Oblique Shock Research Articles

Evaluating sustainability power plant efficiency: Unveiling the impact of power plant load ratio on holding steam ejector performance

Understanding the role of holding ejectors is essential in the broader context of environmental and sustainability research. Through rigorous analysis, optimization of ejector usage is applied to minimize resource consumption and promote cleaner, more sustainable production methods. However, the design and control strategies of the power plant, including the setpoints and tolerances of various components, can also contribute to the power plant load ratio (PPLR). This study explores the impact of PPLR in the power plant cycle on the holding ejectors performance under various condenser temperatures. For comprehensive analysis, the study employs a comprehensive framework encompassing various indicators such as condensing fluid behavior, energy consumption, exergy destruction, air suction in the condenser, and entrainment ratio. The research unveils key findings, including a notable increase in oblique shocks with rising PPLR and a quantifiable rise in the maximum liquid mass fraction in the condensed regime. Additionally, the analysis reveals numerical relationships, such as a 5 % increase in mass flow rate and a 4.9 % rise in energy consumption with a PPLR increase from 0 to +5 %. Negative impacts on performance, such as a 4.6 % reduction in air extraction flow rate from the condenser, are observed with PPLR changes from 0 to −5%.

Inner structures of rapid free-surface granular avalanche over a small bump obstacle: Expansion fan, oblique shock wave, and contact anisotropy

This study investigates the inner flow characteristics of a rapid granular avalanche passing over a small bump obstacle fixed on an inclined chute using the discrete element method. Both the cross-sectional mean flow properties, such as free-surface height, mean flow velocity, and mean stresses, and the inner local flow properties, including granular temperature, coordination number, pressure, contact force orientation, and granular fabrics, were comprehensively investigated. Upstream of the obstacle, a wide compression region where mean stresses strengthen and exhibit anisotropy was observed. Employing the kinetic theory of granular gas, we revealed a smooth supersonic-to-subsonic transition near the obstacle, a phenomenon distinct from typical gas dynamics. These upstream flow phenomena are attributed to the generation of stream-wise-oriented contact force chains as the flow impacts the obstacle. Downstream of the obstacle, a complex non-monotonic expansion–compression–expansion process was observed. We demonstrated that this non-monotonic flow process reflects an inner gasdynamic-like phenomenon characterized by an expansion fan followed by an oblique shock wave. Moreover, the force chains and the inner shock structure were found to significantly influence the evolution of stream-wise velocity profiles. These findings underscore the significance of inner flow structures in shaping the dynamics of granular avalanche flow interacting with obstacles.

Aerodynamic performance optimization of a UAV's airfoil at low-Reynolds number and transonic flow under the Martian carbon dioxide atmosphere

Unmanned aerial vehicles (UAV) on Mars operate in a thin atmosphere of carbon dioxide, where the flow Reynolds number and local sound velocity are significantly reduced compared to the Earth's atmosphere because of low gas density and temperature. Therefore, a Mars UAV with a rotary-wing design can operate in a state of low-Reynolds number (Re = 1-5 × 104) and transonic flow (Matip = 0.7–1.2), where flow separation and shock waves occur simultaneously. The interaction between flow separation and shock waves not only makes the entire flow field more complex but also seriously affects the aerodynamic performance of the airfoil, thereby reducing the payload of UAVs, which previous aircraft designs based on the thick atmospheric environment of the Earth have not considered. Therefore, in this study, a numerical simulation method is used to conduct airfoil optimization research on a characteristic airfoil under low-Reynolds number transonic flow (Re = 33200, Ma = 0.85) conditions to acquire additional insight, improve aerodynamic performance of the airfoil, and enhance the payload. The results were compared with the optimization results of subsonic flow (Re = 19500, Ma = 0.50). The Hicks-Henne bump functions method was selected for airfoil parameterization, and non-dominated sorting genetic algorithms (NSGA)-Ⅱ were used for optimization. Through an analysis of the pressure coefficient (Cp), surface friction coefficient (Cf), and density gradient, the number and intensity of oblique shock waves on the upper and lower surfaces were observed to significantly affect the aerodynamic performance of airfoils under low-Reynolds number flow. The changes in flow separation and transition positions do not directly affect the aerodynamic performance but weaken the oblique shock wave intensity induced by the separation point or directly transform strong shock waves into weaker compression waves because of the absence of flow separation. The optimization results for both subsonic and transonic flow point to a thinner overall shape with a flatter surface with the drag greatly reduced due to the disappearance of shock waves on the lower surface, which significantly improves aerodynamic performance. Compared with the original airfoil, at Ma = 0.85, this design can increase the lift coefficient (Cl) by at least 124 % and lift drag ratio (Cl/Cd) by 177 %.

