Sidewall Compression Research Articles

Numerical study of a scramjet isolator performance under different sidewall compression angles

Scramjet engines, also known as supersonic combustion ramjet engines, are frequently regarded as a compelling alternative for launching payloads into Earth's orbit. These air-breathing engines have streamlined designs with minimal movement of components. However, the successful design of scramjet engines necessitates overcoming various challenges such as managing the high heat fluxes and pressure loads exerted on the engine walls. Additionally, addressing issues such as shockwave-boundary-layer interactions and the potential occurrence of choked flow within the isolator channel are critical considerations during the scramjet design process. Therefore, this study aims to evaluate sidewall compression in the isolator region to deal with the high heat fluxes and pressure loads inside the scramjet isolator. In addition, this work also investigates how the variation in the angle of attack influences the mass flow rate of the intake and at which range of the angle of attack the intake becomes choked. The CFD analyses include contour images of properties such as Mach number, total pressure, heat flux, and pressure distribution on the walls, and the calculation of performance parameters, including the analysis of the second law of thermodynamics. The study involved varying the compression angle within the range of 4° to 10°. The results of this study demonstrate that implementing sidewall compression in the isolator region allows for the effective management of the position of the heat flux and pressure peaks on the upper wall of the isolator. Regarding the pressure distribution along the upper wall of the isolator, the 10°case presented a pressure peak of approximately 130000 Pa while the 4°case presented 155000 Pa. In addition to this significant decrease in the pressure peak value, its location also changed, with an increase of approximately 8 mm downstream of the isolator by decreasing the compression angle from 10° to 4°. This engineering approach presents a viable solution for mitigating the challenges posed by high heat flux and pressure loads in the intake section. The cost of applying such a solution is to decrease the intake performance – a decrease of approximately 30 % in the isentropic efficiency when comparing a case with no sidewall compression with the sidewall compression cases. In the choked flow study, angles of attack ranging from 4 to 30°were considered. The analysis shows that the choked-flow condition gradually occurs as the angle of attack increases beyond 4°, owing to the shock-on-lip condition. The results at approximately 20° indicate that the isolator becomes completely choked once the mass flow rate abruptly decreases – from around 0.30 to 0.15 Kg/s when comparing the 20°-of-AoA case with the 30° one. This work aims to contribute to the early phase of engine design by avoiding critical failures in the scramjet structure owing to aerodynamic load, thermal stress, and engine unstart.

Numerical analysis on the conceptual design of an air-breathing engine inlet working in mach number 3∼5.5

Based on the equal-intensity shock theory, this article designed a supersonic inlet working in Mach number 3.0∼5.5 with the background of an air-breathing engine. The inlet applied the four-shock train mixed compression configuration and inserted a sidewall compression at the beginning of the isolator. Through developing effective 3D RANS computations validated by current experiments, the performance of the designed inlet was identified. The designed inlet self-starts at freestream Mach number Ma∞ = 3.0 under which the total pressure recovery coefficient has dramatic increment, and the aerodynamic choking at the inlet throat no longer presents; the inlet keeps working at all studied flight states with zero angle of attack (AoA) and achieves shock-on-lip at the design point Ma∞ = 5.0. Both positive and negative AoAs can disrupt the equal-intensity shock allocations, which degrade the inlet performance. The inlet obtains maximum total pressure recovery coefficient at zero AoA. The maximum back pressure at Ma∞ = 3.0∼5.5 obtained by the inlet surpasses the requirements and keeps a certain margin. The inlet performance basically meets all the goals proposed by the engine.

