Pseudo-Shock Waves Using Wall Model

A supersonic turbulent flowfield involving the pseudo-shock waves in an isolator of a supersonic combustion ramjet is computed using two different LES codes which are a high-order upwind finite volume scheme, and a sixth order compact differencing scheme utilizing the localized artificial diffusivity method for stabilizing shock waves and employing a wall model to enable the use of coarse mesh. In the validation study where a supersonic turbulent boundary layer flow over a flat plate is examined, both LES codes are well validated using velocity profile in the boundary layer given by the hot-wire anemometry and normal stress given by the laser Doppler anemometry. In particular, the sixth order compact differencing scheme gives closer agreements with these experimental data. Then, the validated LES codes are applied to solve the Mach number 2.5 supersonic turbulent flowfield involving the pseudo-shock waves. It is shown that typical features of unsteady flowfield of the pseudo-shock waves are well obtained by both schemes. Again, it is indicated that the sixth order compact differencing scheme gives closer agreements with the existing velocity data obtained by particle image velocimetry and pressure fluctuation data on the wall surface. Besides, the computational cost of the compact differencing scheme is found to be 1/7 of that for the upwind finite volume scheme, even though a wall model is solved at each grid point on the wall surface. Therefore, the obtained results in the present study allow recommending the sixth order compact differencing scheme with a wall model for simulating supersonic turbulent flowfield in an isolator involving the pseudo-shock waves.

Mean and Turbulent Velocity Field for Jet Injected into Supersonic Flow with Pseudo-Shock Wave

The mean and turbulent velocity field of a transverse jet in a supersonic flow with a pseudo shock wave by stereoscopic particle image velocimetry was measured. Both the mainstream and the injectant were unheated air. The pseudo shock wave was generated and controlled by means of a flow plug driven by a stepping motor. To study the influence of the generation anddevelopment of the pseudo shockwave,measurements are taken for four different cases of the pseudo shock wave location, including one without a pseudo shock wave. Sequences of up to 750 image pairs for each measurement plane were acquired. From these data, the mean and turbulent characteristics of the flowfield were examined, such as mean velocity and vorticity, turbulence intensities, Reynolds stresses, and turbulence kinetic energy.Themeanvelocity andvorticity indicate that the pseudo shockwave gradually lowered the flow speed of the mainstream and increased the size of the counter-rotating vortex pair. The pseudo shock wave strongly disturbed the flowfield, inducing complicated three-dimensional distortions. The pseudo shock wave disturbed theflowfields near thewalls, and the turbulence kinetic energy in this areawas kept at a high level over long streamwise distances. The streamwise fluctuation intensity was about 4 times stronger than that of the vertical and lateral components, and the streamwise velocity components fluctuated even more greatly than the cross-sectional velocity components. The transverse injection moderately affected not only turbulence around the jet, but also the turbulent characteristics caused by the pseudo shock wave. The injection not only produced turbulence around the jet, but also suppressed the velocity fluctuation caused by the pseudo shock wave.

Mean and Turbulent Velocity Field for Jet Injected into Supersonic Flow with Pseudo-Shock Wave

Mean and Turbulent Velocity Field for Jet Injected into Supersonic Flow with Pseudo-Shock Wave

Flowfield characteristics of a transverse jet into supersonic flow with pseudo-shock wave

We performed an experimental investigation of the flowfield of a transverse jet into supersonic flow with a pseudo-shock wave (PSW). In this study, we injected compressed air as the injectant, simulating hydrocarbon fuel. A back pressure control valve generated PSW into Mach 2.5 supersonic flow and controlled its position. The positions of PSW were set at nondimensional distance from the injector by the duct height (x/H) of −1.0, −2.5, and −4.0. Particle image velocimetry (PIV) gave us the velocity of the flowfield. Mie scattering of oil mist only with the jet was used to measure the spread of the injectant. Furthermore, gas sampling measurements at the exit of the test section were carried out to determine the injectant mole fraction distributions. Gas sampling data qualitatively matched the intensity of Mie scattering. PIV measurements indicated that far-upstream PSW decelerated the flow speed of the main stream and developed the boundary layer on the wall of the test section. The flow speed deceleration at the corner of the test section was remarkable. The PSW produced nonuniformity in the main stream and reduced the momentum flux of the main stream in front of the injector. The blowing ratio, defined as the square root of the momentum flux ratio, of the jet and the main stream considering the effect of the boundary layer thickness was shown to be a useful parameter to explain the jet behavior.

