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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. ...

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...

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...

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...

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.

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...

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...

A mixed-compression supersonic inlet with a variable throat can solve the contradiction of the starting capability at low Mach numbers and the total pressure recovery performance at high Mach numbers. However, the variation of the throat area is rather difficult for an axisymmetric supersonic inlet. A new method based on a deforming air cell to control the throat area of an axisymmetric inlet is brought forward and investigated by both experimental and computational methods. The morphing behavior of the deforming air cell in a supersonic wind tunnel is also obtained. The results indicate that the contraction ratio of the inlet can be effectively controlled by the cooperation of the air cell profile under different charging pressures and the cowl inner surface. As compared with a fixed geometry inlet, the total pressure recovery and the corresponding maximum backpressure of the morphing inlet are increased by 25.7 and 31.8%, respectively. In addition, the air cell can withstand the impact of strong oscilla...

The performance of hypersonic inlets is significantly affected by the presence of shock/boundary-layer interactions. To examine the potential of microramps for shock/boundary-layer interaction control in a finite-width duct, a detailed experimental and computational study has been conducted with a separated oblique shock/boundary-layer interaction generated by a 12 deg shock generator at Mach 3.5. Results show that the shock/boundary-layer interaction in the finite-width duct generates complex three-dimensional flow structures with significant swirling nature, and the traditional microramps cannot suppress the separation effectively. Therefore, a type of highly swept microramps with a large chord ratio and small incidence angle is brought forward and investigated. By the precompression effect, the dividing effect, the obstructing effect, and the energizing effect, the highly swept microramps with a height of 0.24 times the boundary-layer thickness show good control capability on the shock/boundary-layer i...

The cowl shock/boundary layer interaction is of great importance to the performance and operability of hypersonic inlets. However, it is usually affected by the expansion waves originated from the convex corner of the ramp surface (often called the shoulder), making the commonly used separation prediction relation unreliable. Therefore, the cowl shock/boundary layer interaction under the interference of the expansion waves is investigated by both theoretical analysis method and computational method in this paper. The results show that the expansion waves bring significant influence on the cowl shock/boundary layer interaction and four kinds of interaction processes exist when their relative position changes. When different interaction processes dominates the flowfield, the expansion waves may bring positive or negative effects on the shock/boundary layer interaction. In particular, while the cowl shock impinges near the shoulder, the interaction processes of shock–shock–expansion wave and shock–expansion ...

Separation due to the shock-wave/boundary-layer interaction in a supersonic/hypersonic inlet always brings negative effects on its performance. A new control method of the shock-wave/boundary-layer interaction based on a two-dimensional bump is brought forward and investigated by both experimental and computational methods. First, the uncontrolled case is studied. The flow turning angle of the incident shock is 12 deg, and the freestream Mach number is 3.5. The results show that the shock-wave/boundary-layer interaction generates complex three-dimensional flow structures with significant swirling nature, which will substantially deteriorate the performance of a supersonic/hypersonic inlet. Then, the controlled case demonstrates that the shock-induced flow separation can be effectively reduced by the coupling of the precompression effect and the acceleration effect of the bump with a height of 0.33 times of the boundary-layer thickness. In addition, the efficiency of the control method for different shock-...

Previous investigations on the submerged inlets show that the counter-rotating vortex pair originated from the side edge is helpful to inhale the external flow into the duct but increases the mixing loss of the inlet. To improve the performance of this type of inlet, a flow-control method based on a bleeding vortex is brought forward in this paper that is expected to use the positive effect of the vortex pair and to reduce the accompanied disadvantages. Computational results indicate that the control method is effective, with the total pressure recovery coefficient increased by 2.8%and the distortion indexdecreased by 51.0%at the design condition as comparedwith the baseline case. Furthermore, the impacts of slot location are also obtained, which underline the fact that it should be chosen carefully according to the transmitting path of the vortex pair. In addition, the drag analysis for the submerged inlet with flow control shows that it is promising for practical use with no extra drag penalty.

To provide information for the detection, prediction, and control of the inlet unstart, the entire process from a started status to an unstarted status of a generic two-dimensional hypersonic inlet is studied experimentally at Mach 5. The movement of a flow plug at the exit of the duct is used to gradually increase the throttling to simulate the unstart process caused by the excessive heat release in the combustor. Simultaneous high-speed schlieren imaging and dynamic surface pressure measurements are used to record the unsteady flow structures and surface pressures of the unstart process. According to the internal and external flows, the unstart process can be divided into four stages, namely, shock train in the combustor, shock train in the isolator, separation bubble in the throat, and unstart. The transient flow patterns of each stage are substantially different and the corresponding dynamics pressures also have prominent time-frequency features, which make the detection and prediction of the inlet unstart based on dynamic pressures possible. Since the sensor placed at the end of the top surface of the inlet contract part, marked by C1, obtains the most abundant time-frequency characteristics which can be used to discern the different stages of the inlet unstart process, the location where it stays is regarded as the first choice for sensor placement to construct a practical inlet-status-monitoring system.

