Stagnation Pressure Research Articles

Bimodal velocity and size distributions of pulsed superfluid helium droplet beams.

We report detailed measurements of velocities and sizes of superfluid helium droplets produced from an Even-Lavie pulse valve at stagnation pressures of 20-60 atm and temperatures between 5.7 and 18.0 K. By doping neutral droplets with Rhodamine 6G cations produced from an electrospray ionization source and detecting the positively charged droplets at two different locations along the beam path, we determine the velocities of the different groups of droplets. By subjecting the doped droplet beam to a retardation field, size distributions can then be analyzed. We discover that at stagnation temperatures above 8.0 K, a single group of droplets is observed at both locations, but at 8.0 K and below, two different groups of droplets with different velocities are detectable. The slower group, considered from fragmentation of liquid helium, cannot be deterred by the retardation voltage at 9 kV, implying an exceedingly large size. The faster group, considered from condensation of gaseous helium, has a bimodal distribution when the stagnation temperatures are below 12.3 K at 20 and 40 atm, or 16.1 K at 60 atm. We also report similar size measurements using low energy electrons for impact ionization, and this latter method can be used for facile in situ characterization of pulsed droplet beams. The mechanism of the bimodal size distribution of the condensation group and the reason for the coexistence of both the condensation and fragmentation groups remain elusive.

Simulation study on near-field structure and flow characteristics of high-pressure CO2 released from the pipeline

The accidental release of high-pressure carbon dioxide (CO2) can cause serious damages to both humans and pipeline equipment. Therefore, it is of great significance to have a deeper understanding about the release characteristics of high-pressure CO2 for improving the safety level of Carbon Capture and Storage (CCS) technologies. Both industrial-scale and laboratory-scale studies have been carried out to predict the release behaviors. In recent years, computational fluid dynamics (CFD) simulation has become a crucial method to study the instantaneous changes and microscopic details of the fluid behaviors. In this paper, the simulation method was employed to study the near-field structure and flow characteristics of high-pressure CO2 released from pipelines. The Peng-Robinson Equation of State (EOS) was used to compute the thermodynamic properties of high-pressure CO2, and SST k-ω model was applied to simulate the structure and physical parameters of the under-expanded jet. In addition, the multi-phase mixture model was introduced to study the phase transition. The non-equilibrium liquid/vapor transition is modeled by introducing ‘source terms’ for mass transfer and latent heat. Compared to the experimental results, the simulation results showed good agreement. Furthermore, the influences of operating conditions, including different stagnation pressure, stagnation temperature, and nozzle size, were analyzed.

Visualization of separation and reattachment in an incident shock-induced interaction

An experimental study is conducted on the qualitative visualization of the flow field in separation and reattachment flows induced by an incident shock interaction by several techniques including shear-sensitive liquid crystal coating (SSLCC), oil flow, schlieren, and numerical simulation. The incident shock wave is generated by a wedge in a Mach 2.7 duct flow, where the strength of the interaction is varied from weak to moderate by changing the angle of attack α of the wedge from 8° and 10° to 12°. The stagnation pressure upstream was set to approximately 607.9 kPa. The SSLCC technique was used to visualize the surface flow characteristics and analyze the surface shear stress fields induced by the initial incident shock wave over the bottom wall and sidewall experimentally which resolution is 3500 × 200 pixels, and the numerical simulation was also performed as the supplement for a clearer understanding to the flow field. As a result, surface shear stress over the bottom wall was visualized qualitatively by SSLCC images, and flow features such as separation/reattachment and the variations of position/size of separation bubble with wedge angle were successfully distinguished. Furthermore, analysis of shear stress trend over the bottom wall by a hue value curve indicated that the relative magnitude of shear stress increased significantly downstream of the separation bubble compared with that upstream. The variation trend of shear stress was consistent with the numerical simulation results, and the error of separation position was less than 2 mm. Finally, the three-dimensional schematic of incident shock-induced interaction has been achieved by qualitative summary by multiple techniques, including SSLCC, oil flow, schlieren, and numerical simulation.