Flow characteristics and propulsive performance of oblique detonation waves induced by a transverse jet

The initiation of oblique detonation waves (ODWs) is a key component of the successful application of oblique detonation wave engines (ODWEs). This paper numerically investigates the initiation of ODWs under the active control of a transverse jet by solving the two-dimensional multi-species Euler equations, focusing on the morphology of the flow fields as well as the relationship between the flow structures and propulsive performance using the concept of thrust potential. Active jet control significantly shortens the initiation length of the ODW. The results reveal that the jet-induced flow field shows four typical patterns depending on the jet momentum flux ratio and wedge angle: shock-induced combustion, a type I pattern, a type II pattern, and a type I-II pattern. For the jet-induced ODW flow field, the propulsive performance declines as the momentum flux ratio increases when the wedge angle is certain. The larger the wedge angle, the greater the magnitude of the decline. The thrust of the flow field consists of two main components: the thrust generated by the mixture that passes first through the oblique shock wave and then through the detonation wave, and the thrust generated by the mixture that passes directly through the ODW front. Since the ODW upstream front has a larger wave angle, the flow loss of the mixture passes through the ODW upstream front is higher and the thrust potential is lower. This work could guide the active control of the initiation of ODWs at low flow losses.

On the determination of the standing oblique detonation wave in an engine combustor using laser absorption spectroscopy of hydroxyl radical

Freejet experiments were conducted in the hypersonic flight-duplicated shock tunnel (JF-12) to understand the detonation combustion mode in the combustor of a standing oblique detonation ramjet (Sodramjet) engine prototype. Besides the traditional schlieren imaging to capture the oblique shock positions in the combustor, we implemented laser absorption spectroscopy of hydroxyl (OH) radical with three laser beams around the shock front. The Sodramjet engine combustor provides an ideal condition to measure OH because its molar fraction within the detonation front is estimated to exceed 0.03, which is much higher than other combustors. However, the harsh environment of freejet shock tunnel test poses a great challenge to the reliability of measurement systems. In this work, near-infrared tunable laser diode sensors centered at 1528 nm were chosen to measure the line-of-sight average of OH partial pressure. Continuous absorption measurements were conducted near the shock front, indicating a thin OH region after shock compression that could only be achieved by detonation combustion. The time-varied OH partial pressure was well predicted by the computational fluid dynamics (CFD) prediction. Our measurement results verified the detonation combustion of Sodramjet engine prototype and showed that in situ, real-time detonation diagnostics based on laser absorption spectroscopy is possible in freejet experiments of shock tunnel.

Design and aerodynamic characteristics of variable-geometry hypersonic inlet based on shock wave elimination

To solve the problem that the control of shock wave elimination is weakened under off-design conditions, the design concept of a variable-geometry inlet scheme that combines the variable-geometry cowl (translating and diagonalizing) with regulating shock wave elimination is introduced in this paper. The variable-geometry inlet is designed by the theories of oblique shock wave and isentropic wave as well as the Oswatitsch theory. Regulatory law of the variable-geometry cowl based on shock wave elimination is obtained by the geometric relationships between cowl compression angle, cowl shock wave angle, and optimal control point or range. Numerical simulations are conducted to investigate flow field characteristics, control mechanism, and working performance of the inlet. Results reveal that expansion waves have a significant impact on the cowl shock wave and boundary layer interaction, and flow separation. Furthermore, variable-geometry inlet with translating and diagonalizing cowl based on the regulation of shock wave elimination effectively controls and even completely eliminates the flow separation. In terms of inlet performance, the total pressure loss of the variable-geometry inlet decreases such that the total pressure recovery coefficients of the translating cowl and diagonalizing cowl inlets are increased by maximum values of 3.39 % and 9.97 %, respectively. However, the mass flow coefficient of translating cowl inlet decreases, whereas that of the diagonalizing cowl inlet is equivalent to that of the fixed-geometry inlet. The working range can be widened by changing the internal contract ratio of the inlet through translating or diagonalizing the cowl. The results confirm that the scheme of variable geometry inlet with diagonalizing cowl is practicable and reliable and has important guiding significance and value for inlet design.

