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A generic external-compression inlet model was tested to investigate the effect of duct height on terminal shock/boundary-layer interactions, with special sidewall arrangements to eliminate corner separations. Experimental results showed that the separation length significantly decreased by 41% as the duct height was reduced from 60 mm [Formula: see text] to 40 mm [Formula: see text]. To obtain the influence of the duct height in a wider range of values, Reynolds-averaged Navier–Stokes simulations were further performed. The scales of the [Formula: see text]-shock structure and separation region decrease as the duct height decreases. Shrinkage of the scale of the λ-shock structure results from a decrease in separation length. The downstream shift in the separation point and the upstream shift in the reattachment point both contribute to this decrease, but the upstream shift in the reattachment point is a dominant reason. The underlying mechanisms of the variation in the reattachment position as the duct height decreases include two aspects. The first is the stronger pressure gradient in the wall-normal direction immediately behind the shock system, which produces a stronger radial pressure gradient above the bumplike separation bubble, quickly turning the streamlines toward the bottom wall and enclosing the bubble. The second results from the postshock acceleration of the irrotational flow above the separated layer as the duct height decreases. An accelerating mainstream with a favorable pressure gradient above the separated layer increases the mixing strength and momentum exchange between the low-speed layer near the wall and high-speed flow above the separated layer.

A generic external-compression inlet test model is designed and fabricated to investigate the flow characteristics of terminal shock/boundary-layer interaction for inlet applications with and without upstream lateral confinement. When the leading edge of the sidewall is positioned 150 mm upstream of the duct entrance, the bifurcation height of the [Formula: see text] shock, the centerline separation length, and the pressure rise ratio at [Formula: see text] are close to reported values in the literature, although the corner separation herein is minor as compared with previous experiments with thick sidewall boundary layers. However, when the leading edge of the sidewall is aligned with the duct entrance, an obvious increase in the bifurcation point height can be observed, and an increase in centerline separation can be inferred from schlieren photographs. The significantly intensified separation is verified directly by surface oil flow visualization. In addition, the focus previously present in the corner region is replaced by a reattachment node located near the junction region of the bottom plate and sidewall leading edge. The intensification of the separation is attributed to a local high-pressure region caused by the impingement of the postshock subsonic flow onto the sidewalls.

The recognition of supersonic inlet flow pattern has become a research hotspot in recent years. In this paper, the dual external pressure supersonic inlet is taken as the research object. To explore the flow characteristics of the inlet, time-mean processing on the inlet pressure signal collected by sensors is conducted first, and the features of the inlet pressure data in time domain and frequency domain are extracted, respectively. As feature selection (FS) plays an important role in classification tasks and has been recently studied as a multi-objective optimization problem, two objectives of FS are considered and an improved non-dominated sorting genetic algorithm NSGA2 with hybrid mutation operators using support vector machines (SVM) as classifiers is proposed, aiming to simultaneously select feature subsets and optimize SVMs hyper-parameters. In addition, a way to deal with variation transgression is proposed to make the mutation operator of the single-objective evolution fit well in the multi-objective evolution algorithm. Experimental results on 31 sensor datasets demonstrate that our proposed algorithm can achieve competitive classification accuracy while obtaining a smaller size of feature subset compared with particle swarm optimization algorithm and some multi-objective optimization algorithms using single-objective evolution mutation operators.

The variable geometry supersonic inlet tends to decrease the throat area to reduce the Mach number upstream of the terminal shock, so as to reduce the flow loss. However, excessive Internal Contraction Ratio (ICR) exposes the inlet to a greater risk of unstart, which inevitably results in a process of increasing the throat area to aid the inlet restart. In the above throat regulation process, the inlet undergoes the start, unstart, and restart states in turn. In order to reveal the flow structure and mechanism of this process, a two-dimensional unsteady numerical simulation combined with a dynamic mesh technique were employed. The shock-on-lip Mach number of the studied inlet is 4.0 and the flight angle of attack is + 6°. Analysis was focused on the state with a freestream Mach number of 3.0. The results clearly show that the flow response hysteresis appears, and restart is only realized when the throat area is obviously increased as compared to that of unstart due to the historical unstart flow structure. In addition, three typical flow fields were analyzed, and it is found that the separation ahead of the inlet was the key factor affecting the hysteresis. Finally, unstart and restart boundaries of the inlet were discussed, and the factors influencing its deviation from the typical boundaries of dual-solution area were analyzed. The newly predicted unstart and restart boundaries are much closer to the CFD results.