Research and development on inlets for rocket based combined cycle engines

Rocket based combined cycle (RBCC) engine is regarded as one of the most efficient propulsion systems for reusable space transportation vehicles and hypersonic cruise vehicles. It can perform excellently over a very broad range of Mach numbers and flight altitudes through multi-mode operations by effectively integrating high thrust-to-weight ratio of the rocket engine and high specific impulse of the airbreathing engine. The RBCC inlet, commonly operating in an extremely wide flight speed range covering static, subsonic, transonic, supersonic, and hypersonic, plays a vital role in the RBCC propulsion system and flight vehicles. This paper presents a comprehensive and systematical review of the research progress on the inlets for RBCC engines in different countries and different historical periods. A number of key technical and scientific issues specially concerning RBCC inlets are sorted out, including the wide range RBCC inlet design and optimization, inlet/embedded rocket/combustor matching and control, dynamic operation, etc. Several typical RBCC inlet applications and the corresponding research are discussed, including the Strutjet inlet, A5 inlet, GTX inlet, and the Astrox's inlet in the U.S., the RBCC inlet in JAXA, and a sidewall compression RBCC inlet in China. Finally, experience gained and some recommendations are summarized for future research of the RBCC inlet.

Flowfield characteristics in sidewall compression inlets

In this work, the flow fields in three-dimensional (3D) sidewall compression inlets are investigated theoretically and numerically. Combining spatial dimension reduction theory with the shock dynamics theory, a detailed analysis of the flowfield in the compression part of an inlet is proposed. Using this analytical approach, wave configurations are determined and flowfield parameters are solved; numerical simulations are conducted to validate these theoretical results. The effects of the attack angle and yaw angle on the pressure ratio, wave structures, and total pressure recovery performance are discussed in detail. Findings offer an effective way to study shock/shock interaction (SSI) in sidewall inlets, which is essential to the designing of such inlets.

Experimental Study of Single Expansion Ramp Nozzle Performance Using Pitot Pressure and Static Pressure Measurements

In order to overcome the drag at hypersonic speed, hypersonic flight vehicles require a high level of integration between the airframe and the propulsion system. Propulsion system based on scramjet engine needs a close interaction between its aerodynamics and stability. Hypersonic vehicle nozzles which are responsible for generating most of the thrust generally are fused with the vehicle afterbody influencing the thrust efficiency and vehicle stability. Single expansion ramp nozzles (SERN) produce enough thrust necessary to hypersonic flight and are the subject of analysis of this work. Flow expansion within a nozzle is naturally 3D phenomena; however, the use of side walls controls the expansion approximating it to a 2D flow confined. An experimental study of nozzle performance traditionally uses the stagnation conditions and the area ratio of the diverging section of the tunnel for approaching the combustor exit conditions. In this work, a complete hypersonic vehicle based on scramjet propulsion is installed in the test section of a hypersonic shock tunnel. Therefore, the SERN inlet conditions are the real conditions from the combustor exit. The performance of a SERN is evaluated experimentally under real conditions obtained from the combustor exit. To quantify the SERN performance parameters such as thrust, axial thrust coefficient Cfx and lift L are investigated and evaluated. The generated thrust was determined from both static and pitot pressure measurements considering the installation of side walls to approximate 2D flow. Measurements obtained by a rake show that the flow at the nozzle exit is not symmetric. Pitot and pressure measurements inside the combustion chamber show nonuniform flow condition as expected due to side wall compression and boundary layer. The total axial thrust for the nozzle obtained with the side wall is slightly higher than without it. Static pressure measurements at the centerline of the nozzle show that the residence time of the flow in the expansion section is short enough and the flow of the central region of the nozzle is not altered by the lateral expansion when nozzle configuration does not include side walls.

Experimental Investigation of Sidewall Compression and Internal Contraction in a Scramjet Inlet

In this paper, we present the results of an extensive measurement campaign to investigate the effects of external and internal sidewall compression and the variation of internal contraction on the performance and flowfield of a scramjet inlet. Experiments were conducted in the H2K wind tunnel of DLR, German Aerospace Center, in Cologne, Germany, at Mach 7. The performance was evaluated by static and total pressure ratios, kinetic energy efficiency, and mass capture ratios. The flowfield was analyzed using wall pressure distributions, pitot pressure, and Mach number profiles at the isolator interface to the combustion chamber and infrared thermography on the external ramps. The results show that the combination of a two-ramp inlet with external sidewall compression is not suitable for increasing the inlet’s compression capability as it induces strong separation and vortex structures in the external part, which strongly increase spillage and impair the starting behavior. Thus, no significant increases in internal contraction are possible, which inhibits any gains in the performance of the inlet. With internal sidewall compression, strong increases of the pressure ratio can be achieved at the cost of total pressure losses. With internal sidewall compression as well as without sidewall compression, the inlet is still self-starting at internal contraction ratios well above the Kantrowitz limit.