Fluid–Structure Interactions between Oblique Shock Trains and Thin-Walled Structures in Isolators

Understanding aeroelastic issues related to isolators is pivotal for the structural design and flow control of scramjets. However, research on fluid–structure interactions (FSIs) between thin-walled structures and the isolator flow remains limited. This study delves into the FSIs between thin-walled panels and the isolator flow, as characterized by an oblique shock train, by quantitatively analyzing 11 flow parameters assessing the structural response, separation zones, shock structures, flow symmetry, and performance. The results reveal that an FSI triggers panel flutter under oblique shock train conditions, with the panel shapes exhibiting a combination of first- and second-mode responses, peaking at 0.75 of the panel length. Compared to rigid wall conditions, isolators with a flexible panel at the bottom wall experience downstream movement of the separation zones and shock structures, reduced flow symmetry, and minor changes in performance. Transient fluctuations occur due to the panel flutter. Two flexible panels at the top and bottom walls have a comparatively lesser influence on the averaged parameters but exhibit more violent transient fluctuations. Furthermore, the FSI effects under oblique shock train conditions are contrasted with those under normal shock train conditions. The flutter response under normal shock train conditions is more pronounced, with a larger amplitude and higher frequency, driven by the heightened participation of the first-mode response. The effects of FSIs under normal shock train conditions on the averaged parameters are the opposite (with a larger influence) to those under oblique shock train conditions, with significantly more drastic transient fluctuations. Overall, this study sheds light on the complex and substantial influence of FSIs on the isolator flow, emphasizing the necessity of considering FSIs in future isolator design and development endeavors.

Effects of the outlet contraction ratio on the performance of liquid kerosene/air two-phase rotating detonation combustors

Experiments and numerical simulations were conducted across a range of contraction ratio to investigate the detonation combustion characteristics of liquid aviation kerosene/air rotating detonation combustor. The research indicates that extremely low and high contraction ratio adversely affect the stable operation of the two-phase rotating detonation combustors. At a contraction ratio that is too low (19.03 % in the RDC discussed in this paper), the atomization and mixing properties of the two-phase fuel are poor, with a lower reaction pressure in the fresh mixture, hindering the initiation of detonation waves. On the other hand, a contraction ratio (85.7 % in the RDC discussed in this paper) that is too high enhances the energy of the detonation wave; it triggers oblique shockwave feedback, preventing the fresh mixture from entering the combustor and thus leading to the quenching of detonation waves. With an increase in the contraction ratio of outlet configurations, there is a noticeable rise in both the peak pressure and combustion efficiency of detonation waves. In contrast, the wave height gradually decreases, and the wave velocity shows a pattern of initially increasing and then decreasing. Along the axial position inside the two-phase rotating detonation combustor, the upstream detonation wave is caused by the reflected shockwaves increasing the pressure of the upstream fresh mixture. In contrast, the downstream detonation wave results from increased fresh mixture pressure and enhanced fuel atomization and mixing due to the increased contraction ratio of outlet configurations and wave feedback. So, the detonation wave exhibits a strong-weak-strong distribution characteristic.

