Flow Characteristics and Thrust Augmentation Effects of Concentric Canister Gas Jets

The productivity of most gas condensate wells is reduced significantly by condensate banking when the bottom hole pressure falls below the dew point. The most important parameter for determining condensate well productivity is the effective gas permeability in the near well region, where very high velocities can occur. An understanding of the characteristics of high-velocity gas-condensate flow is necessary for accurate forecasts of well productivity. A number of laboratory experiments have demonstrated that gas condensate relative permeabilities increase at high velocity, reducing the negative impact of condensate banking on well productivity. On the other hand, inertial (non-Darcy) flow effects can reduce the effective gas permeability and lead to lower productivity. This paper presents results of relative permeability measurements on a low permeability sandstone core, using a 5-component gas-condensate fluid. The experiments used a pseudo-steady-state technique at high pressure and high velocity, measuring relative permeability under conditions similar to the near-well region of a gas-condensate reservoir. By carrying out measurements at a range of interfacial tensions and velocities, the results can be used to distinguish between high capillary number and inertial flow effects, and to quantify the impact of these two conflicting phenomena. The experiments suggest that the inertial flow coefficient in a 3-phase gas-condensate-water system is about 50% higher then in the equivalent 2-phase gas-water system. The results of these experiments have been modelled through a correlation of relative permeability versus capillary number, together with an inertial flow correction to the gas permeability. The paper discusses these models and demonstrates how they can be used to calculate gas-condensate well performance in full-field reservoir simulation.

Three-dimensional shock/boundary layer interactions (SBLIs) in the hypersonic inlet generate the separation vortex, which affects the flow uniformity of the inlet and can even cause the inlet to unstart. This study experimentally investigates the separation vortex produced by a crossing SBLI in a supersonic quiet wind tunnel. Using a nanoparticle-based planar laser scattering method, the tomography-like three-dimensional structures of the separation vortex on the transverse, streamwise, and horizontal planes are demonstrated. The semi-elliptical separation vortex is formed from a pair of anti-rotating vortices and exhibits eddies around the vortex core; it remains stable in the absence of the expansion effect. Additionally, fractal dimension analysis reveals that the separation vortex core experiences stable streamwise development, while its outer edge is fragmented and dissipated due to the shearing effect of the mainstream. This investigation provides valuable insight for potential flow control to mitigate separation issues in hypersonic inlets.

Shock/boundary layer interaction (SBLI) is an undesirable phenomenon, occurring in high-speed propulsion systems. The conventional method to manipulate and control SBLI is using a bleed system that involves the removal of a certain amount of mass of the inlet flow to control boundary layer separation. However, the system requires a larger nacelle to compensate the mass loss, larger nacelles contribute to additional weight and drag and reduce the overall performance. This study investigates a novel type of flow control device called micro-ramps, a part of the micro vortex generators (VGs) family that intends to replace the bleed technique. Micro-ramps produce pairs of counter-rotating streamwise vortices, which help to suppress SBLI and reduce the chances of flow separation. Experiments were done at Mach 5 with two micro-ramp models of different sizes. Schlieren photography, surface flow visualization and infrared thermography were used in this investigation. The results revealed the detailed flow characteristics of the micro-ramp, such as the primary and secondary vortices. This helps us to understand the overall flow physics of micro-ramps in hypersonic flow and their application for SBLI control.

Passive methods of controlling shock/boundary-layer interactions (SBLIs) consist of a porous surface covering a cavity or a plenum located in the region of the SBLI. The present study focuses on the flowfield downstream of a Mach 1.42 SBLI controlled with various passive devices such as a conventional porous plate, a microporous plate, streamwise slots, a conventional mesoflap array, and a hybrid flap array. Qualitative analysis of the flowfield for the various control devices investigated was achieved with spark shadowgraph visualizations and surface oil-flow visualizations. Quantitative analysis was accomplished by measuring surface static pressure distributions and boundary layer velocity profiles. The flowfields downstream of the slot-controlled and hybrid flap array-controlled SBLIs were found to be highly three-dimensional, whereas the flowfields were predominantly two-dimensional for the remainder of the control devices. It was found that only the conventional mesoflap array had an improved total pressure recovery compared to the baseline solid wall.