Flow Mechanism in a Hypersonic Sidewall Compression Inlet with a Rectangular-to-Circular Isolator

The flow phenomenon and mechanism in sidewall compression hypersonic inlets with a rectangular-to-circular isolator are investigated in this paper. The two schemes were introduced to study the flowfield of the inlets with rectangular-to-circular cross-section shape transition. The results show that 1) the gross performance of inlets with a rectangular-to-circular isolator are better than the original inlet; 2) the common feature of the crossflow in the isolator of the three schemes is the existence of a pair of counter-rotating flow-directional vortices near the bottom wall, rotating on the boundary layer toward each other; 3) scheme 3 has a stronger sidewall vortex than scheme 2 (it produces a rolldown effect to improve the low-speed flow and the corner flow); and 4) the rectangular-to-circular process on sidewall compression inlets does not affect the startability. The backpressure limit of scheme 1 is the highest, which is significantly higher than schemes 2 and 3. This is mainly because the boundary layer accumulates more easily near the bottom of the symmetry plane in the schemes with a rectangular-to-circular process.

Numerical investigation on hydrogen combustion in a scramjet with 3D sidewall compression inlet

Numerical simulations were employed to analyze the flowfield of a scramjet with three-dimensional (3D) sidewall compression inlet, and the effect of inlet distortion on the mixing and combustion process was examined. The numerical approach solved the compressible Reynolds Averaged Navier–Stokes (RANS) equations supplemented with a finite rate chemical reacting model for the combustion of hydrogen fuel and air. Turbulence closure was achieved using Menter shear-stress transport (SST) model. To verify the accuracy of the simulation, the computed wall pressure was compared with the experimental data of the direct-connect combustor test. The metrics employed in the simulations included qualitative assessments related to flow structure as well as quantitative values of fuel mixing efficiency, combustion efficiency and static pressure distribution. Intake sidewalls were found to strongly affect the inlet flow structure, which became more complex in the nonuniform flowfield. The shock train system affected the combustion region located upstream of the injection and led to pairs of asymmetric separation bubbles. Nevertheless, the shock train system dissipated due to the reactions, the combustion patterns of each fuel jets in downstream region were nearly identical, and the degree of improvement of mixing and combustion efficiency near the downstream injectors was less than that near the upstream injectors.

Effects of Sidewall Compression and Relaminarization in a Scramjet Inlet

This paper presents the numerical simulations and the performance analysis of a scramjet inlet as part of a combined experimental and numerical study. A well-validated finite volume flow solver was used to simulate a scramjet inlet with a double ramp configuration for outer compression, including varying degrees of sidewall compression. The computed wall pressure and heat transfer in the symmetry plane are in close agreement with the measurements, and the numerical results indicate that the weak sidewall compression alters the inlet performance significantly. The effects of partial relaminarization over the expansion corner, before the interior part of the inlet, is isolated and investigated in both the experiment and simulation. It is shown that relaminarization of a boundary layer is predicted accurately using the current numerical methods. This work represents a contribution to the understanding of the effects of sidewall compression and relaminarization in designing a scramjet inlet.

Mechanism and Control of Combustion-Mode Transition in a Scramjet Engine

A sidewall compression scramjet engine operated in two combustion modes under Mach 6 flight condition, weak- and intensive-combustion modes. The weak mode occurred below the overall fuel equivalence ratio (Φ) of around 0.4. Transition from the weak mode to the intensive mode occurred at Φ ~ 0.4, accompanied by a sudden increase in thrust. Mechanisms of the transition were numerically investigated in this study. Our simulations captured the sudden increase in thrust at the mode transition. In the weak mode, combustion occurred in only a region near the topwall where an igniter was installed. The combustion region expanded toward the cowl with boundary-layer separation at the mode transition. Our simulations demonstrated that low ignition capability resulted in the weak mode. We demonstrated that the presence of additional igniters on the sidewalls improved the ignition capability and achieved the intensive mode in the entire Φ range.