To realize the control of the ramp shock system of hypersonic inlets, the current study involves a fluidic shock control technique featured by fixed geometry. First, a theoretical analysis and design method for this technique is developed, including the calculation of the shock control requirement of a typical hypersonic variable inlet, the analysis of the shock control capability of a set of S-shaped surfaces attached to a ramp surface, the estimation of the required secondary-to-primary mass flow ratio, and the realization of an aerodynamic curved surface by distributed air injections. Then, the correctness of the method and the underlying flow mechanism are verified by computational fluid dynamics with a two-dimensional case. Finally, a three-dimensional test model is designed and investigated by experiments and computations to demonstrate the capability of the fluidic shock control technique, examine the three-dimensional effects, and verify the computational fluid dynamics tool. The results show that the shock control capability of this technique can meet the requirement of hypersonic variable inlets operating from Mach 4 to 6. At a secondary-to-primary mass flow ratio of 1 % or less, the first ramp shock can be consistently maintained on the cowl lip from Mach 5 to 6.

Wind-tunnel tests are conducted to study the characteristics of the shock train in the curved variable-section diffuser for hypersonic inlets. The test model is equipped with a forebody, a contracting entrance, a dump mixing duct, and an aft plug. Tests are performed at nominal freestream Mach numbers of 4.0, 5.0, and 6.0, and the corresponding inlet Mach numbers of the diffuser are 2.05, 2.59, and 3.06, respectively. Results indicate that at an inlet Mach number of 2.05, the surface pressure distributions in the shock train are similar at different backpressure ratios and can be well predicted by the modified Waltrup formula, but the length of the shock train is increased by 32 % due to the curved duct, the incident shock waves, and the incident expansion waves at the inlet plane. At higher inlet Mach numbers, the similarity disappears and the distributions of the surface pressure are not easy to predict. At different measurement points in the shock train, the instantaneous surface pressures vary obviously and almost synchronously, suggesting the oscillatory motions of the shock train. At an inlet Mach number of 2.59, the base frequency of the oscillations is about 19 Hz and the maximum fluctuating range of the instantaneous surface pressure is as high as 21.6% of the ideal pressure rise of the shock train. The static pressure fluctuations at different points in the shock train correlate strongly but are almost unrelated with those of the exit survey point, which indicates that the oscillation of the shock train is not likely induced by the pressure fluctuations far downstream of the shock train.

Wind-tunnel tests were conducted to better understand the flow pattern of a hypersonic inlet at an unstable-unstarted condition that is caused by the downstream mass flow choking. High-speed shadowgraph flow visualization and instantaneous pressure measurements are employed to reveal the oscillations of the shock system and the unsteady process of the duct flow. Results show that the shock system both inside and in front of the inlet duct oscillates intermittently when the inlet is unstarted. With the increase of the downstream throttling ratio, the fundamental frequency of the unsteady flow process goes up and the shock-on-lip time goes down, but the process whereby the shock system is disgorged and swallowed remains unchanged, lasting for 15 ms. When the throttling ratio is 89 and 91%, the fundamental frequency is 8 and 23 Hz, respectively. At the smaller throttling ratio, the aerodynamic oscillatory phenomenon of the inlet can be divided into three stages: namely, mass filling up, shock system disgorging and swallowing, and near-throat flow pattern establishing and backpressure propagating. At the larger throttling ratio, the first stage disappears and the third one shortens.

Combined with a UAV of the shape like Global Hawk, a new inlet is advanced to obtain high performance in both Radar Cross Section(RCS) and aerodynamic drag. Efforts are made to achieve this goal such as adopting a top-mounted inlet configuration, utilizing the diverterless technique and putting forward a new shape of entrance. A design method is brought forward and verified by wind tunnel tests. Results indicate: (1) Despite the negative effect of the front fuselage and the absence of the conventional boundary diverter, the performance of the top mounted diverterless inlet advanced here (Ma: 0.50-0.70, α: − 4°-6°, σ > 0.975) is equivalent to that of conventional S shaped inlet with diverter; (2) The integration of the inlet with the fuselage is realized by the utilization of a special inlet section and the diverterless technique, which disposes the whole inlet in the shield of the head of UAV, improving the drag characteristics and the stealthy performance of the aircraft; (3) The bump which is equal to the local boundary layer thickness in height can divert the boundary layer effectively. As a result, no obvious low total pressure zone is found at the outlet of the inlet; (4) According to the experimental results, negative angle of attack is favorable to the total pressure recovery and positive angle of attack is favorable to the total pressure distortion, while yaw brings bad effects on both; (5) The design of cowl lip is of great importance to the inlet performance at yaw, therefore, further improvement of the inlet performance will rely on the lip shapes of the cowl chosen.