Study of Wall Static Pressure Distribution on Flat Surface by Impinging Submerged Jet from Non-circular Orifice

The distribution of wall static and stagnation (CP and CPO) pressure coefficient on a flat rectangular element by impinging air jet from the hexagonal orifice is obtained from experimentation. The past research studies helped to identify key parameters such as orifice geometry, jet exit-to-plate-distance (Z/dj), test section inclination (θ), jet Reynold number (Re), lateral distance-to-jet diameter (X/dj), test surface type and geometry, for better and acceptable results. The experimental outcome helps to know the effect of identified key parameters on wall static and stagnation pressure on a rectangular test plate in a confined flow path. The independent nature of wall static pressure is observed for all jet Reynold number (10000 ≤ Re ≤ 50000). Higher pressure coefficient values were observed at lower Z/dj = 1, X/dj = 0 and θ = 0. A significant drop in CP values are seen with the increase in Z/dj, X/dj and θ. The experimental CP and CPO contribution of confined flow are compared against the unconfined flow, around 48% to 58% enhancement is observed when confinement is used. Experimental pressure drop measurements across orifice were made and pressure loss coefficient (PC) for hexagonal orifice of confined and unconfined condition are reported.

Analysis of the Simulation Conditions of the Aerodynamic Heating in Subsonic High-Enthalpy Air Jets from the VGU-4 HF Plasmatron

The domain in the “total enthalpy–stagnation pressure” coordinates and, correspondingly, the ranges of the velocity and altitude of the atmospheric entry of a body with the nose bluntness radius of 1 m, for which the conditions of the local modeling of heat transfer to the stagnation point are fulfilled in subsonic high-enthalpy air jets flowing around cylindrical flat-nosed models, 20 to 140 mm in diameter, are established. The region of the trajectory of the European experimental IXV vehicle entry into the atmosphere, in which the local modeling of the convective heating of the vicinity of its nose, assuming 1 m in radius, is possible in the VGU-4 HF plasmatron, is determined.

Modes of plasma-stabilized combustion in cavity-based M = 2 configuration

This paper considers the modes of plasma-stabilized supersonic combustion in a cavity-based flameholding configuration. The testing was performed in the SBR-50 supersonic facility at the University of Notre Dame under the following conditions: Mach number M = 2, stagnation pressure and temperature P0 = 1–3.2 bar, T0 = 295–750 K, and an ethylene injection rate of ṁC2H4 = 0–8 g/s. A rail type electric discharge is employed in this work, which represents a modification of the previously explored quasi-DC (Q-DC) configuration. Two principally different cases are realized and described in detail: in- or over-the-cavity combustion, and bulk combustion characterized by a significant increase of pressure downstream of the fuel injection ports. Rail discharge performance is compared to the performance of Plasma-Injection Modules (PIMs) using a 1D analysis. Dimensionless analysis of the plasma-assisted combustor performance indicates reasonable agreement of the in-cavity combustion mode with the Ozawa stability diagram, while the plasma-assisted bulk combustion is essentially distinct from this formalism.

Numerical Investigation of the Centrifugal Compressor Stability Improvement by Half Vaned Low Solidity Diffusers

Centrifugal compressors for the fuel cell vehicles often operate near the surge line compared with the turbocharger compressors. Low solidity and half vaned diffusers are recognized as good ways to improve the stability of the centrifugal compressor. The presented work investigated four diffuser configurations (i.e., the vaneless diffuser (VLD), full-height low solidity vaned diffuser (LSVD), hub-side half vaned diffuser (HVD) and shroud-side half vaned diffuser (SVD)) through steady-state and unsteady numerical simulations. The results show that the best performance is achieved by the LSVD, HVD and SVD at the design, surge and choke conditions. The flow rate at the surge operating point of the HVD has decreased by 15.53% compared with the LSVD, and 9.21% compared with the VLD. At near surge operating point, a longitudinal suction side passage vortex is formed on the hub of the LSVD and rotates as circumferential stall cells. A hairpin vortex is formed along the leading edge and is dragged by the main flow along the suction side as a local vortex shedding. The mechanism of the stability improvement by half vaned diffusers is that the tip leakage vortex migrates from the clearance side to the vane mounting side and replenishes the low-momentum zone on the mounting side. The best position where the half vaned diffuser should be mounted is based on the impeller outlet flow conditions, namely, the location of the wake region, where the meridional velocity and relative stagnation pressure is low.