Effect of perforated wall in controlling the separation due to SWBLI at Mach no. 5 to 9

Abstract The mixed compression scramjet engine intake was numerically simulated to study the characteristic of shock wave induced boundary layer separation leading to the formation of the separation bubble (SB). The analysis employed a 2D-RANS method with SST k-ω turbulence model at different Mach numbers. The intake is designed as a three-ramp intake for improved performance. As the Mach number varies, the size of the separation bubble formed due to the interaction of oblique shock waves with the boundary layer also varies, affecting both intake the efficiency and overall efficiency of the engine. Apart from traditional control techniques, the most preferred bleed technique is incorporated. Localised & distributed bleed techniques are designed and analysed within the intake at different locations. These techniques result in a reduction in the size of the separation bubble within the intake. Establishing perforation in the engine intake also increases the intake efficiency and overall engine performance.

Research on the Combustion Mode and Thrust Performance of Rotating Detonation Scrarmjet Engines

Decades of research in hypersonic science and engineering have shown that scramjet engines are the preferred choice for powering hypersonic vehicles. In these engines, oblique shock combustion and rotating detonation combustion are two common combustion modes. This article primarily focuses on the numerical simulation and theoretical analysis of these two combustion modes. The research indicates that as the combustible mixture injected into the combustion chamber increases, the combustion mode transitions from deflagration combustion to self-sustaining detonation combustion and then to forced detonation combustion to maintain stable combustion in the chamber. Under self-sustaining detonation combustion mode, as the amount of combustible mixture injected into the chamber increases, the number of detonation waves gradually increases from one to multiple and the angle between the detonation waves and the inflow changes from acute to right angles. In the forced detonation combustion mode, the minimum shock wave intensity required to initiate detonation combustion can be obtained by drawing a tangent from the Rayleigh line. The structure generating oblique shock waves in the combustion chamber needs to match the shock wave intensity. In the detonation combustion mode, both the exit total temperature and specific impulse of the combustion chamber increase with the increase in Mach number at the inlet, while the pressure ratio decreases. This article provides a reference for the design of the combustion chamber in supersonic combustion ramjet engines.

Influence of double-side symmetric oblique laser shock peening on shape deviation, surface integrity, and fatigue properties of the blades in small-sized blisk

Double-sided symmetric oblique laser shock peening (DSOLSP) was adopted to experiment on a small-sized blisk with restricted space. The influences of different laser energies on the shape deviation, surface roughness, microhardness, and residual stress of the blades were studied. The optimum process parameters were selected based on the shape deviation results of the blades and the microstructural evolution of the surface layers was investigated. The strengthening effect of DSOLSP treatment on notched blades was evaluated by vibration fatigue tests. It was found that DSOLSP can effectively cover the fatigue vulnerable regions without interference. Two-way bending deformation was formed after DSOLSP, and the shape deviation of the blades could be controlled within ±0.02 mm. With the increase of laser energy, the surface roughness and the depth of work-hardened layer increased. The full-thickness compressive residual stress (CRS) field was induced. A gradient microstructure was generated on the surface of the blade, and the size of average nanograins on the topmost surface is approximately 36.5 nm. The fatigue strength of the DSOLSP treated blades with notches was increased by 25.9 %. The enhancement can be attributed to the synergistic strengthening of CRS and gradient microstructure.

On the aeroelastic bifurcation of a flexible panel subjected to cavity pressure and inviscid oblique shock

The aeroelasticity of a panel in the presence of a shock is a fundamental issue of great significance in the development of hypersonic vehicles. In practical engineering, cavity pressure emerges as a crucial factor that influences the nonlinear dynamical characteristics of the panel. This study focuses on the aeroelastic bifurcation of a flexible panel subjected to both cavity pressure and oblique shock. To this end, a computational method is devised, coupling a high-fidelity reduced-order model for unsteady aerodynamic loads with nonlinear structural equations. The solution is meticulously tracked by continuous calculations. The obtained results indicate that cavity pressure plays a pivotal role in determining the bifurcation and stability characteristics of the system. First, the system exhibits hysteresis behaviour in response to the ascending and descending dynamic pressures. The evolution of hysteresis behaviour originates from the phenomenon of cusp catastrophe. Second, variations in cavity pressure induce three types of bifurcation phenomena, exhibiting characteristics akin to supercritical Hopf bifurcation, subcritical Hopf bifurcation and saddle-node bifurcation of cycles. The system's response at the critical points of these bifurcations manifests as long-period asymptotic flutter or explosive flutter. Lastly, the evolution of the dynamical system among these three types of bifurcations is an important factor contributing to the discrepancies observed in certain research results. This study enhances the understanding of the nonlinear dynamical behaviour of panel aeroelasticity in complex practical environments and provides new explanations for the discrepancies observed in certain research results.