Accurate simulation of compressible turbulent flows and shock–vortex interactions remains a core challenge in computational fluid dynamics, especially when resolving fine-scale vortical structures alongside strong discontinuities. This paper introduces a hybrid weighted essentially non-oscillatory (WENO) scheme aimed at balancing the demands of shock-capturing and turbulence resolution in compressible flows. The method delivers improved accuracy in capturing both classical flow discontinuities and complex vortical structures typical of turbulent flows. The proposed scheme is tested across core case studies, including one-dimensional shock–entropy wave interactions, two-dimensional double-vortex pairing, and three-dimensional Taylor–Green vortex transition to turbulence. Results show that this hybrid scheme provides sharper resolution of discontinuities and better captures fine-scale turbulent structures. In the double-vortex pairing case, the method reduces numerical dissipation by nearly 20%, compared to earlier versions of WENO schemes, enabling a more precise depiction of vortex dynamics and mixing. For the Taylor–Green vortex, the scheme detects more turbulent structures than the 11th-order method, improving predictions of kinetic energy dissipation and enstrophy evolution. These advancements are vital for applications in science and engineering involving compressible turbulence and shock–boundary layer interactions, where accurately resolving both discontinuities and vortical features is essential.

Laminar hypersonic flows at Mach 7.10 with unit Reynolds numbers of 5.2×104, 1.04×105, and 4.14×105 m−1 over a 30°/55° double-wedge configuration were studied to investigate the spatial–temporal characteristics of the flow in a time-accurate manner. Close comparisons were made between previous kinetic and current continuum approaches to test the validity of the continuum assumption, especially considering the presence of large gradients associated with shock–shock and shock–boundary layer interactions, as well as spanwise instabilities. Previous results from direct simulation Monte Carlo, which inherently predicts rarefied effects such as velocity slip and temperature jumps, were found to be in very close agreement with the current work, even for the lowest Reynolds number. The impact of velocity slip and temperature jumps on flow and surface parameters was investigated, and comparisons were made with a no-slip and constant temperature wall model. The temporal and spatial variation of two- and three-dimensional flows were thoroughly investigated using two-dimensional (2D), three-dimensional (3D) periodic sidewall boundary conditions, and a full 3D configuration consistent with existing experimental data. Close comparisons among the 2D and 3D cases were made. The characteristics of 2D periodic oscillations were reported for the moderate Reynolds number case for the first time. The presence of spanwise instabilities, even at a relatively low free stream pressure of about 100 Pa, establishes that the flow field depends on spanwise effects and is fully 3D. High-fidelity numerical schlieren videos captured strong spanwise oscillations for 3D configurations.

Shock/boundary layer interaction (SBLI) is an undesirable phenomenon, occurring in high-speed propulsion systems. The conventional method to manipulate and control SBLI is using a bleed system that involves the removal of a certain amount of mass of the inlet flow to control boundary layer separation. However, the system requires a larger nacelle to compensate the mass loss, larger nacelles contribute to additional weight and drag and reduce the overall performance. This study investigates a novel type of flow control device called micro-ramps, a part of the micro vortex generators (VGs) family that intends to replace the bleed technique. Micro-ramps produce pairs of counter-rotating streamwise vortices, which help to suppress SBLI and reduce the chances of flow separation. Experiments were done at Mach 5 with two micro-ramp models of different sizes. Schlieren photography, surface flow visualization and infrared thermography were used in this investigation. The results revealed the detailed flow characteristics of the micro-ramp, such as the primary and secondary vortices. This helps us to understand the overall flow physics of micro-ramps in hypersonic flow and their application for SBLI control.

Slot is one of the features that control Shock wave-boundary layer interaction (SBLI), which is generally used to prevent strong interference from shockwaves to the boundary layer in supersonic flows. With this feature, the height of the triple point of λ shock significantly increases, and this increase causes a decline in shock power and pressure drop rate. In the current paper, the main focus is on the monitoring of the geometrical effect of slot as an influential parameter on the structure of the shock and flow characteristics by using numerical methods. Therefore, the averaged implicit Navier-Stokes equations and two equation standard k-ω turbulence models for the numerical simulation of the flow field have been used. Results indicate that the numerical results are fairly consistent with the experimental data. Because of the increase in the number of slots (n), and the leading leg of the λ shock is located within the slot, the height of the triple point increases. However, because of the increasing drops due to viscosity, the total pressure changes are negligible. In addition, with an increase in this parameter, changes in the static pressure caused by the leading leg of the shock have increased. By increasing the width of the slots, the height of the triple point has had an upward trend up to s = 8 mm and then had nearly constant values. In this mode, the static pressure changes resulting from the leading leg of the shock are negligible. For increasing the number or the width of slots, the re-expansion waves formed within the slot are removed because of the reduction in the severity of the changes in the boundary layer. To simulate and compare the results with the data obtained from the experimental tests, results from the Cambridge University's wind tunnel tests have been used.