Impact of Fuel Injection Distribution on Flame Holding in a Cavity-Stabilized Scramjet

This study investigated the impact of fuel injection distribution on flame stabilization and heat release in a cavity-stabilized scramjet. Experiments were conducted using a hydrogen air heater that provided vitiated air at an average stagnation temperature of 1570 K and stagnation pressure of 12.7 atm, accelerated to a Mach number of 2.2 at the combustor entrance. Ethylene was injected using two fuel injection locations, one upstream of the cavity and another in the cavity. Three different fuel injection distributions were studied for global equivalence ratios of 0.2, 0.3, and 0.4. Measurements included high-speed chemiluminescence and shadowgraph, as well as wall-mounted pressure transducers. It was found that when the cavity fuel flow rate was low, the flame stabilized on the cavity shear layer and resulted in a gradual pressure rise in the combustor. A sufficiently high fuel flow rate in the cavity led to flame stabilization upstream of the cavity, in the fuel jet wake, and a more confined combustion zone. Shadowgraph imaging showed that injection of the fuel from the cavity floor at a sufficient momentum penetrated the shear layer over the cavity and enabled flame propagation upstream of the cavity. It was concluded that the heat release distribution and pressure profile in the combustor could be significantly influenced by the fuel injection distribution through its impact on the flame stabilization mode.

Numerical investigation of cavity-induced enhanced supersonic mixing with inclined injection strategies

In this study, an inclined injection system with a cavity and grooves arranged in series is developed to enhance the supersonic mixing in a scramjet combustor. Three-dimensional steady Reynolds-averaged Navier–Stokes equations, the two-equation shear stress transport k-ω model, and the finite-rate/no turbulence-chemistry interaction model are used to simulate the flow field structures with and without grooves at different inclination angles. The numerical method used in the study has been verified using experimental data from the open literature, and the static wall pressure distribution of the numerical simulation is in good agreement with the experimental data. The numerical results reveal that the grooved configuration has a significant influence on the flow field structure. The downstream streamline of the structure with grooves is more turbulent and produces more streamwise vortices because the geometry of the grooves enhances the shear effect between the vortex and the high-speed incoming flow. In the case of the grooved configuration, the mixing coefficient at the fuel orifices is slightly lower, but the fuel mixing effect in the downstream is enhanced and gradually exceeds the grooveless configuration. Moreover, the grooved configuration can obtain the best mixing coefficient at a jet angle of 30°. The grooved configuration has a greater stagnation pressure loss than the grooveless configuration.

Experimental investigation of hypersonic flight-duplicated shock tunnel characteristics

Hypersonic air-breathing propulsion is one of the key techniques for future aviation and the ground aerodynamic testing for full scale test models with sufficient test time at flight conditions is of fundamental importance for verifying hypersonic air-breathing engines. Based on the backward detonation-driven concept, the hypersonic flight-duplicated shock tunnel (or JF-12 shock tunnel) has been successfully constructed and calibrated. This facility is capable of reproducing airflow for Mach numbers ranging from 5 to 9 at an altitude of 25–50 km, with a test duration of more than 100 ms. To quantify the performance of the shock tunnel, experiments were conducted to investigate the aerodynamic characteristics of the test flows and the effects of several critical techniques that play important roles in the operation of the shock tunnel. The stagnation pressure was constant within ±5 $$\%$$ and the average stagnation pressure varied by less than 0.048 $$\%/$$ ms. The variation of the stagnation pressure in repeated experiments is less than 2.0 $$\%$$ , indicating the good repeatability of the wind tunnel. The non-uniformity of the Mach number in the core flow field at the nozzle exit was within ±2.5 $$\%$$ . Additional, a uniform flow field is established upstream of the nozzle exit. The axial gradients of the flow field are small since the Mach number varies less than 1.7 $$\%/$$ m. Findings regarding the ignition technology, diaphragm ruptures, detonation driver capacity, incident shock-wave decay, and tunnel operation mode are also presented. The findings of this study are not only helpful for operating the shock tunnel, but can also assist the future development of hypersonic wind tunnels.

In situ nozzle reservoir thermometry by laser-induced grating spectroscopy in the HELM free-piston reflected shock tunnel