Experimental Investigation of a Hypersonic Inlet with and Without Sidewall Compression

The inlet of a Mach 8 cruise vehicle has been experimentally investigated in a hypersonic blow down wind tunnel with test times of up to 30s. Two test series have been carried out, the first one with a two-dimensional inlet configuration and the second one with the inlet featuring additional side-wall compression. The objective of the tests was to determine the performance of the inlet at on- and off-design conditions, i.e. at angle of attack and angle of yaw. Additionally, the starting behavior of the inlet has been determined, i.e. the contraction ratio at which the inlet starts. Results were interpreted from schlieren visualizations, IR thermography, static and Pitot pressure readings and mass flow measurements. The off-design investigation of both inlet configurations revealed the sensitivity of the inlet towards a cowl shock induced separation on the second compression ramp. If the cowl shock was not ensured to fall into the bleed duct behind the second ramp, no inlet flow could be established. The comparison of starting contraction ratios showed the advantage of the inlet with side-wall compression, i.e. the inlet with sidewall compression starts at higher contraction ratios.

Structural and hydrodynamic simulation of an acute stenosis-dependent thrombosis model in mice

Platelet activation under blood flow is thought to be critically dependent on the autologous secretion of soluble platelet agonists (chemical activators) such as ADP and thromboxane. However, recent evidence challenging this model suggests that platelet activation can occur independent of soluble agonist signalling, in response to the mechanical effects of micro-scale shear gradients. A key experimental tool utilized to define the effect of shear gradients on platelet aggregation is the murine intravital microscopy model of platelet thrombosis under conditions of acute controlled arteriolar stenosis. This paper presents a computational structural and hydrodynamic simulation of acute stenotic blood flow in the small bowel mesenteric vessels of mice. Using a homogeneous fluid at low Reynolds number (0.45) we investigated the relationship between the local hydrodynamic strain-rates and the severity of arteriolar stensosis. We conclude that the critical rates of blood flow acceleration and deceleration at sites of artificially induced stenosis (vessel side-wall compression or ligation) are a function of tissue elasticity. By implementing a structural simulation of arteriolar side wall compression, we present a mechanistic model that provides accurate simulations of stenosis in vivo and allows for predictions of the effects on local haemodynamics in the murine small bowel mesenteric thrombosis model.

Experimental Investigation of a Two-Dimensional and a Three-Dimensional Scramjet Inlet at Mach 7

The current experimental study at Mach 7 was carried out to investigate the viscous flow effects in a scramjet inlet with a fixed geometry operating at Mach 7.5-8. To reduce the risk of inlet unstart, an optional passive boundary-layer bleed was integrated at the throat and tested for different bleed dimensions. The passive bleed significantly reduces the lip-shock-induced separation bubble on the ramp. But the correct location of the bleed gap is important for the reduction of the separation bubble. To investigate the compression behavior, different backpressures were applied and the effects on the internal flow structure were analyzed by means of static pressure along the isolator and pitot pressure profiles at the exit of the isolator. Because the reference configuration has a very small internal contraction ratio to guarantee self-starting, the internal compression ratio was increased in three steps by means of additional sidewall compression along the internal flowpath. This includes internal contraction ratios above the starting limit. This measure led to a significant increase of the static pressure ratio at the end of the isolator. The complex flow structures are explained by means of detailed static and pitot pressure plots.

Performance enhancement of three-dimensional hypersonic inlet with sidewall compression

Extensive tests and analysis have been implemented to enhance the overall performance of the three-dimensional hypersonic inlets with sidewall compression. The goal of this research is to improve the mass flowrate (MFR) at design point, while keeping good starting characteristics at low Mach number. A series of inlet models have been designed and tested at Mach 4 and 6 in a free-jet tunnel. The present study focuses on the influence of the cowl deflection angle and the position of the sidewall compression on the inlet overall performance. The results indicate that the cowl deflection angle had great influence on the inlet starting characteristics, and the sidewall compression allocation is a key point of elevating the MFR at design Mach number. Remarkably, enhancement of the overall performance of the three-dimensional hypersonic inlet has been achieved.