Improvement of fatigue life of Ti-6Al-4V alloy treated by double-sided symmetric oblique laser shock peening

To improve the anti-foreign object damage (FOD) ability of Ti-6Al-4V alloy blisk blades, a blade simulation specimen was designed and subjected to double-sided symmetrical oblique laser shock peening (DSOLSP) treatment. Surface roughness, residual stress distribution, and microstructural evolution of simulated specimens of Ti-6Al-4V alloy without and with DSOLSP were investigated, and the strengthening effect was evaluated by flexural fatigue test. The results showed that after DSOLSP treatment, surface roughness of simulated specimens was reduced from Ra0.97 µm to Ra0.6 µm. DSOLSP treatment generated a compressive residual stress (CRS) field along the entire thickness direction of the simulated specimen. Nanostructures with a depth of ∼500 nm was observed at topmost surface layer, and the average size of the nanograins was about 35.2 nm. Compared with the base material, the flexural average fatigue life of the simulated specimen was increased by 8.45 times. Moreover, mathematical statistical analysis confirms that DSOLSP has a significant effect on the fatigue life of simulated samples. The development of CRS fields throughout the entire thickness direction, the grain refinement on the surface, and the reduction of surface roughness are primarily responsible for the enhancement of the flexural fatigue life of the DSOLSP simulated specimen.

Effects of high-temperature and entropy layer of Mach 12 oblique shock wave

In the current study, we numerically investigate the high-temperature inviscid flows over the 15-degree wedges with the sharp and blunt leading-edges. Firstly, the vibrational non-equilibrium effects in sharp wedge flow with different deflection angles are analyzed. Then, the influence of the vibrational non-equilibrium effects on the flow parameters behind the oblique shock wave is provided. Lastly, we discuss the high-temperature non-equilibrium flow over the 15-degree wedge with the entropy layer.

A study on shock-wave/boundary-layer interaction with varying flow deflection angle

Shockwave boundary layer interaction is studied for oblique shocks generated from a shock generator with a varying flow deflection angle. A two-dimensional shock generator with a wedge-shaped geometry is fixed inside the test section and mounted to the wind tunnel's top wall. The shock generator is made to pitch up and down so that the net flow deflection angle changes because of the pitch motion. As a result, the shock strength changes for a varying flow deflection angle of the shock generator with a fixed incoming Mach number of the flow. The imposed adverse pressure gradient on the boundary layer by this incident shock changes, which influences the shape and structure of the shockwave boundary layer interaction. Three different shockwave boundary layer interaction structures are observed with high-speed schlieren from the experiments conducted for the current study. Schematics for these structures' types are illustrated, and a model is presented to build the transition criteria for the flow deflection angles for which the transition would take place from one structure to the other. The two transition criteria, one for type 1 to type 2 transition and the other for type 2 to type 3 transition, are validated with the observed results from the experiments. Assumptions made for the transition criteria are also validated with experimental results.

Numerical study on altitude-compensating mechanism of a permeable nozzle

A conceptual adaptive altitude compensation nozzle with a permeable wall (permeable nozzle for short) is proposed in the present study. The typical operating states of the nozzle are demonstrated using numerical methods and the underlying mechanism of its altitude compensation is interpreted. The numerical results show that, at low nozzle pressure ratios, the external atmosphere enters the nozzle through its permeable wall section, which inhibits the backflow at the nozzle wall, weakens the strength of the separated oblique shock, thus improves the low-altitude thrust performance of the nozzle. However, at high nozzle pressure ratios, the exhaust gas leaves the nozzle through its permeable wall, which weakens the expansion of the exhaust gas and reduces the thrust performance. The permeable nozzle is considered to possess altitude compensation ability when the enhancement in its low-altitude thrust performance surpasses the deterioration in high-altitude thrust performance. The thrust composition analysis of the permeable nozzle reveals that its altitude compensation ability derives from the internal pressure rise due to the atmospheric air penetration into the nozzle at low nozzle pressure ratios.

