Oblique Shock Train Research Articles

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.

Asymmetry of oblique shock train and flow control

An oblique shock train generally forms an asymmetric structure in a Mach-2.7 flow field within a duct. To study the flow structure and interaction between oblique shock trains and upstream shocks, a ramp with equal width was installed inside a Mach-2.7 straight duct to generate an incident shock and an oblique shock train interaction. A Schlieren system, transient pressure measurements and particle image velocimetry were used to capture quantitative and qualitative shock structure information. Results show that the asymmetric separation deflection of the oblique shock train occurs randomly in the symmetrical straight duct. The separation deflection of the oblique shock train was steady with upstream shock interactions. Under backpressure conditions, the rate of movement of the oblique shock train increases rapidly when it passes through the separation regions generated by the ramp, and the deflection direction of the asymmetric separation may switch. Based on the characteristics of the oblique shock train and upstream shock interaction, a flow control method was used to generate asymmetric upstream flow conditions, providing active control of the oblique shock train deflection direction.

Oblique shock train motion based on schlieren image processing

In this paper, shock train motion in a Mach number 2.7 duct is studied experimentally, and large numbers of schlieren images are obtained by a high-speed camera. An image processing method based on Maximum Correlation Detection (MCD) is proposed to detect shock train motion from the schlieren images, based on which the key structures, e.g., separation positions and separation shock angles on the top and bottom walls, can be analysed in detail. The oscillations of the shock train are generated by rhombus and ellipse shafts at various excitation frequencies. According to the analysis of MCD results, the distributions of the frequency components of shock train oscillation generated by the two shafts are distinctly different, in which the motion generated by the ellipse shaft is much smoother; shock train motion is mainly characterized by the oscillation of separation position while the separation shock strength is not so sensitive to downstream disturbance; there is a hysteresis loop relation between the downstream pressure and separation position.

Supersonic combustion diagnostics with dual comb spectroscopy

Supersonic engine development requires accurate and detailed measurements of fluidic and thermodynamic parameters to optimize engine designs and benchmark computational fluid dynamic (CFD) simulations. Here, we demonstrate that dual frequency comb spectroscopy (DCS) with mode-locked frequency combs can provide simultaneous absolute measurements of several flow parameters with low uncertainty across a range of conditions owing to the broadband and ultrastable optical frequency output of the lasers. We perform DCS measurements across a 6800–7200 cm−1 bandwidth covering hundreds of H2O absorption features resolved with a spectral point spacing of 0.0067 cm−1 and point spacing precision of 1.68 × 10−10 cm−1. We demonstrate 2D profiles of velocity, temperature, pressure, water mole fraction, and air mass flux in a ground-test dual-mode ramjet at Wright-Patterson Air Force Base. The narrow angles of the measurement beams offer sufficient spatial resolution to resolve properties across an oblique shock train in the isolator and the thermal throat of the combustor. We determine that the total measurement uncertainties for the various parameters range from 1% for temperature to 9% for water vapor mole fraction, with the absorption database/model that is used to interpret the data typically contributing the most uncertainty (leaving the door open for even lower uncertainty in the future). CFD at the various measurement locations show good agreement, largely falling within the DCS measurement uncertainty for most profiles and parameters.

Comparison between sonic jets with one jet orifice and two opposite orifices into a Ma-3.0 supersonic crossflow through circular ducts

Large-eddy simulations and nanoparticle-based planar laser scattering experiments are conducted to investigate various physical aspects of transverse sonic jets injected into a Ma-3.0 supersonic crossflow through a circular pipe. Configurations with one jet and two opposite jets are compared. For the single jet, a separation shock is generated by the recirculation zone on the opposite wall, and this intersects with the jet shear layer to push several jet plumes into the near-wall region. For the two jets, the bow shocks interact with each other, forming an oblique shock train. All of the shocks promote vortex breakage in jet wakes. A counter-rotating vortex pair is generated in the jet near-field region, enhancing the local mixing. A near-wall region in the jet lee between the counter-rotating vortex pair branches exhibits a low fuel mass fraction. The jet fluid in the downstream near-wall region is entrained by the crossflow upstream of the jet. The interaction between the bow shocks and shear layers of the two jets induces recirculation zones in the lee of the jet, which enhance the fuel mixing. This explains the phenomenon whereby the total pressure recovery coefficient and mixing efficiency of two jets are higher than those of the single jet.