Effects of body-side compression on forward-swept sidewall compression inlets

Forward-swept sidewall compression inlets have been tested at Mach 4 to study the effects of body-side compression surfaces on inlet performance. The measurements include mass flow capture and mean surface pressure distributions obtained during simulated combustion-induced pressure increases downstream of the inlet. Inlet performance is evaluated using starting and unstarting characteristics, mass capture, mean surface pressure distributions, and permissible back-pressure limits. The results indicate that inlet performance can be improved with selected body-side compression surfaces placed between the inlet side walls.

側壁圧縮型極超音速インレットの内部流れ場に関する実験的研究

Sidewall compression type inlet models for the Scramjet engine were tested in NAL's hypersonic wind tunnel at Mach 7.1. Its contraction ratio, sweep angle, entrance Mach number and cowl size were set in variable manner along the experimental plan. The inlets were investigated with the flow visualization and pressure measurement techniques. As results, without the sweep angle the pressure near the open side was less than near the body side due to the spill from the inlet. It had agreement with the 2-dimensional calculation near the body only. With the sweep angle 45 degree, the lower pressure region due to the spill became small and the performance was higher. The some kinds of the flow field in the inlet owing to the variation of the configuration and basic data of the sidewall compression type inlet were obtained.

Hypersonic crossing shock-wave/turbulent-boundary-layer interactions

Experimental data for two three-dimensional intersecting shock-wave/turbulent-boundary-layer interaction flows at Mach 8.3 are presented. The test bodies, composed of two sharp fins fastened to a flat plate test bed, were designed to generate flows with varying degrees of pressure gradient, boundary-layer separation, and turning angle. The data include surface pressure and heat transfer distributions as well as mean flowfield surveys both in the undisturbed and interaction regimes. The persistence of an extensive low-pressure region throughout the flowfield demonstrates that a sidewall compression inlet is not an efficient pressure increasing device. The data have been obtained in sufficient detail to validate existing or future computational models of these hypersonic flows.

Numerical study of the effects of reverse sweep on scramjet inlet performance

A comparative numerical study of performance parameters of a similar and an opposite sweep sidewall compression inlet is made. The focus of the study is the investigation of the impact of alternate backward-forward sweep on the compression sidewalls as opposed to back-ward sweep on all the sidewalls. Two equivalent scramjet inlet configurations are designed for this purpose. These inlets have the same wetted areas of compression and expansion and same height and width; but in one inlet all the compression surface are swept back (similar sweep inlet) whereas in the other inlet, alternate surfaces are swept backward and forward (opposite sweep inlet). The cowl closure in both cases begins at the start of the throat region. A three-dimensional Navier-Stokes code is used to calculate the flow through these inlets. Results of these calculations are used to compare the two designs for their performance and flow quality. Effects of boundary- layer ingestion on the performance and overall flow features are also investigated.

Numerical and experimental investigation of airframe-integrated inlet for high velocities

Airframe-integrated, three-dimensional fixed geometry inlets for hypersonic Mach numbers are under consideration. The numerical investigation is conducted using the three-dimensional Euler's code. The program calculates local and integral flowfield performances in duct taking into account displacement thickness along the wetted surface on the condition that the flow is nonseparated and the gas is in thermodynamic equilibrium. Sweep sidewall compression inlets with and without struts and transformed types of three-dimensi onal inlets have been examined. The results of flowfield calculations at the exit section and along the duct as a function of Mach number M are given. Two models of three-dimensi onal inlets with and without struts have been tested at A/0, = 2-6. The results of numerical and experimental investigations are compared. Nomenclature Ac = inlet frontal area AN = inlet passage area normal to axis Ox AT = area of captured stream tube at the inlet throat A T = relative throat area, = A T/A c