In this paper, a small supersonic wind tunnel was designed and built to study the flow characteristics of a supersonic impulse turbine cascade by experiment. The flow was visualized by means of a single pass Schlieren system. The supersonic cascade with 3-dimensional supersonic nozzle was tested over a wide range of pressure ratio. Highly complicated flow patterns including shocks. nozzle-cascade interaction and shock boundary layer interactions were observed.

A high-order in space central difference scheme with flux-corrected transport has been employed for numerical simulation of unsteady three-dimensional compressible fluid flow about the supercritical airfoil NLR7301 in the transonic regime with shock-boundary layer interaction. The high order in space enables to reproduce well vortical structures and to predict transition to turbulent flow as well as to perform DNS. Numerical results of the method are compared with experimental mean flow data.KeywordsDirect numerical simulationhigh-order difference methodtransonic flowtransition to turbulent flowsupercritical airfoil

PurposeThis study aims to investigate the prevalence of shock boundary layer interaction (SBLI) in air-breathing intake system is highly undesirable since this leads to high pressure gradients, typical stream mutilation and pressure drop. A novel flow control mechanism is incorporated in this research holding an array configuration of passive flow control device (micro ramps [MR]) that is adapted to improve the boundary layer stability.Design/methodology/approachTwo geometric variants of the MR, namely, MR40 and MR80 is considered which reduce the pressure drop during SBLI. The incidence oblique shock wave angle of 34° is considered for the modelling. Large eddy simulation (LES) turbulence model was used with subgrid models of Wall modelled LES, Smagorinsky–Lilly to compute the unsteady effects of SBLI control using micro vortex generators. The unsteady results are compared with steady Reynold’s average Naviers–Stoke’s equation for calibrating the turbulence models.FindingsThe array configuration of MR80 reduces the pressure drop by 22% as compared with no ramp configuration and also reduces the flow distortion in hypersonic inlet. The most affected region of the MR is in the vicinity of center-line. Quantitative results prove that the upstream influence of the shock waves has been largely reduces by MR80 array configuration as compared to single MR80 pattern configuration. Different vortex structures found in the experiments was exclusively predicted using LES.Originality/valueThis paper substantiates the requirement of MR array configuration for transferring the momentum from free stream to the boundary layer and thereby energizing the boundary layer. This process of energization delays the flow separation in hypersonic flow.

Shock-boundary layer interaction (SBLI) is frequently met in supersonic engine inlet flow and external flow. A detailed study on the mechanism of reduction of shock induced flow separation by micro-vortex generator (MVG) is carried out by high order implicit large eddy simulation (ILES). To generate the fully developed turbulent inflow, a series (20,000) of turbulent profiles are given by our previous DNS results. The mechanism of reduction of shock induced flow separation by MVG was originally considered as a result of streamwise vortex mixing. It was thought that the mixing would result in a plump turbulent velocity profile which has stronger separation resistance. It was claimed that turbulent flow has so strong mixing that the velocity profile becomes plump, which lead to reduction of shock induced flow separation. However, according to our LES study reported in this paper, the real mechanism of flow separation reduction is that the shock wave breaks down and disappears when the ring-like vortices generated by MVG are passing through the shock. On the other hand, the vortex structures never break down and is influenced very little when they pass the shock wave. Therefore, the shock induced flow separation is reduced by moving ring-like vortices which were generated by MVG through the manner to break down the shock wave, but mainly not by the streamwise vortex mixing which produces a plump velocity profile in the boundary layer. Details of the investigation on the mechanism are reported in this paper.

Pump as turbine (PAT) is an efficient, simple, and cost-effective equipment combining pump and turbine and is one of the excellent energy recovery devices. It is helpful to master the flow characteristics of the key component impeller for the further optimization and design of the PAT. To analyze the unsteady flow features in the impeller of a double-suction pump operating as a turbine, numerical simulations were conducted using the shear stress transport (SST) k-ω turbulence model at the designed operating conditions. By utilizing proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) methods on the unsteady velocity field of a single cycle, the dominant modes up to the fourth order, along with their respective space–time information, can be extracted. The velocity field and vorticity field analysis were performed on the first four modes extracted using two different methods. Additionally, the vortex structures were extracted using the Ω method. The analysis demonstrates that the POD and DMD methods effectively decompose the intricate flow characteristics within the impeller into dynamic–static interference modes, fundamental modes, and dissipative modes. The dynamic–static interference mode is dominant, reflecting the flow characteristics influenced by the stationary components within the impeller. The vortex structure is mainly small tubular vortex and point vortex. The fundamental mode captures the steady flow field characteristics caused by the blade channel geometry. The vortex structure is mainly continuous tubular vortex and the diameter becomes larger. The dissipative mode reflects the flow separation generated on the blades by disturbances from the stationary components. The vortex structure is dominated by point vortex and discontinuous tubular vortex. Comparing the outcomes of the two modal analysis methods shows that the POD method has a distinct advantage in showcasing key changing nodes. In contrast, the DMD method is superior in isolating modes with a single frequency and in determining their stability.

