The complex wave structures generated by a ship significantly influence the wave-making resistance, which is a critical factor in optimizing vessel performance and hull design. Traditional measurement methods, such as wave height gauges, are limited by sparse measurement points, low spatial resolution, and potential interference with the wave surface, making them unsuitable for full-field wave surface measurements in complex wave systems. To address these limitations, a non-contact measurement method for ship wave surfaces based on multi-view stereoscopic vision is presented. The proposed scheme includes: applying neural networks for spatial calibration and the three-dimensional reconstruction of feature points, using an ant colony optimization particle tracking algorithm for perspective matching, and conducting principal component analysis for still water baseline correction. Experimental results demonstrate that the proposed method can realize the wide-area free-surface measurement of the ship wave and effectively capture typical wave characteristics at various ship speeds. Compared to traditional measurement, the stereoscopic vision approach offers enhanced flexibility and suitability for real-time measurement of dynamic wave fields over large areas.

In liquid environments, nanoscopic soft materials typically adopt a cap-like shape to maintain their stability. Therefore, their morphology characterization and identification remain challenging in the liquid phase. The present study employs Atomic Force Microscopy (AFM) to achieve high-resolution imaging of subaqueous micro-nano blisters, polymer droplets, and surface bubbles. By analyzing the morphological changes in various scanning forces, the morphological characteristics of these nanoscopic soft materials are investigated. Subsequently, nanoindentation tests are conducted to analyze the interaction between the probe and the solid-liquid-gas interfaces, and their mechanical properties are obtained. The results show that under a scanning force of 0.50 nN, all blisters, droplets, and bubbles exhibited cap-like shapes. Under a higher scanning force (5.0 nN), the blister morphology remained constant, the droplet volume decreased, and the bubble disappeared. Force-distance curves at the vertex under a load of 3.0 nN indicate that all the three experienced elastic deformation. The probe has to overcome greater adhesion force to detach the droplet, while bubbles display a two-stage elastic deformation. Furthermore, due to the influence of anchoring effects, the considered objects exhibit stronger resistance to deformation near the edge of the spherical cap. The modulus of the polymethyl methacrylate (PMMA) nanofilm is independent of the applied load during elastic deformation caused by blister and it is estimated to be around 3.38 GPa. The interfacial tension of small-sized PDMS droplets underwater is approximately 37.3 mN/m, while the gas-liquid interfacial tension of surface bubbles is approximately 32.5 mN/m.

Rotating structures are commonly used in the transportation and energy generation fields. As a non-contact optical measurement technique, the Three-Dimensional Digital Image Correlation (3D–DIC) allows for the full-field spatial measurement of displacement, deformation and vibration of rotating objects in operation, which provides important data support for damage detection, health monitoring and aerodynamics structural optimization design. A review of the extensive studies on the development and application of 3D–DIC is presented. First, the principle of 3D–DIC and its rotating object measurement setup are introduced. Particularly, in cases when the deformation involves a large rotation, a de-correlation problem emerges. To address this, many scholars have proposed initial value estimation and shape function design methods for large-rotation objects, which are investigated extensively here. Several application cases utilizing 3D–DIC for damage detection and vibration measurements are introduced, showing its unique advantages compared with traditional deformation measurements methods. Lastly, the applica-tion prospect and development trend of this technology are prospected.

Icing in the form of impinging supercooled droplets accounts for the most cases of aircraft icing. This review provides an in-depth discussion over single sessile droplet icing, impinging of a droplet, and the combined icing of an impinging droplet. Special focus is placed on icing of an impinging supercooled large droplet, which is of significant importance in the study of aircraft icing. The effects of a few key parameters, such as the impinging speed and droplet size, are studied and analyzed. Meanwhile, by studying the factors influencing the dynamics of droplet impact icing, the icing phenomena of room temperature droplets and supercooled droplets are investigated, and it turns out that the initial thermophysical properties, heat transfer mechanisms and phase transition modes vary drastically, and the kinetic behaviors such as nucleation time, ice crystal type, interfacial heat transfer characteristics and phase advancing velocity as well as the droplet morphology are significantly different. The effects of the impact velocity, particle size and surface wetting on droplet impact icing are indicated, providing a theoretical basis and reference for the design and development of anti-deicing technology for aircraft.