Experimental determination of test gas caloric quantities in high-enthalpy ground testing is impeded by excessive pressure and temperature levels as well as minimum test timescales of short-duration facilities. Yet, accurate knowledge of test gas conditions and stagnation enthalpy prior to nozzle expansion is crucial for a valid comparison of experimental data with numerical results. To contribute to a more accurate quantification of nozzle inlet conditions, an experimental study on non-intrusive in situ measurements of the post-reflected shock wave stagnation temperature in a large-scale free-piston reflected shock tunnel is carried out. A series of 20 single-shot temperature measurements by resonant homodyne laser-induced grating spectroscopy (LIGS) is presented for three low-/medium-enthalpy conditions (1.2–2.1 MJ/kg) at stagnation temperatures 1100–1900 K behind the reflected shock wave. Prior limiting factors resulting from impulse facility recoil and restricted optical access to the high-pressure nozzle reservoir are solved, and advancement of the optical set-up is detailed. Measurements in air agree with theoretical calculations to within 1–15%, by trend reflecting greater temperatures than full thermo-chemical equilibrium and lesser temperatures than predicted by ideal gas shock jump relations. For stagnation pressures in the range 9–22 MPa, limited influence due to finite-rate vibrational excitation is conceivable. LIGS is demonstrated to facilitate in situ measurements of stagnation temperature within full-range ground test facilities by superior robustness under high-pressure conditions and to be a useful complement of established optical diagnostics for hypersonic flows.

Rotating detonation waves in annular gap with variable stagnation pressure

We consider a three-dimensional unsteady flow with one, two, or four rotating detonation waves arising in an annular gap of an axially symmetric device between two parallel planes perpendicular to its symmetry axis. The corresponding problem is formulated and studied. It is assumed that there is a reservoir with quiescent homogeneous propane–air combustible mixture with given stagnation parameters; the mixture flows from the reservoir into the annular gap through its external cylindrical surface toward the symmetry axis, and the parameters of the mixture are determined by the pressure in the reservoir and the static pressure in the gap. The detonation products flow out from the gap into a space bounded on one side by an impermeable wall that is an extension of a side of the gap. Through a hole on the other side of the gap and through a conical output section with a half-opening angle of $$45^{\circ }$$ , the gas flows out from the engine into the external space. We formulate a model of detonation initiation by energy supply in which the direction of rotation of the detonation wave is defined by the position of the energy-release zone of the initiator with respect to the solid wall situated in a plane passing through the symmetry axis. After a while, this solid wall disappears (burns out). We obtain and analyze unsteady shock-wave structures that arise during the formation of a steady rotating detonation. We measure and compare thrust characteristics for different numbers of detonation waves. It was shown that the thrust weakly depends on the number of waves while the specific impulse increases with increasing number of waves. The study of rotating detonation at various values of the stagnation pressure was conducted. The calculation results show that at a stagnation pressure less than the critical value, the realization of rotating detonation is impossible. It is established that, depending on the stagnation pressure, qualitatively different shock-wave structures can be observed. The analysis is carried out within single-stage combustion kinetics by the numerical method based on the Godunov scheme with the use of an original software system developed for multi-parameter calculations and visualization of flows. The calculations were carried out on the Lomonosov supercomputer at Moscow State University.

Direct current electric field regulates endothelial permeability under physiologically relevant fluid forces in a microfluidic vessel bifurcation model.

Previous in vitro studies have reported on the use of direct current electric fields (DC-EFs) to regulate vascular endothelial permeability, which is important for tissue regeneration and wound healing. However, these studies have primarily used static 2D culture models that lack the fluid mechanical forces associated with blood flow experienced by endothelial cells (ECs) in vivo. Hence, the effect of DC-EF on ECs under physiologically relevant fluid forces is yet to be systematically evaluated. Using a 3D microfluidic model of a bifurcating vessel, we report the role of DC-EF on regulating endothelial permeability when co-applied with physiologically relevant fluid forces that arise at the vessel bifurcation. The application of a 70 V m-1 DC-EF simultaneously with 1 μL min-1 low perfusion rate (generating 3.8 dyn cm-2 stagnation pressure at the bifurcation point and 0.3 dyn cm-2 laminar shear stress in the branched vessel) increased the endothelial permeability 7-fold compared to the static control condition (i.e., without flow and DC-EF). When the perfusion rate was increased to 10 μL min-1 (generating 38 dyn cm-2 stagnation pressure at the bifurcation point and 3 dyn cm-2 laminar shear stress in the branched vessel) while maintaining the same electrical stimulation, a 4-fold increase in endothelial permeability compared to the static control was observed. The lower increase in endothelial permeability for the higher fluid forces but the same DC-EF suggests a competing role between fluid forces and the applied DC-EF. Moreover, the observed increase in endothelial permeability due to combined DC-EF and flow was transient and dependent on the Akt signalling pathway. Collectively, these findings provide significant new insights into how the endothelium serves as an electro-mechanical interface for regulating vessel permeability.