NOWLEDGE of the skin-friction drag is of primary importance in the design of advanced aircraft and/or aerospace vehicles. Considerable analytical and experimental studies have been conducted to predict the skin friction for flow conditions that are either laminar or fully turbulent. For mixed flow conditions, the total skin-friction drag coefficient will vary widely depending on the location of the transition point because of the large differences between the drag coefficients for laminar and turbulent flow at identical Reynolds numbers. When transition from laminar to turbulent flow occurs at low speed, the well-known formula of Prandtl-Schli chting1 has provided, although empirical in nature, a quick but accurate estimate of the skin-friction drag. Bertram2 has proposed a semi-empirical method for estimating the skinfriction drag on a delta wing at hypersonic speeds wherein transition is present. However, reviewing his results (Ref. 2, Fig. 15), they do not appear correct. The purpose of this note is to present a method that is similar to Bertram's but which is more rigorous and less complicated and provides numerical results of the skin-friction drag coefficient as a function of the transition on a delta wing. Consider a flat plate delta wing in a hypersonic flow at small values of angle of attack. Depending on the freestream Mach number, unit Reynolds number, leadingedge thickness, surface roughness, etc., a portion of the flow on the wing will become turbulent. For a delta wing at small angles of attack, chordwise strip theory is generally valid provided the spanwise pressure gradient induced by the leading-edge-bluntness or boundary-laye r-shock interaction is small. It may be noted that ARA has successfully obtained and correlated pressure and force data from a large (48 in. long) 70° swept blunted-leading-edge delta wing at Mach numbers 3 to 8; the data have shown that, except near the nose, chordwise strip theory is a good approximation. Not only is the theoretical treatment of the transition location difficult, but, in addition, its experimental determination is also quite illusive. It is generally accepted that the beginning and end points are determined, respectively, by the minimum and maximum locations of the axial variations of either the surface-pitot-pressure or the heat-transfer rate. There are several thoughts, however, for defining the actual transition location. For instance, some investigators define the location as the point of inflection between the minimum and maximum axial surface-pitot-pressure or heat-transfer rate; others use the end point of the transition region. At moderate supersonic speeds the streamwise distance between the minimum and maximum pressure or heat-transfer rate is

To investigate the spatiotemporal characteristics of vortex structures within a centrifugal pump and explore the application of modal decomposition techniques in three-dimensional vortex feature extraction, this study employs computational fluid dynamics (CFD) to numerically simulate the unsteady flow in a centrifugal pump under flow rates of 80 m 3 /h (low-flow condition), 100 m 3 /h (design flow condition), and 120 m 3 /h (high-flow condition). By integrating the Omega-Liutex method with modal decomposition techniques, including Proper Orthogonal Decomposition (POD) and Dynamic Mode Decomposition (DMD), a detailed characterization of the vortex structures is performed. The results indicate that, compared to other vortex identification criteria, the Omega-Liutex vortex identification technique effectively captures the flow structures within the centrifugal pump. By applying the POD method to analyze the vortex characteristics in the impeller region under three different flow conditions, the analysis revealed that tip vortices and wake vortices dominate the energy distribution across all flow conditions, with an energy proportion ranging from 34.7% to 40.3%. Additionally, under both low-flow and design-flow conditions, the first and second-order modal structures of the vortex in the impeller region demonstrate periodic behavior. The dominant vortex structures in the volute region exhibit significant variations under different flow conditions. Under low-flow and design-flow conditions, the primary vortex structures in the volute region include the tongue-shedding vortex and the wall-attached vortex, with a nonlinear interaction observed between the first and second-order modes. Under high-flow conditions, the primary energetic vortex structure in the volute region is the recirculating vortex in the diffuser section. A comparison of the vortex areas for all modes in the impeller and volute regions under three different flow conditions reveals that the primary mode exhibits the smallest vortex area. The first-order mode obtained through the DMD method clearly reveals the overall flow characteristics within the centrifugal pump. Among the first four modes, there are always two modes exhibiting identical vortex structures. In the comparison of vortex feature extraction, the POD method can identify large-scale vortices with relatively low energy in the impeller and volute regions, whereas the DMD method extracts vortex structures more clearly and comprehensively.