Centralized gradient-based reconstruction for wall modeled large eddy simulations of hypersonic boundary layer transition

In this study, we introduce a robust central Gradient-Based Reconstruction (GBR) scheme for the compressible Navier-Stokes equations. The method leverages transformation to characteristic space, allowing selective treatment of waves from the compressible Euler equations. By averaging left- and right-biased state interpolations, a central scheme is achieved for all but the acoustic waves, which require upwinding for stability. Distinct differences were observed between transformations using either primitive or conservative variables. We evaluated the method's robustness and superiority using benchmark problems, including the two-dimensional shock entropy problem, two-dimensional viscous shock tube, and three-dimensional inviscid Taylor-Green vortex. Subsequently, we assessed the method in the context of Wall Modeled Large Eddy Simulations (WMLES), where coarse grids are used to reduce computational cost but also introduce substantial numerical dissipation. Using WMLES, we simulated oblique shock impingement on a Mach 6 disturbed boundary layer and a Mach 7.7 flow over a 15∘ compression ramp. Our findings reveal that: 1) transformation to characteristic space using conservative variables leads to more accurate results; 2) minimizing numerical dissipation through centralized interpolation is crucial. In the compression ramp case, boundary layer separation was shifted slightly upstream, and there was an over-prediction of wall heating, likely attributable to the equilibrium-assuming wall model. Overall, this work showcases the method's potential in accurately capturing complex flow dynamics with reduced numerical dissipation.

Design and Analysis Aerospike Blunt Body at Hypersonic Speed

Hypersonic vehicles have the drag and heating problem when they operate at high speeds, to avoid or minimize this problem, aero spikes are utilized. This investigation finds a solution for reducing the drag and heating. Using ANSYS fluent designed an aero spike, which has flow around a blunt body at hypersonic Mach number 6 with angle of attack zero. It is analyses for L/D ratio of 1.5 and 2. The front of the body is replaced by an aero spike which has weak oblique shocks and well-built detached shock. This develops a circular region between shock and blunt body. This works as a streamlined profile to reduce heat as well as drag.

Numerical Investigation of Supersonic Flow over a Wedge by Solving 2D Euler Equations Utilizing the Steger–Warming Flux Vector Splitting (FVS) Scheme

Supersonic flow over a half-angle wedge (θ = 15°) with an upstream Mach number of 2.0 was investigated using 2D Euler equations where sea level conditions were considered. The investigation employed the Steger–Warming flux vector splitting (FVS) method executed in MATLAB 9.13.0 (R2022b) software. The study involved a meticulous comparison between theoretical calculations and numerical results. Particularly, the research emphasized the angle of oblique shock and downstream flow properties. A substantial iteration count of 2000 iteratively refined the outcomes, underscoring the role of advanced computational resources. Validation and comparative assessment were conducted to elucidate the superiority of the Steger–Warming flux vector splitting (FVS) scheme over existing methodologies. This research serves as a link between theoretical rigor and practical applications in high-speed aerospace design, enhancing the efficiency of aircraft components.

Ten-moment fluid model with heat flux closure for gasdynamic flows

The gasdynamic 10-moment equations are revisited, applying a Chapman-Enskog type expansion to achieve the closure for heat flux. The resulting gradient-based closure performs favorably against previously employed closures in regions of shocks when compared to the kinetic results obtained from Direct Simulation Monte Carlo simulations. The 10-moment model can capture finite kinetic effects due to non-Maxwellian velocity distribution functions, as compared to the 5-moment Navier-Stokes equations, and remains hyperbolic-parabolic with real wave speeds under all initial conditions. Self-consistent kinetic boundary conditions are derived without consideration of a distinct Knudsen layer. The model is applied to a number of canonical gasdynamics problems in one and two dimensions such as steady normal and oblique shocks, as well as the Sod shock tube problem.