Development of a New Supersonic Rotor-vane Ejector using Computational Fluid Dynamics

An innovative design of a supersonic rotor pressure-exchange ejector is introduced in this paper. In this design, momentum is exchanged between supersonic primary flow and secondary flow using an idle rotor. A CFD code developed to model the 3-D compressible, viscous and turbulent flow of air inside the new design of ejector. Roe approach and Spallart-Allmaras methods used to analyze flow inside the ejector. The flow inside the ejector was modeled by using a structured grid and air was employed as the working fluid in both primary and secondary streams. The Mach number of the motive flow was set at 2. Momentum exchanged between the primary and secondary flows because of direct contact between those. In addition to that, rotation of idle rotor and mechanical blades entrained the secondary flow to the ejector. Enthalpy, entrained mass flow rate and created vacuum presented for the flow inside the ejector for different configurations of the rotor and ejector until an optimum case was achieved. Also, uniformity of the flow at discharge section compared between ejectors. For the optimum case with the presented geometry, the ultimate rotor speed of 50000 rpm was obtained and an increase of 47% in entrainment ratio achieved with respect to the stationary blades. To study the flow field in more details, the contours of the Mach number and stagnation pressure were compared according to the different sections of computational domain.

Experimental Study on the Diagnostics of Internal Compressible Flows using Outer Surface Nozzle Temperature

Cold spray (CS) is one of the thermal spraying methods which make coatings by spraying melted minute particles on substrate surfaces. Solid particles are sprayed onto a substrate at high speed to form a coating accelerated by a supersonic gas flow using a convergent-divergent nozzle. Therefore, it is important to understand the gas flow condition inside the nozzle that accelerated the gas flow. However, diagnosing internal gas flow is almost impossible for commercial CS nozzles because the nozzle is fabricated with hard metal alloy to prevent abrasions of the nozzle. Therefore, it is impractical to drill pressure taps to the CS nozzle to diagnose the gas flow. To solve this problem, authors propose non-intrusive flow-diagnostic methods which uses outer surface temperature of the CS nozzle and temperature recovery factor of the working gas. This research shows that the accuracy of this method is improved by measuring the outer surface temperature of the entire axial direction of the nozzle, and that the temperature recovery factor of the internal flow depends on the stagnation pressure. Therefore, the temperature recovery factor of the internal flow should be treated as a function of stagnation pressure.

Flame stabilization enhancement by microjet-based virtual shock wave generators in a supersonic combustor

Two flow control methods, namely, the microjet array (MJA) and the conventional shock wave generator (SWG), were employed to enhance the chemical reaction in a strut-based supersonic combustor. The Mach number, stagnation temperature, and stagnation pressure of the inflow were 2.92 K, 1650 K, and 2.60 MPa, respectively. The combustor was fueled by hydrogen and liquid kerosene. The case without any flow control method (the baseline case) was also reported for comparison. It was found that the recirculation flow at the strut base was essential to flame stabilization. A self-sustaining hydrogen flame was achieved in the baseline case. However, a flame blow-off was observed when kerosene was injected into the combustor. This was attributed to the short residence time in the recirculation flow and the long ignition delay time of kerosene. To extend the operation envelope of the combustor, additional shock waves were introduced into the combustor using MJAs and SWGs. The shock waves successfully enlarged the recirculation flow at the strut base and significantly enhanced the chemical reaction. Self-sustaining hydrogen/kerosene flames were achieved in the cases with MJAs and SWGs. Although the chemical reaction in the case with MJAs was slightly less vigorous than that in the case with SWGs, these two cases shared the same combustion mode. The MJAs had the potential to replace the conventional SWGs in terms of flame stabilization enhancement.

Drag Control by Hydrogen Injection in Shocked Stagnation Zone of Blunt Nose

The main motivation of the current study is to propose a high-pressure hydrogen injection as an hybrid active flow control technique in order to manipulate the flow-field in front of a blunt nose during hypersonic flight. Hydrogen injection can lead to self-ignition under the right environment conditions in a stagnation zone, and may cause thermal heat addition through combustion and provide the counterjet effect together by pushing bow shock upstream. The axisymmetric numerical simulations for the hemispherical blunt nose are performed at a Mach 6 freestream flow with 10000 Pa pressure and 293 K temperature. The sonic and supersonic hydrogen and air injections are compared for drag reduction at the same stagnation pressure ratio PR and momentum ratio (RMA). The sonic air and hydrogen injection scenarios show similar performance in terms of drag reduction and similar SPM flow features, but hydrogen injection has a mass flow rate 3.76 times lower than air. Supersonic hydrogen injection at Mj 2.94 behaves differently than supersonic air injection and can achieve up to 60 % drag reduction at lower PR and LPM mode with lower mass flow rate. Additionally, air injection achieves a drag reduction of 40 % in SPM mode at higher PR with very high mass flow rate.