In this paper, the restart processes of two generic rectangular hypersonic inlets with different internal contraction ratios have been experimentally studied under a freestream Mach number of 5.0. ...

The submerged inlet is an attractive configuration for advanced helicopters due to its high stealth performance and low external drag. In this paper, a submerged inlet, integrated with a ROBIN helicopter fuselage and a simplified power output shaft, is experimentally and numerically investigated to obtain the basic flow characteristics under a freestream velocity of 23.6 m/s. The results indicate that the pylon ahead of the inlet induces a horseshoe vortex. Though the vortex is ingested into the inlet, it has little effect on the internal flows and can be neglected. When the airflow enters into the inlet, it interacts with the shaft with a large incidence angle, yielding a vortex pair. At the leeside of the shaft, the two side flows of the shaft impinge at the center plane, generating a local high-pressure region at the azimuthal angle of 180°, which forces the boundary layer to roll up a counter-rotating vortex pair. In addition, the airflow adjacent to the cowl lip accelerates rapidly, resulting in a local low-pressure region at the azimuthal angle of 0°. Therefore, the inlet duct has a strong circumferential pressure gradient, which originates from an azimuthal angle of 180° to 0° and induces a vortex pair at the azimuthal angle of 0°. The three vortex pairs are the main origins of the distortion at the duct exit plane, among which the one near the cowl lip with the azimuthal angle of 0° plays the dominant role. Additionally, as the velocity ratio increases from 3.9 to 5.5, the circumferential pressure gradient and the cowl lip vortex get intensified, which causes that the total-pressure recovery coefficient drops by 0.5% and the distortion index increases by 28%. A submerged inlet, integrated with a ROBIN helicopter fuselage and a simplified power output shaft, is experimentally and numerically investigated. Three vortex pairs, which locate at the azimuthal angle of 0°, the leeside of the shaft, and 180° of the inlet surface, are the main origins of the distortion of the inlet, among which the one near the cowl lip with the azimuthal angle of 0° plays the dominant role. As the velocity ratio increases, the circumferential pressure gradient gets intensified, leading to stronger vortex pairs.

Herein, a curved isolator with sidewall contraction of a hypersonic scramjet is extracted and simplified. The flow characteristics in the isolator with an entrance Mach number of 3.46 are studied. The results show that the lateral contraction of the sidewall generates a strong shock and its subsequent reflected shock, which induce glancing shock wave/boundary layer interaction (GSWBLI), rolling up large-scale vortices adjacent to the symmetry plane. As the downstream backpressure increases, the shock train propagates upstream and interacts with the glancing shock and vortices inevitably, forming a typical shock train/glancing shock/boundary layer interaction (STGSBLI) phenomenon. Different from the shock train in a straight isolator, it can be divided into two distinct parts, i.e., the center part and the side part. For the center part, it behaves in a quasi-two-dimensional manner, which is primarily associated with STGSBLI, forcing the low-momentum subsonic flow to principally accumulate near the symmetry plane and weakening the spanwise pressure gradient in the center part. For the side part, it locates behind the glancing and reflected shocks and is much shorter than the center part, resulting in a streamwise extension of the supersonic flow region near the sidewall. As the shock train moves upstream, the quasi-two-dimensional region enlarges and extends to the sidewall. It is demonstrated that the aforementioned two parts of the shock train can also be discovered in curved sidewall-contraction isolators with different centerlines. Therefore, the STGSBLI is the dominant and universal physical phenomenon in isolators with sidewall contraction.

