Opposing jet flow control technology has a broad application prospect and can achieve effective drag reduction control under the long penetration mode (LPM). In order to further study the opposing jet mode transition mechanism and the influence of freestream conditions on drag reduction control of hypersonic vehicles, wind tunnel tests were conducted on the blunt body model under different freestream total pressures (0.1 and 0.4 MPa), Mach numbers (Ma = 5, 6) and jet total pressure ratios. The structure of the spatial flow field and the surface pressure distribution of the model were obtained by using high-speed Schlieren and electronic pressure scanning valve. The results show that the LPM is extremely unstable, with strong unsteadiness and asymmetry, and the flow field presents axial and radial oscillations. At the transition phase, the flow field oscillates axially and changes continuously between LPM and short penetration mode (SPM). The SPM is stable and the flow field structure is axisymmetric. As the total pressure ratio increases, the surface pressure distribution of the model corresponds to the phase of jet mode transition. The critical total pressure ratio of the jet is affected by the freestream total pressure and Mach number. As the freestream total pressure increases, the critical total pressure ratio increases, and the total pressure ratio interval corresponding to LPM becomes larger. With the increase of incoming Mach number, the minimum total pressure ratio and critical total pressure ratio of sonic jet decreases, and the total pressure ratio interval corresponding to LPM becomes smaller.

Wind tunnel testing serves as a critical method for studying the performance of high-speed aircraft. Enhancing the measurement accuracy of infrared characteristics in wind tunnel tests can provide more reliable data support for predicting aircraft infrared signatures and designing stealth performance, embodying significant engineering value. However, in wind tunnel tests, the interference of background radiation inside the wind tunnel makes it difficult to accurately obtain the true infrared radiation (IR) of the aircraft, directly affecting the execution of wind tunnel tests and the use of test data. The CFD (computational fluid dynamics) method verified by simulation accuracy was used to simulate the state of a conventional hypersonic blowdown wind tunnel with the incoming Mach number at the test section inlet in the range of 3–10. A numerical study was conducted on correction methods for IR measurement in the 3–5 μm band at a lateral 90${\text{°}} $ direction for the aircraft and lateral jet, aiming at the interference of wind tunnel background radiation. A progressive measurement procedure based on background field separation was proposed. Through a stepwise experimental design involving cold-model interference calibration and hot-target measurements, the interference signals were stripped. The relationships between the intrinsic radiation of the aircraft surface and lateral jet and the total IR received by the detector were established respectively. The interference factors in IR measurement were identified. The correction formulas for measuring the intrinsic IR of the aircraft surface and lateral jet were derived. The results show that the influence of interference factors on the IR of lateral jet is relatively small, ranging between 1.6% and 2.8%. When the incoming flow Mach number in the wind tunnel test section is between 3 and 5, the influence of interference factors on the IR of the aircraft surface exceeds 10%. Especially at a Mach number of 3, the proportion of interference signals reaches 70.72%.

The lateral jet significantly impacts the external flow field by generating complex shock interactions and affecting the aerodynamic characteristics of the aircraft. To investigate the aerodynamic effects of dual lateral jets under high pressure ratio conditions, a three-dimensional numerical simulation method based on the Reynolds-averaged Navier-Stokes (RANS) equations was developed. The accuracy of the simulation method is verified through single-jet experiments, followed by a comparative analysis of aerodynamic characteristics for different nozzle diameters using the double jet configuration. The results show that the interaction factor for the larger diameter decreases by over 5% compared to the smaller diameter at negative angles of attack. This indicates an adverse effect on the aircraft overload for the larger diameter. At negative angles of attack, the thrust line for the larger diameter is further offset towards the head, with a maximum difference of 2.0%, which reduces the stability of the aircraft. However, at positive angles of attack, the thrust line for the larger diameter is offset towards the tail, with a maximum difference of 0.98%. This slightly improves the stability of the aircraft. Through the flow analysis, it is found that the separation shock for the larger diameter moves forward compared to the smaller one, while the reattachment shock at downstream of the jet reaches the aircraft earlier. This research provides a reference for the configuration selection of lateral jet applied in high-speed aircraft design.

The reaction control system requires fundamental information concerning the process of transverse Jet Interaction (JI) with a crossflow. The effects of JI properties on hypersonic aerodynamic interference, have been the subject of several wind tunnel experimental investigations. In wind tunnel facilities, it is more convenient to perform JI tests with jet fluids of different compositions and with test models of scaled size from those of the actual vehicle. There are a myriad of scaling issues/differences from flight to wind tunnel, and a review of scaling theories has been presented and verified, with calibrated computational fluid dynamics simulations based on HB-1 standard model, as well as NASA Space Shuttle flight data from open literatures. In this work it is found that, in the case of strong jet flow that can penetrate the incoming shock wave, the momentum ratio as the scaling parameter is appropriate, while in the case of weak jet flow that cannot penetrate the incoming shock wave, the mass flow ratio is more suitable as the scaling parameter.