Aero-engine nacelle lip icing poses a significant threat to engine safety. Understanding the icing characteristics on the nacelle lip surface provides a crucial basis for analyzing icing hazards and designing protective measures. This study focuses on a self-designed nacelle model, employing a combination of numerical simulations and ice wind tunnel experiments to systematically analyze the icing characteristics and influencing factors on the nacelle lip surface. The findings show that an increase in intake airflow and a decrease in incoming airflow speed result in a reduction of the actual airflow attack angle at the lip, which in turn reduces the ice coverage on the inner surface and increases the ice coverage on the outer surface. As the incoming temperature and droplet diameter decrease, the ice coverage on both the inner and outer surfaces of the nacelle decreases, while the ice thickness increases with lower temperatures but decreases with smaller droplet diameters. The droplet diameter has a minor effect on ice thickness. Furthermore, sensitivity analysis of the average ice thickness for different parameters indicates that the droplet size has the most significant impact on nacelle lip icing, followed by intake airflow, incoming wind speed, and incoming temperature. The results of this study provide important theoretical support for the research on icing mechanisms and the design of anti-icing measures for aircraft engine nacelle.

The phenomenon of droplets hitting the vibrating wall widely exists in industrial production and nature, and the vibration phase of the wall affects the velocity and acceleration state of the wall, which in turn affects the dynamic characteristics of the impact droplets. Therefore, it is of great significance to determine the influence of different vibration phases on the dynamic behavior of droplets hitting the vibrating wall. The process of droplets hitting the vibrating wall of different collision initial phases was observed through high-speed camera experiments, and the influence of the initial phase of collision on the spreading and oscillation characteristics of the droplets after hitting the vibrating wall under different vibrational conditions was analyzed. The results show that the initial phase of the collision determines the relative position and motion trend of the wall, and the spreading time of the droplets increases first and then decreases with the increase of the initial phase of the collision. The maximum spreading factor decreases first and then increases with the increase of the initial phase of the collision. According to the force analysis and study of the oscillation characteristic curve of the droplet oscillation process, it is found that the value of the maximum central height of the initial phase of the 90° collision is the largest, and the initial phase of the droplet collision has little influence on the stability time of the droplet. When the amplitude of the wall vibration increases, the change of the droplet spreading characteristics and oscillation characteristics with the initial collision phase is basically unchanged. The trend of droplet spreading characteristics and oscillation characteristics with the initial collision phase is mainly frequency-related.

Focused Laser Differential Interferometer (FLDI) has been widely used in the spatial measurement of hypersonic boundary layer instability. However, at present, the FLDI system used in China can only measure the density change of one or two spatial points at a time, and the experimental measurement along the streamwise direction or normal direction in the boundary layer cannot be carried out efficiently, and the correlation of the density fluctuation signal between the two measuring points obtained from two wind tunnel runs is poor. In this work, a setup of multi-point FLDI measurement system is designed and built, which can simultaneously obtain the second mode instability waves characteristics of multiple measuring points in the hypersonic boundary layer in a single wind tunnel operation period. Based on the Φ0.25 m low noise Mach 6 hypersonic Ludwieg tube wind tunnel in Huazhong University of Science and Technology (HUST), the instability waves measurement experiment of a sharp cone’s boundary layer with a freestream Reynolds number of 1.19 × 10⁷ m⁻¹ is carried out by using the multi-point FLDI measurement system. The experimental results show that the multi-point FLDI system successfully captures the typical second mode instability waves with the main frequency in the range of 316.4−322.3 kHz, which is agree with the main frequency 319.3 kHz of those measured by a PCB sensor at almost the same measuring point. Furthermore, through the cross-correlation analysis of the density fluctuation data obtained based on the multi-point FLDI system, it is found that the phase velocity of the instability waves between the adjacent measuring points is the same, which is consistent with the propagation velocity characteristics of the second mode instability waves in the boundary layer. The experimental results verify the data obtained by multi-point FLDI are quiet reliable. Above all, the multi-point FLDI system has the advantages of high efficiency, low cost, high accuracy and high spatial resolution, so it is expected to be applied to the basic research investigation in the future, such as hypersonic boundary layer instability and receptivity mechanism.