Görtler vortices (GVs) in dual-incident shock-wave/turbulent-boundary-layer interactions (dual-ISWTBLIs) are experimentally investigated in a Mach 2.48 flow. A double-wedge shock generator with two deflection angles of 8° and 5° is used to produce two incident shock waves (ISWs). Flow structures of the experiments with three different shock-wave distances were visualized by the ice-cluster-based planar laser scattering technique at two orthogonal planes ( x– y and x– z planes). The images in the x– y plane present three types of flow patterns of dual-ISWTBLIs corresponding to the first type with a triangle-like separation, the second type with a quadrilateral-like separation, and the third type with two isolated interactions induced by the two ISWs. The images in the x– z plane indicate that the GVs exist in the first type of dual-ISWTBLI originating in the vicinity of the apex of the separation region and cover nearly the whole spanwise range of the reattachment region. By comparison, the GVs intermittently occur in the limited spanwise range of the reattachment region in the second type of dual-ISWTBLI. No GVs are observed in the third type of dual-ISWTBLI because no visible separation is induced under the experimental conditions considered in this situation. In addition, based on the wall-pressure distribution in the former two types of dual-ISWTBLIs, this paper proposes a method to estimate the mean-flow streamline curvature in the reattachment region, thereby obtaining the criteria for the existence of GVs, according to which reasonable explanations for the different distributions of GVs in the two types of dual-ISWTBLIs are provided.

As the basis of protection control, supersonic inlet plays an important role in a supersonic air-breathing propulsion system. To reduce the negative effects of buzz flow on the supersonic i...

To greatly improve the supersonic inlet stability at low cost of structural weight and complexity, a novel buzz suppression strategy based on fixed-geometry air bleed is developed. It is designed t...

A deformable shape-memory alloy (SMA) bump is introduced to control the succeeding cowl-shock/boundary-layer interactions in a supersonic inlet with an operating Mach range of 2.0–3.8. The deformat...

A new fluidic method to control the hypersonic inlet external shock system based on local secondary flow recirculation and the related prediction method for the profile of the restructured shock ar...

The plasma synthetic jet is a new active flow control technique, which has great potentials for supersonic flow control A plasma synthetic jet actuator (PSJA) for supersonic flow control which operates under low ambient pressure is designed In order to explore the transient jet flowfield, high-speed Schlieren and electronic measurement systems are utilized to test the single-shot operating characteristics of the actuator under two ambient pressure conditions. The evolution of the jet boundary, which depicts the transient jet flowfield and can be used to estimate the flow control capability, is captured. It is found that the ambient pressure is the primary reason to affect the arc energy deposition, which directly determines the velocity of blast wave and jet front. In addition, the flow patterns of PSJA under various ambient pressures show some similarities. The jet evolution can be divided into three stages, i.e. pressure dominant stage, the inertia dominant stage and vortex ring dominant stage. During the pressure dominant stage, the PSJA could be treated as jet. For the inertia dominant stage, the Froude number decreases from 2263.82 to 21.73, indicating the inertia effect gets weakened with an intensified buoyancy effect, but the inertia effect still holds the dominant position, and this stage may be considered as the end mark of the injection process. As for the vortex ring dominant stage, large scale vortex ring becomes significant and the jet front velocity is very low but with apparent fluctuation.

To study the duct curvature effects on a shock train in rectangular hypersonic inlet/isolator models, four curved isolators are designed and tested at a freestream Mach number of 4.92. Strong inter...

A rectangular external-compression supersonic inlet with partially isentropic compression is experimentally studied at a freestream Mach number of 2.0. During the test, the inlet operates at the li...