Oscillation of the shock train under synchronous variation of incoming Mach number and backpressure

Experiments were conducted to characterize shock train oscillation under the simultaneous variation of the incoming Mach number and backpressure. Under steady and low-frequency oscillatory backpressure (2 Hz), the incoming Mach number varied from 1.8 to 2.4. According to the intersection of downgoing background wave with bottom front leg, Mach stem, and top front leg of the normal shock train leading edge, the normal shock train/background wave interaction can be divided into three types. Two types of oblique shock train/background wave interaction exist. The downgoing (upgoing) background wave upstream of the oblique shock train can cause the upgoing (downgoing) shock in the shock train leading edge to become the dominated shock. Two modes of shock train oscillation were found: oscillation mode 1, in which the shock train oscillated in the favorable gradient region of the relaxing boundary layer, and oscillation mode 2, where the shock train enters the adverse pressure gradient region caused by the impingement of background wave. Compared with mode 1, mode 2 leads to a larger upstream movement of the shock train and more intense pressure fluctuation. The oscillation of the shock train is caused by instability in the separation region behind the shock train leading edge. The oscillatory backpressure only affected the motion of shock train during each oscillation period. The overall movement trend of shock train is determined by the incoming Mach number and the mean value of backpressure. The increase of incoming Mach number and backpressure can lead to the enhancement of shock train oscillation.

Forced oscillation phenomenon of normal/oblique shock trains in a rectangular duct at Mach 2 and 3

The forced shock train oscillation is a common phenomenon in a practical working scramjet isolator. To evaluate the similarities and differences between the forced normal shock train oscillation and the forced oblique shock train oscillation, experiments were conducted in a rectangular duct at Mach 2 and 3. The normal/oblique shock trains were forced to oscillate for different back-pressure levels by changing the mechanical throttling ratio. High-frequency pressure measurements were mainly utilized for data acquisition and analysis. The frequency contents, the intermittency characteristics and the perturbation propagation of forced shock train oscillations were investigated in this work. Experimental results illustrate that the forced oscillation at Mach 3 is a composite pattern, which is dominated by the excited frequency and coupled with the inherent unsteadiness. This phenomenon is not obvious for the forced oscillation at Mach 2. For the forced oscillations at Mach 2 and 3, the maximum zero-crossing frequency is close to the downstream excitation frequency, and the whole intermittent region almost holds this value, which is very different from the self-excited oscillation pattern. By using the cross-spectral analysis, the frequency-dependent correlation was evaluated. It shows that the excitations and its harmonic contents can propagate upstream to affect the intermittent region and its downstream. Besides the excitation frequency and its harmonic contents, low-frequency contents (tens to more than a hundred Hz) can propagate nearby the separation region, and the propagation to downstream the separation bubble is very weak. Considering the time-lag between different transducers, the propagation speed of perturbations can be obtained. It implies that the propagation speed is independent of the magnitude of the back-pressure, but affected by the velocity of the incoming flow. Specifically, the higher the freestream Mach number is, the greater the propagation speed will be.

Separation induced low-frequency unsteadiness in a supersonic combustor with single-side expansion

Systematic studies on separation induced low-frequency unsteadiness in a canonical supersonic combustor are implemented through wind tunnel experiment and numerical simulation. With an inflow Mach number of 3, cold flow analysis has been carried out to focus on the key impact factor of flow instability. Dynamic flow features are captured by high-frequency pressure signals, and three-dimensional Reynolds-Averaged Navier-Stokes simulation is performed to represent the typical unsteady movement of the shock train. The separated flowfield shows an intrinsic instability, whose feature is the large-amplitude and low-frequency streamwise movement of the oblique shock train. The oscillation of shock train is in a broadband frequency range, and pressure signals obtained from different streamwise regions behave various features. The intermittent region and the backpressure-affected region are two major resources of oscillation energy. Numerical results represent variable-speed shock train motions with multiple amplitudes, and broadband behaviors in experiments are captured. The autocorrelation analysis shows that the broadband behavior of the unsteadiness is not caused by the white noise. From the coherence analysis, it is found that two kinds of oscillation modes (independent and synchronous) exist in the flowfield. The independent mode exists extensively in the unstable flow, while the synchronous mode only appears occasionally and is always suppressed in the very-low-frequency band (below 80 Hz). Repeated experiments indicate that signals from these two oscillation modes superpose randomly. The phase analysis reveals that the backpressure is the original source of this complicated unstable separated flow.