Experimental study of Stagnation Pressure Reaction Control for mid-calibre non-spinning projectiles

Increased precision of gun-launched ammunition can reduce collateral damage and increase stand-off distance. It can maximize the tactical advantage of each shot and reduce the number of rounds needed to achieve mission success. Worldwide, technology for controlled smart munitions is being researched and demonstrated in various implementations. Although the use of aerodynamic control fins is an effective control means, such fins are complex to integrate with the projectile, they increase radar reflectivity (with associated risk of counter battery fire) and will reduce payload. The Netherlands Organisation for Applied Scientific Research has been developing a Stagnation Pressure Reaction Control technology without aerodynamic fins, for aerodynamically unstable projectiles. The on-off character of this control technology requires a tuned control algorithm in order to achieve stable and controllable projectile behaviour. This article discusses simulated projectile behaviour under different control algorithms and presents achieved test object performance in a supersonic wind tunnel experiment performed in 2018. The experiment demonstrated feasibility of the combination of the Stagnation Pressure Reaction Control Technology and the tuned control algorithm for a non-spinning, aerodynamically unstable test object, hinged through its centre of mass. In a Mach 2 wind tunnel flow, this test object was stabilized and put under a controlled angle of incidence up to 1.5 degrees with angular error bandwidths up to 0.3 degrees, requiring and realising actuator response times on the order of 2 ms. For higher stable angles of incidence, required for correcting disturbances such as wind gusts, the SPRC technology could be integrated into projectiles with a reduced margin of instability. The short response time would provide such a projectile with a high level of (end game) manoeuvrability, for instance against moving targets.

Lean blowoff behavior of cavity-stabilized flames in a supersonic combustor

Experiments are performed to investigate the near blowoff characteristics of cavity-stabilized flames in an ethylene-fueled scramjet combustor. Heated air enters the combustor with conditions of Ma = 2.52, stagnation pressure p=01.6 Mpa and stagnation temperature T0= 1400 K. The lean blowoff and ignition limits are obtained for different-sized cavity flameholders with upstream flush-wall fuel injection, and the flame structures and dynamics near blowoff are captured with high-speed chemiluminescence imaging. Though the flow residence time is believed to increase with increasing cavity size, it is observed that the lean ignition and blowoff limits do not vary monotonically with the cavity size. Under the same fuel injection conditions, the flame intensity in larger cavities is weaker than that in smaller cavities though the residual flame can survive for an obviously longer period in larger cavities. This may be attributed to the leaner mixture within larger cavities when the lean blowoff limits are approached. Thus, there exists an optimal cavity size which is the most suitable for the ignition and flame stabilization near lean blowoff conditions. This optimal cavity size is determined by the competition between the flow residence time and the local equivalence ratio that greatly depends on the interactions between the fuel jet and the cavity.

Analysis of flow-field in a dual mode ramjet combustor with boundary layer bleed in isolator

A two-dimensional Reynolds averaged Navier Stokes (RANS) simulation of a dual mode ramjet (DMRJ) combustor is performed, modeling the University of Michigan dual-mode combustor experimental setup operating in reacting mode with different equivalence ratios (φ). The simulations are carried out using a k-ω SST turbulence model and a steady diffusion flamelet model for non-premixed combustion. Air enters the isolator at Mach 2.2, stagnation pressure and temperature of 549.2 kPa and 1400 K respectively. Hydrogen is injected transverse to the flow direction and upstream of the cavity flame holder to simulate ramjet (φ = 0.29) and scramjet (φ = 0.19) modes of operation. Wall static pressure plots are used to validate numerical results against experimental data. Analysis of flow separation in ramjet mode due to the presence of a shock train in the isolator is carried out by means of numerical Schlieren images overlapped with contours of negative axial velocity, showing the effects of shock wave boundary layer interaction (SWBLI). Active control through wall normal boundary layer bleed in the separated flow region is implemented, which weakens the shock train and moves it downstream closer to the cavity. Bleed results in an improved stagnation pressure recovery in ramjet mode, with a marginal increase in combustion efficiency.