A supersonic inlet is one of the key components in a supersonic air-breathing propulsion system and is the basis for protection control. The overall system performance can be greatly influenced by its flow patterns, so it plays a crucial part and is necessary to develop methods for monitoring its flow patterns to ensure stable and safe operation. This issue can be viewed as a time series classification (TSC) task. Traditionally, several manually-engineered features are extracted as the indicators to evaluate the operation status, but this process can be heavily dependent on the professional experience. In this paper, a novel neural network called DTW-SLFN-KF is proposed, which integrates Dynamic Time Warping (DTW) and Kalman Filter (KF) into a single-hidden-layer neural network (SLFN) architecture to directly determine flow patterns from the dynamic sensor signals. The proposed network first adopts a DTW layer as the feature extractor to automatically extract robust features, and exploits the flexible alignment ability of DTW to keep the temporal continuity and deal with the temporal distortions. Then, these features are fed into an SLFN for classification. After that, to make full use of the extracted features and improve the classification performance of SLFN when the network structure is fixed, KF is applied as a linear post-processing technique to get the predicted output of SLFN closer to the true output. Experimental results demonstrate that the proposed DTW-SLFN-KF network has better comprehensive performance for monitoring the flow patterns of supersonic inlet in terms of monitoring accuracy and real-time performance when compared with other competitive methods.

The unthrottled flows in curved isolators are studied experimentally and numerically. Wind-tunnel tests are conducted at a freestream Mach number of 4.92. The internal flowfield is visualized and m...

A fluidically variable hypersonic inlet with fixed geometry is investigated in this paper by both experimental and numerical methods. The results show that the inlet maintains the shock-on-lip cond...

A rectangular hypersonic inlet is tested at the low-speed mode (Ma 5), the design mode (Ma 6), and the overspeed mode (Ma 7) to enrich the understanding of hypersonic-inlet unstart. A flow plug is placed at the duct exit to simulate the combustion-induced high pressure and initiate the inlet unstart. High-speed schlieren imaging and time-resolved pressure measurements are used simultaneously to record the unsteady-flow structures and surface pressures of the unstart process. The results indicate that the unstart processes are basically the same for the low-speed mode and the design mode. An unstart phenomenon dominated by the ramp-side flow separation, which exhibits violent fluctuations along the entire duct with a low frequency, is observed at these two operating modes. While operating at the overspeed mode, the strong shear layer, induced by the Mach reflection of the coalesced external oblique shock, divides the internal flow into two parts. The cowl side is filled by fully subsonic flow after the com...

Fractal features of the turbulent/non-turbulent interface (TNTI) in shock wave/turbulent boundary-layer interaction (SWBLI) flows are essential in understanding the physics of the SWBLI and the supersonic turbulent boundary layer, yet have received almost no attention previously. Accordingly, this study utilises a high spatiotemporal resolution visualisation technique, ice-cluster-based planar laser scattering (IC-PLS), to acquire the TNTI downstream of the reattachment in a SWBLI flow. Evolution of the fractal features of the TNTI in this SWBLI flow is analysed by comparing the parameters of the TNTI acquired in this study with those from a previous result (Zhuang et al.J. Fluid Mech., vol. 843, 2018a).

Sodium taurocholate cotransporting polypeptide (NTCP) is a carrier protein encoded by the human SLC10A1 gene, acting as the principal transporter of conjugated bile salts from the plasma into hepatocytes. Although NTCP was cloned as early as in 1994 and its function has been studied extensively, clinical description of NTCP deficiency remains rather limited thus far. The patients in this paper were 2 female monozygotic twins, who were referred to our hospital at the age 2 years with the complaint of persistently-raised total bile acids (TBA) for 21 months. At age 3 months, they were both diagnosed to have cholestatic liver disease due to raised serum TBA and direct bilirubin (DBIL) with the fraction >20% of the elevated total bilirubin (TBIL). Thereafter, their jaundice subsided and the DBIL levels recovered gradually, while serum TBA remained raised persistently. In view of their refractory hypercholanemia but negative symptoms and signs, SLC10A1 genetic analysis was performed for all family members to evaluate the possibility of NTCP deficiency. As a result, the twins were both homozygotes, while the parents, carriers, of the reportedly pathogenic variant c.800C>T (p.Ser267Phe). These findings suggested that NTCP deficiency may be a unique genetic factor causing transient cholestasis in early infancy, as well as, persistent hypercholanemia in pediatric patients.