Premixed flame stability and transition to detonation in a supersonic combustor

Simulations of a supersonic, reacting, premixed flow in a channel were performed to investigate the effect of flow speed on ignition, flame stability, and transition to detonation. The configuration studied was a rectangular channel with a supersonic inflow of stoichiometric ethylene–oxygen, a transmissive outflow boundary, and no-slip adiabatic walls. The compressible reactive Navier–Stokes equations were solved by a high-order numerical algorithm on an adapting mesh for inflow Mach numbers, M∞, of 3 to 10. For M∞= 3, the fuel-oxidizer mixture does not reach a sufficient temperature for autoignition. Boundary layers that form on the top and bottom walls deflect the incoming flow, resulting in the formation of an oblique shock train. For M∞ ≥ 5, the fuel-oxidizer mixture ignites in the boundary layers and the flame front expands into the channel. The flame front becomes unstable and turbulent with time due to a Rayleigh–Taylor (RT) instability at the interface between the low-density burned gas and high-density unburned gas. Detonation is initiated in several locations at the flame front and in the unburned gas through an energy-focusing mechanism. As M∞ increases, the time scales for growth of the RT instability at the flame front and eventual detonation increase significantly. Despite the difference in time scales, the flame evolution process is qualitatively independent of M∞: ignition in the boundary layer, laminar flame expansion, growth of an RT instability at the flame front, turbulent flame expansion, and deflagration-to-detonation transition.

Unsteady behavior of oblique shock train and boundary layer interactions

The aim of present investigation is to analyze the unsteady oblique shock train and boundary layer interactions during the self-excited and forced oscillation. The oblique shock train is generated in a Mach 2.7 ducted flow and controlled by a downstream elliptical shaft. Cyclic rotating of the shaft leads to the forced oscillation. A Schlieren system as well as transient pressure measurements and particle image velocimetry have been used to capture quantitative and qualitative shock structure information. Results show that the behaviors of unsteady SBLIs structure are highly related to the dynamics of shock motion. For both self-excited oscillation and forced oscillation, the asymmetrical characteristics of first X-shock was found to be negatively correlated with shock velocity. There exist some relative motions between the first X-shock and the second shock, but the absolute variations are very weak. At lower excitation frequency, the relative motion is not noticeable to the oscillation amplitude, it could be treated as a rigid motion in the duct. At higher excitation frequency, the relative motion amplitude is significant to the oscillation amplitude, and the relative movement of shock cells becomes the dominant motion. There is a hysteretic effect and phase lag between the shock position and downstream pressure perturbation when the shock train travels along a different path for upstream and downstream movements, and the hysteretic effect becomes weaker with increasing frequency.

Experimental Study on Self-Excited and Forced Oscillations of an Oblique Shock Train

Two types of unsteady shock train motion, including self-excited oscillation and forced oscillation, were experimentally studied and compared in this work. The experiments were conducted in a rectangular isolator under Mach 3 incoming conditions. High-speed schlieren technology and high-frequency pressure measurements were adopted to do this research. The mean pressure distributions, the pressure standard deviation distributions, and the spatial frequency distributions were estimated for two types of unsteady shock train oscillations. It was found that the shock train may self-excite oscillations for Mach 3 incoming flow at relatively low frequencies (below 200 Hz) under steady backpressure conditions. With the downstream backpressure forcing, the shock train oscillated in a composite pattern, which was dominated by the imposed frequency content and coupled with the inherently self-excited frequency content. And, the self-excited oscillation was restrained, but was not eliminated, by the forced oscillation.

Response of an oblique shock train to downstream periodic pressure perturbations

An experimental investigation of the response of an oblique shock train to downstream periodic pressure perturbations was conducted. The oblique shock train is generated in a Mach 2.7 ducted flow and controlled by a downstream elliptical shaft. Cyclic rotating of the shaft leads to a periodic oscillatory motion of the oblique shock train. Six cases of perturbation frequency are studied. The results indicate that the downstream pressure perturbations propagate upstream to cause the oblique shock train to oscillate with a translational motion back and forth and the wall pressure to fluctuate with the same frequency. There is no distinct relative motion between the first oblique shock and the second shock during the motion process of the oblique shock train. The entire oblique shock train exhibits its behaviour of rigid motion and the strength of the first oblique shock of the oblique shock train is nearly stable during its periodic motion. There is a clear hysteresis effect in that the oblique shock train travels along a different path for the upstream and downstream motions. A simple analytical model was built based on these experimental data to analyse the oblique shock train dynamics.