Transonic flows have presented an enduring challenge to experimental research due to their intricate and unsteady flow characteristics. This study investigated the megahertz-frequency Particle Image Velocimetry (MHz–PIV) technique to enhance the resolution of small time-scale flows under the transonic flow conditions. During the measurement, five high-speed cameras alternately and quickly captured images of the same measurement area, and thus obtained ultra-high time resolution particle image data. By employing image processing techniques, optical distortion correction and identification of the common area were achieved. The application of the ensemble correlation algorithm, coupled with spectral analysis of the compressible turbulent flow field based on the velocity field, contributed to a comprehensive analysis. The experiment validated the high-frequency sampling capability of MHz–PIV, which significantly reduces the technology’s dependence on camera performance. This approach offers a refined measurement technique with high spatiotemporal resolution for transonic experiments.

Inertia force and aerodynamic force are often coupled in flapping motion. In order to study the aerodynamic characteristics of a bat flexible membrane wing in flapping motion, it is necessary to separate the inertia force and the aerodynamic force to obtain the aerodynamic force. By setting up a photographic platform based on multi-vision, images of flexible membrane wings with different properties were captured, and a multi-vision algorithm was used to reconstruct the deformation of the flexible membrane wing, so the inertia force can be calculated from the deformation. A six-axis force sensor was used to obtain the real-time force of the flexible membrane wing, then the aerodynamic force can be obtained by eliminating the inertia force, and the law between the inertia force and the aerodynamic force was analyzed. A standard model verifies that the deformation error of this method is 2.36%. The results show that the highly flexible wing membrane has a significant deformation during the flapping process, which is related to both inertia force and aerodynamic force. And with the increase of the thickness of the wing membrane, the inertia force and the aerodynamic force are increased to different extents.

The thrust vectoring technology is the key technology of high-performance aircraft, which can be divided into the mechanical and the fluidic thrust vectoring technology. The fluidic thrust vectoring technology has advantages of simple structure, fast deflection response and low energy consumption, so it has been widely studied. However the problems of jump and nonlinearity of the control law exist in the current rectangular fluidic thrust vectoring nozzles. Meanwhile the maximum deflection angle is small, which seriously limits the engineering application of this technology. A new type of the rectangular multi-wall passive thrust vectoring nozzle is developed in this paper. In order to figure out whether this new nozzle can realize the vector control of the jet with high linearity and larger deflection angle, two control strategies of the multi-section wall under multi-wall passive secondary flow and the variation of the initial passive secondary flow rate are employed. The characteristics of the dynamics and flow field of the jet are studied by means of experiment and numerical simulation. The results show that these control methods improve the linearity of the control law of the force vector angle, and the effective deflection force vector angle is increased. The linearity of the control law increases from 62% to 90.8%, and the maximum deflection force vector angle increases from 18° to 22°. When the initial inlet and outlet area ratio of the passive secondary flow is 1.16, the line-arity is further increased to 94.9%. It is found that there is no obvious separation bubble structure near the wall of the nozzle during jet deflection, which preliminarily explains the principle of the linear continuous control law of the nozzle.

To address the energy consumption issue in data centers, this study investigates the Single-Phase Immersion Liquid Cooling (SPILC) system by an experimental platform. The influential mechanisms of the coolant flow rate and chip power on the SPILC performance are analyzed. The results indicate that the heat transfer of chips located at inlet regions is more influenced by board layout than that of the chips at the outlet region. Moreover, the chip case temperature shows a trend of slight increase followed by a decrease as the flow rate increases. There exists a critical flow rate beyond which further increase does not significantly improve the SPILC performance. Additionally, the chip case temperature rises significantly with increasing power, and lower flow rates result in greater temperature increases. When the flow rate does not exceed 4.0 L/min, the chip power has almost no effect on the system pressure drop. Moreover, the influence of the flow rate on the power usage effectiveness of SPILC systems is significant when the total chip power is low, but this effect is greatly weakened as the chip power increases.