As the basis of protection control, supersonic inlet plays an important role in a supersonic air-breathing propulsion system, so it is of great significance to ensure the safe and stable operation by monitoring its flow patterns. From the perspective of machine learning, this issue can be viewed as a time series classification (TSC) task. Traditionally, several manually-engineered features are extracted as the indicators to evaluate the operation status, which can be heavily dependent on the professional experience and time-consuming. This paper proposes a novel Dynamic Time Warping-Radius Basis Function (DTW-RBF) network to directly determine the flow patterns from the dynamic input signals. DTW-RBF network replaces the Euclidean distance in the static RBF kernels with the DTW distance, which exploits the elastic matching ability of DTW to align the input signals to the kernels. Then, the second-order Levenberg-Marquarelt (LM) optimization algorithm is used to allow the efficient training process of the proposed network. In order to determine the appropriate locations for sensor placement and enhance the robustness and reliability of monitoring flow patterns by a single sensor, an optimal subset of sensors is further selected for ensemble through multi-objective optimization and fuzzy decision. Experimental results demonstrate that the proposed DTW-RBF network works efficiently for TSC tasks on benchmark time series datasets, and has better comprehensive performance for monitoring supersonic inlet flow patterns in terms of classification accuracy and test time. The ensemble classifier further increases the classification accuracy with still meeting the real-time requirements.

Placing the right amount of sensors in key locations is critical for system monitoring. In real applications, the determination of sensor placement is a compromise between monitoring performance and the costs of installation and maintenance. Given the well-interpretability of sparse learning, this paper proposes an efficient data-driven method to obtain the optimal sensor subset from the entire candidate sensor set. In order to make our model more robust to outliers and overcome the limitation of inconsistent coefficients for multiple class optimization problem, our proposed method introduces a special norm to realize the similar sparse structures of coefficients. Considering that the redundant data cannot effectively improve the real-time condition monitoring performance of engineering systems, our proposed method also includes a redundant information elimination model, which is rarely investigated in data-driven methods for optimal sensor placement problem, and this elimination model is designed by exploring the diversity of measurement data of different sensors. What’s more, we provide an alternating iteration algorithm to solve the non-smoothness convex problem of our proposed data-driven method, and the proof of its convergence has also been presented. The optimal sensor subset can be determined by the rank of the coefficients obtained by the alternating iteration algorithm. Finally, the effectiveness and feasibility of our proposed method are verified by a large number of experiments, including validation experiments on benchmark data sets and a real engineering example on the inlet model of hypersonic aircraft engine.

A hypersonic inlet with side compression has been tested at a freestream Mach number of 6.0 to enrich the understandings of hypersonic inlet unstart. A flow plug is placed at the duct exit to simulate the combustion induced high pressure and to initiate the inlet unstart. High-speed schlieren imaging and time-resolved pressure measurements are used simultaneously to record the unsteady flow structures and surface pressures of the unstart process. The inlet operates in a big buzz mode with a base frequency of 30 Hz and exhibits a series of unsteady flow patterns similar to those of rectangular hypersonic inlets in a buzz cycle when the throttling ratio is 87.4%. During the upstream propagating process of the unstart shock system, the propagation velocity in the two ends of the duct is higher than that in the middle section of the duct, with a minimum value around the isolator. Once the separation bubble induced oblique shock is expelled over the cowl lip, a supersonic reverse flow with a Mach number of 1.5...

A new plasma synthetic jet actuator with low exit inclination angle is proposed. Transient flow features under an ambient pressure of 13 kPa are experimentally tested. It's found that the plasma-induced flow can be classified into two stages, namely, initial shock diffraction dominated stage and the stratified flow stage. Regarding the shock diffraction dominated stage, a curved primary shock following a rotational jet will be formed; during the latter stage, it's dominated by the buoyancy and inertial effects jointly, leading to an upward and forward moving jet. The initial jet can accelerate to a maximum speed of ∼300 m/s, and the high speed of the jet can maintain for about 60-100 μs with its trajectories during the whole lifetime showing a piecewise linear feature. The repeated working of the actuator is examined as well, and the results demonstrate that the actuator can operate stably with the frequency ranging from 100-2 kHz. Although the increase of operating frequency lowers the moving speed of the primary shock and the jet to some extent, it retards the duration of the high-speed jet. Additionally, the effect of deposited energy per cycle on the flow characteristics of the actuator is studied. As the deposited energy decreases, the primary shock and jet move more slowly. In conclusion, the deposited energy per cycle of 82 mJ and the operating frequency of 1 kHz can induce a relatively high-speed jet with a longer duration.