Experimental and numerical investigation on hysteresis characteristics and formation mechanism for a variable geometry dual-mode combustor

As part of our efforts to study the hysteresis characteristic and formation mechanism for a variable geometry dual mode combustor, a series of geometry path continuous adjustment experiments and numerical simulation were conducted in the variable geometry dual mode combustor with a Mach number of 3, a divergent ratio ranging from 1.6 to 2.54 and a fuel equivalence ratio ranging from 0.6 to 1.0. Experimental results indicated that the wall static pressure of feature points had an obvious hysteresis phenomenon with the geometry path continuous variation. For given feature points, the hysteresis becomes much smaller with the increasing of the divergent ratio. The interaction between the oblique shock train motion and combustion heat release distribution was adequately considered to explicate the mechanism of the hysteresis formation. It was found that the hysteresis phenomenon was produced from the unstable positive feedback effect of the oblique shock train motion in the isolator. Moreover, the effect of the hysteresis on the total pressure loss, irreversible entropy loss and combustion performance were investigated numerically and experimentally. It was therefore strongly believed that the study of the hysteresis characteristic and formation mechanism on the combustion performance could be very significant to improve combustion performance for the variable geometry dual mode combustor, especially for a wide range of flight Mach numbers.

Experimental characteristics of oblique shock train upstream propagation

The structure and dynamics of an oblique shock train in a duct model are investigated experimentally in a hypersonic wind tunnel. Measurements of the pressure distribution in front of and across the oblique shock train have been taken and the dynamics of upstream propagation of the oblique shock train have been analyzed from the synchronized schlieren imaging with the dynamic pressure measurements. The formation and propagation of the oblique shock train are initiated by the throttling device at the downstream end of the duct model. Multiple reflected shocks, expansion fans and separated flow bubbles exist in the unthrottled flow, causing three adverse-pressure-gradient phases and three favorable-pressure-gradient phases upstream the oblique shock train. The leading edge of the oblique shock train propagates upstream, and translates to be asymmetric with the increase of backpressure. The upstream propagation rate of the oblique shock train increases rapidly when the leading edge of the oblique shock train encounters the separation bubble near the shock reflection point and the adverse-pressure-gradient phase, while the oblique shock train slow movement when the leading edge of the oblique shock train is in the favorable-pressure-gradient phase for unthrottled flow. The asymmetric flow pattern and oscillatory nature of the oblique shock train are observed throughout the whole upstream propagation process.

Computational study and error analysis of an integrated sampling-probe and gas-analyzer for mixing measurements in supersonic flow

Concentration probes are employed in supersonic flow mixing measurements. Because the typical design of such probes is essentially based on an inviscid, adiabatic, quasi-1D analysis, the scope of this work is to understand better and quantify the severe impact of viscous effects on the probe’s internal gasdynamics and the associated uncertainties in the measured quantities via a computational fluid dynamics analysis. Specifically, the focus is on the augmented errors due to the aforementioned viscous effects when coupled with various cases of probe-flow misalignment, which is a typical scenario encountered in mixing measurements of binary gas compositions (air and helium in the present work) in vortex-dominated flows. Results show phenomena such as shock induced boundary layer separation and the formation of an oblique shock train. These flow features are found to noticeably affect the accuracy of the composition measurement. The errors associated with the inviscid, adiabatic, quasi-1D analysis of the probes are quantified in this study.

Mechanism and Prediction for Occurrence of Shock-Train Sharp Forward Movement

A numerical simulation was conducted using the commercial computational fluid dynamics software FLUENT to study the influence of inner flow structures on the motion characteristics of a separation oblique shock train caused by backpressure. Simulation results indicated there was a separation point of the boundary layer at the head of the shock train. The separated flow expanded in the isolator and led to movements of the shock train. With the shock train moving forward in the isolator, the leading edge of the shock train passed through the peaks and valleys of pressure along the forward path. As the pressure gradient at the separation point switched between positive and negative gradients, the state properties of the shock train movements transformed between a slow forward movement and a sudden sharp forward movement. The sudden sharp forward movement of the shock train always occurred when the separation point surmounted reflection points of the background wave system, which implies that the forward motion characteristics of shock train are related to the background flow structures. So, the occurrence of a sharp forward movement of the shock train can be predicted from the flow structures of a fully started state.

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

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