Effect of Local Flow Parameters on Skin-Friction Coefficient in Hypersonic Airbreathing Engines

Effects of Rex, Re , Tw=Taw, and Me on the local skin-friction coefficient (Cf ) in a hypersonic airbreathing combustor with various combustion modes were examined. Measurements were performed in the hypersonic airbreathing combustor with airflowMach number and unit Reynolds numbers at the entrance of combustor of 2.5 and 1:6 10 3:5 10 m , respectively. The ranges of total temperature, total pressure of airflow and fuel equivalence ratio were varied, i.e., 1000–2000 K, 1.0–1.8 MPa, and 0–0.4, respectively, to change Rex, Re , Tw=Taw, and Me. Direct skin-friction measurements were conducted. The state of flow was also surveyed in the streamwise locationwhere skin frictionwasmeasured to attain core-flowparameters, velocity profile, andmomentum thickness. When boundary-layer transition occurred due to combustion, evaluation ofCf by the vanDriestmethod using either by Rex or by Re was not effective. When there was separation with a long oblique shock train and a pseudo shock wave system in the precombustion region due to combustion, evaluation of Cf by the method using Rex was not effective because the origin of the turbulent boundary layer was changed in the downstream combustor.

Effects of Supersonic Nozzle Geometry on Characteristics of Shock Wave Structure

Interaction between the normal shock wave and the turbulent boundary layer in a supersonic nozzle becomes complex with an increase of a Mach number just before the shock wave. When the shock wave is strong enough to separate the boundary layer, the shock wave is bifurcated, and the 2nd and 3rd shock waves are formed downstream of the shock wave. The effect of a series of shock waves thus formed, called shock train, is considered to be similar to the effect of one normal shock wave, and the shock train is called pseudo-shock wave. There are many researches on the configuration of the shock wave. However, so far, very few researches have been done on the asymmetric characteristics of the leading shock wave in supersonic nozzles. In the present study, the effect of nozzle geometry on asymmetric shock wave in supersonic nozzles has been investigated experimentally.

G050062 コールドスプレー用ノズル内の超音速流れに関する実験的研究

Cold spray is a technique to make a coating on a wide variety of mechanical or electric parts by spraying solid particles accelerated by a high-speed gas flow through a converging-diverging nozzle. Experimentally it is well known that the particles start to deposit on the substrate when the impact velocity exceeds material dependent threshold value. Therefore, the particle velocity at impact is very important in the cold spray. In order to investigate the gas flows in the nozzle, we measured the wall-static pressure of the part of diverging nozzle by using pressure sensors attached to the nozzle. Also, quasi-one-dimensional analysis of the gas flow in the cold spray nozzle was conducted for comparison. As a result, the wall-static pressure from nozzle throat to downstream flow decreases smoothly at all the stagnation pressures and increases from the front position of pseudo-shock wave. In addition, from the comparison of experimental and analytical results, it is clarified that the gas flow in the cold spray nozzle is turbulent and can be regarded as a quasi-one-dimensional flow.

Effects of Supersonic Nozzle Geometry on Characteristics of Shock Wave Structure

Interaction between the normal shock wave and the turbulent boundary layer in a supersonic nozzle becomes complex with an increase of a Mach number just before the shock wave. When the shock wave is strong enough to separate the boundary layer, the shock wave is bifurcated, and the 2nd and 3rd shock waves are formed downstream of the shock wave. The effect of a series of shock waves thus formed, called shock train, is considered to be similar to the effect of one normal shock wave, and the shock train is called pseudo-shock wave. There are many researches on the configuration of the shock wave. However, so far, very few researches have been done on the asymmetric characteristics of the leading shock wave in supersonic nozzles. In the present study, the effect of nozzle geometry on asymmetric shock wave in supersonic nozzles has been investigated experimentally.