To further explore the shear-layer-induced buzz diversity recently discovered in an overspeed supersonic inlet, an external-compression supersonic inlet is specially designed and carefully studied through wind-tunnel test at its design Mach number of 2.0. It is observed that the flow instability begins with the little buzz mode featuring a stable terminal shock and a lightly unsteady shock-induced separation bubble. Then a combination of another two stronger buzz styles replaces the little buzz. Of them the medium buzz is characterized by a locally destabilized terminal shock and the varying cowl-side reverse flow, whereas the big buzz is distinguished by intense separation unsteadiness. Interestingly, the medium buzz cannot stay long and vanishes eventually. Instead of an ordinary buzz case in the design Mach number situation, current buzz flows are found pretty similar to the overspeed inlet case superficially and fundamentally under the influence of two strong shear layers, albeit with a clear difference in the duration of the medium buzz. Accordingly, the possibility of buzz flow diversity and the related shear-layer effect proposed recently get further supported. Besides, it is indicated that the overspeed operation is not the only situation allowing for the diversified inlet buzz.

The turbulent–non-turbulent interface (TNTI) of supersonic turbulent boundary layers is a fundamental but relatively unexplored physics problem. In this study, we present experimental results from fractal analysis on the TNTI of supersonic turbulent boundary layers, and test the applicability of the additive law for these flows. By applying the nanoparticle-tracer planar laser scattering (NPLS) technique in a supersonic wind tunnel, we obtain data covering nearly three decades in scale. The box-counting results indicate that the TNTI of supersonic turbulent boundary layers is a self-similar fractal with a fractal dimension of 2.31. By comparing data sets acquired from two orthogonal planes, we find that the scaling exponent does not depend on direction, consistent with the validity of the additive law for the TNTI of turbulent boundary layers in a scale range with the large-scale limit not exceeding approximately $0.05\unicode[STIX]{x1D6FF}$.

An ultracompact serpentine inlet, which contains a compact diffuser as short as 2.3 times the exit diameter, is experimentally and numerically studied. An internal bump is introduced to red...

Under the influence of duct curvature, cross-sectional area variation and internal struts, the internal flow field within a curved annular duct becomes rather complicated and contains strong secondary flow. In this paper, the secondary flow characteristics in an annular duct with struts are experimentally and numerically investigated. The results show that large pressure gradients exist on the bends of hub and shroud. Meanwhile, two counter-rotating vortex pairs appear both along the hub-side and shroud-side surfaces. The hub-side vortex pair of which the vortex cores travel downstream parallelly evolves from the horseshoe vortex which is induced by the leading edge of the upstream strut, whereas the shroud-side vortex pair originates from the strut trailing edge and the corresponding vortex cores develop in a divergent way. Additionally, the effects of the duct exit Mach number on the secondary flow characteristics are also studied. As the exit Mach number increases, the streamwise pressure gradients increase and lead to more intense vortices, higher total pressure loss and larger flow distortion.

The influence of the secondary flow injection angle on a fluidic shock control technique that is used to control the ramp shock of supersonic and hypersonic inlets is studied in this paper. The differential equation of the momentum thickness of the boundary layer, the injection ratio, and the secondary flow injection angle is first deduced. Then, the relation of the secondary flow injection ratio with the displacement of the first ramp shock is established. Based on this, a theoretical prediction method of the required secondary mass flow ratio for a given shock control requirement is developed in an iterative manner and then validated by both numerical simulations and wind-tunnel tests. The theoretical analysis shows that the momentum thickness development rate of the boundary layer is closely related to the dimensionless tangential velocity of the secondary flow and the injection ratio. Although the desired displacement of the first ramp shock remains the same, the required secondary mass flow ratio dec...