Severe flow oscillation induced by low-mass-flow of inlet directly threatens flight safety. The unsteady simulation on the low-mass-flow induced instability phenomenon of the Bump inlet under over-speed operating conditions at Ma₀=1.8 is conducted. The effects of different throttling levels on the inlet is studied. The results show that when the inlet mass flow is reduced to 48% of the through-flow mass flow rate, the inlet enters a periodic surge state. Based on the flow characteristics, the instability process can be divided into three typical stages, i.e., terminal shock expelling out of duct, backflow development, and the internal flow re-establishment. Driven by the high pressure inside the tube, the terminal shock propagates forward, with its movement speed initially decreasing and then increasing. As the terminal shock merges with the lip shock, it triggers large-scale flow separation. The interaction between this separated flow and the incoming freestream generates unsteady vortical structures, which disrupt the inlet shock system. In the later stage of the backflow development phase, significant reverse flow is observed near the inlet, and a zero-velocity interface forms in the vicinity of the outlet section. Subsequently, the pressure inside the duct gradually decreases, allowing the internal flow field to be reestablished. As the second throat diameter decreases (corresponding to a further reduction in mass flow), the maximum pressure ratio inside the duct increases from 6.61 to 6.84, and the normalized backflow propagation distance ratio increases from 2.43 to 3.33.

To induce the near-wall low-momentum fluids moving laterally, an aerodynamic design method of hypersonic forebody/compression surface with the controlled lateral surface pressure is proposed. The basic principle is that the lateral pressure distribution on one section after the external conical flow field is prescribed, and the spatial coordinates of the section can be derived inversely through the coordinate transform, then the forebody/compression surface can be obtained by the stream tracing method. The numerical results demonstrate that the lateral pressure gradient dominates the motion of the near-wall low-momentum fluids on the forebody/compression surface. For the conventional forebody/compression surface, it has strong lateral pressure gradient on the 1^st stage of the forebody, which can induce a lateral flow with a deflection angle below 3° at the design point (Ma = 7.0 and H = 28 km). However, it almost has no lateral pressure gradient on the subsequent compression surfaces, and the lateral flow is also fairly weak. The controlled lateral pressure distribution forebody can intensify the lateral pressure gradient about 7 times within the sector-angle ranging from 0° to 40°, the deflection angle increases about 5° on the 1^st stage of the forebody, and the lateral pressure gradient increases significantly with the deflection angle increased over 7° on the 2^st and 3^rd stages of the forebody. Consequently, the boundary layer thickness decreases about 20%, and the total-pressure recovery coefficient in the sector region of the inlet increases about 1.56%.

A flow control method using multiple micro-vane vortex generators is proposed to improve the fullness of boundary layer profiles over the forebody ramp of hypersonic vehicles and to reduce the potential risk of flow separation in the inlet. The flowfield characteristics and mixing mechanism at an incoming Mach number of 7.0 are numerically analyzed. The effect of the installation angle of vortex generators on the flowfield is studied as well. Results show that micro-vane vortex generators can generate a local large-sideslip-angle and low-pressure flow in the near-wall region. At the lateral sides of mocro vanes, glancing shocks and expansion waves are generated, which induce the lateral mixing of low-momentum fluid with the mainstream. When installation angles are positive, with the increase of installation angle, the mixing effect gets stronger but the pressure loss increases. Micro vanes with negative installation angles lead to the strongest mixing effect, however, they also bring significant total pressure loss. Compared with the one without flow control, the shape factor of the boundary layer gets smaller by using vortex generators, among which those with an installation angle of 15 degrees lead to the smallest shape factor and consequently the fullest boundary layer profile, which has better anti-separation capability under reverse pressure gradients.

Metal-organic frameworks (MOFs) materials exhibit inherent advantages in microwave absorption (MA) due to their unique compositional and structural characteristics. However, it remains a significant challenge to systematically elucidate the coordinated effects of component modulation and morphology design on electromagnetic parameters during the complex modification process to enhance the MA performance of MOFs. In this study, the two-dimensional (2D) Co-based MOFs material, layered zeolitic imidazolate framework (ZIF-L), was meticulously selected as a structure design template and combined with an efficient ion-exchange component modulation method, aiming to stimulate the potential advantages between the two in the synergistic modulation of electromagnetic parameters. Specifically, the ion-exchange method enhanced various polarization mechanisms, including interface, dipole, and defect-induced polarizations. Meanwhile, the 2D template facilitates the formation of a robust conductive network on the surface and ensures the high efficiency and uniformity of the ion-exchange process. Ultimately, the resulting material achieved an adequate absorption bandwidth (EAB) of 6.47 GHz at a thickness of 2 mm, a 78% improvement over the untreated sample (EAB = 3.64 GHz). This research not only provides valuable theoretical insights into the application of 2D MOFs materials in MA but also offers innovative perspectives for the design and development of energy-saving and environmentally friendly microwave absorption materials (MAMs).

To ensure the efficient matching of the supersonic Bump inlet with the engine across the entire flight envelope, this paper explores the dominant mechanism for enhancing its stable operational margin. Taking the four-lip forward-swept cowl Bump inlet as the subject of study, the evolution of the three-dimensional flow structure from supercritical to subcritical conditions at the designed Mach number is analyzed. The research findings indicate that under low mass flow (or high back pressure) conditions, the interaction between the normal shock and the boundary layer on the compression surface of the inlet, which combines a forward-swept cowl and a conical bump, generates a three-dimensional separation vortex that is expelled to the exterior of the inlet entrance. In contrast, an inlet with the same forward-swept cowl combined with a flat wedge experiences a “quasi-two-dimensional” flow separation under low mass flow conditions, with most of the separated flow being ingested into the inlet to result in a narrower stable margin as compared with the Bump inlet. Therefore, the generation of a three-dimensional separation vortex and its expulsion out of the entrance of the duct is the core mechanism for enhancing the aerodynamic performance of the Bump inlet and broadening its stable operational margin.