By designing an array integrated electrostatic atomization device with double-layer cross space shielding electrodes, using an optical visualization system and a high-speed spray particle size analyzer, the evolution characteristics of the jet morphology of the integrated electrostatic atomization of stable multiple jets were accurately recorded, and the voltage range of the integrated electrostatic atomization of stable multiple jets and the size distribution characteristics of spray droplets under different integration degrees and supply flow rates were discussed. The results show that the voltage range for stable atomization of multiple jets in the array integrated device is higher than that in the single capillary device, and the number of stable jet strands is weakly affected by the applied voltage. Under the same single capillary supply flow rate, the increase of the total supply flow rate due to the improvement of the integration degree has no obvious effect on the average droplet size, droplet size distribution range and cumulative volume distribution of the overall spray, but can cause the peak value of the volume distribution to decrease accordingly. At the same integration level, the average droplet size increases nonlinearly with the increase of the total supply flow, and the growth rate decreases with the increase of the total supply flow. The array integrated microchannel stabilized multi jet atomization system can significantly increase the supply flow while maintaining the particle size distribution of spray droplets in a narrow range, showing good monodispersity.

This paper is based on a circulating water tunnel, employing high-speed photography and pressure sensors to measure the flow characteristics of the ventilated cavity around a conical axisymmetric body. The focus of the discussion is on the flow characteristics of ventilated cavities at high Froude numbers and the pressure evolution on the axisymmetric body surface for different cavity shapes. The research results indicate that under high Froude numbers, gravity effects can be neglected, and ventilated cavities around the conical axisymmetric body exhibit three typical flow patterns: foamy cavity (FC), intermittent and transparent cavity (ITC), and continuous and transparent cavity (CTC). Among these, FC cavities are found within a small ventilation rate range, accompanied by some small-scale cavity shedding at cavity’s trailing. The transient pressure distribution on the axisymmetric body surface exhibits high-frequency, low-amplitude fluctuations, with the maximum pressure coefficient fluctuation amplitude of 0.18. ITC cavities represent a transitional flow pattern between FC and CTC, characterized by the shedding of large-scale cavity clusters from the cavity’s trailing, resulting in low-frequency, high-amplitude fluctuations in the transient pressure distribution on the axisymmetric body surface, with a maximum pressure coefficient fluctuation amplitude of 0.49. CTC cavities are observed within a large ventilation rate range, exhibiting similar characteristics in cavity shedding and pressure pulsation to ITC cavities. Under specific Froude numbers, during the transition from FC to CTC, the pressure on the axisymmetric body surface at different measurement points exhibits a trend of initially decreasing, then increasing, and finally stabilizing.

As a new type of micro aircraft configuration, the jellyfish-like flying aircraft has the advantages of low noise and flexible maneuvering, which has attracted the attention of the academic community. In this paper, a pair of two-dimensional plates rotating around fixed axes is used as a simplified model of jellyfish-like flapping wings. The disturbed flow fields in static water are measured by the time-resolved particle image velocimetry. By controlling the frequency and angular amplitude of the flapping motion, the influence of the motion parameters on the vortex characteristics and evolution law is studied. The vortex generation, shedding, and interaction processes are analyzed by using the phase average method and the circulation tracking technique. The forming mechanism of thrust is explained from the perspective of vortex dynamics. The experimental results provide certain references for the design of the jellyfish-like flying machines.

In recent years, relevant studies have shown that wing sections with humpback whale-like leading-edge protuberances can effectively delay stall and improve aerodynamic performance after stall. To investigate the flow separation phenomenon of the biomimetic leading-edge protuberances wing section, a particle image velocimetry (PIV) experimental study on the NACA 63₄-021 wing section and modified wing sections with single and double leading-edge protuberances is carried out. The distributions of statistics such as the average velocity of the wing section suction surface and the turbulent kinetic energy at different angles of attack are compared. The proper orthogonal decomposition (POD) method is adopted to analyze the experimental data. The flow structures during the stall process are analyzed from an energy perspective. The study finds that: both single and double leading-edge protuberances wing sections have two-step stall characteristics. When a unilateral stall occurs, there is a stable attached flow at the convex peak, which suppresses the extension of the stall zone to the other side of the convex peak and plays a role similar to a wing blade. Compared to the single leading-edge protuberance wing section, the wing section with double leading-edge protuberances undergoes stall mode transition at smaller angles of attack. The energy of each mode of POD decomposition is related to the stall mode at which the wing section is located. The characteristics of the flow structure in the flow field change with the switching of stall modes. The flow separation mode of the single leading-edge protuberance wing section exhibits instability during the transition from unilateral stall to bilateral stall. This study provides an experimental reference for the theoretical research and potential engineering application of biomimetic leading-edge protuberances wing sections.