Identifying the main sound sources under typical aircraft operating status and estimating their radiated sound power are of great significance for aircraft low-noise design and noise pollution assessment. In this study, a beamforming algorithm is employed to develop a method that utilizes a sound array to identify the external main airframe sound sources and estimate their radiated sound power levels during both aircraft ground operation and low-altitude flight. Additionally, a method for analyzing the contribution of these main sound sources is proposed. A flight test technical plan is then presented, which centers around the measurement of sound–aircraft–environment parameters. The proposed methods are validated through a series of sound tests conducted on a fighter plane and a large transport aircraft, utilizing a well-designed non-uniform linear acoustic array and a planar multi-arm spiral acoustic array. The results demonstrate the efficacy of the proposed methods in accurately identifying the main sound sources within the frequency range of 200–2000 Hz, as well as estimating the radiated sound power levels of the aircraft engine and airframe sound sources, thereby enabling the determination of their respective contributions to the overall sound power.

Magnetic Resonance Velocimetry (MRV) is becoming a crucial experimental technology for flow field visualization, flow heat transfer mechanism investigating, and numerical model validation. In recent years, the MRV technology has been employed in experimental investigations of coolant channel flow fields in nuclear reactor core fuel assemblies, providing fundamental data for core thermal hydraulic characteristics research and safety analysis. A literature overview of the research progress using MRV technology in experimental studies on the flow field of reactor core fuel assembly channels is presented in this paper. The three-dimensional coolant flow field within a 2 × 2 simplified rod bundle channel was measured using MRV, and quantitative comparisons were conducted with PIV measurement results from previous literature to investigate the key parameters influencing the measurement accuracy of MRV. In this study, MRV tests were performed under two conditions of Re = 12500 and 25000, respectively. The results indicate that the axial velocity from MRV within the investigated channel has a maximum discrepancy of 5.05% compared to the PIV data reported in the literature.

A huge change in the development mode of the aerospace field is taking place due to modern digital technology. As the cradle of aircraft, the development of wind tunnel must deeply grasp the new trend of digital development, actively adapt to the digital development mode of Model-Based System Engineering (MBSE), and strive to promote the digital transformation of wind tunnel design, construction and operation. This paper introduces the development and application results based on MBSE in relevant industries at home and abroad, compares and analyzes the current situation of digital development of wind tunnels, points out the shortcomings of the existing development mode, and constructs a digital development mode based on MBSE. The design, construction, operation, and maintenance process of this model and its application in the development of a large wind tunnel are introduced in detail. Finally, in order to meet the requirements of model-based digital development of wind tunnels, a collaborative design and simulation platform is constructed, which provides effective support for further improving the level of digital development of wind tunnels.

The dynamic characteristics and temperature effects of fast-response pressure-sensitive paint are the key factors for determining whether it is applicable to unsteady flow fields. This paper proposes a method for dynamic calibration of fast-response PSP (Pressure-Sensitive Paint), establishes a controllable calibration standing wave field through a dynamic calibration system that can adjust the temperature and frequency, and researches on the influence of temperature on pressure sensitivity and frequency response characteristics of the developed PC–PSP (Polymer Ceramic Pressure-Sensitive Paint). Dynamic calibration reflects the PSP characteristics more accurately than static calibration. Frequency changes have a significant effect on the pressure sensitivity of PC–PSP, and the influence of temperature on the pressure sensitivity of PC–PSP is not significant, while the temperature sensitivity of PC–PSP cannot be ignored in measurement. At the same temperature, the gain decreases and the phase lags behind as the frequency increases. At different temperatures, with the increase of the temperature, the cut-off frequency increases, and the phase lag gradually weakens. The dynamic pressure sensitivity at 20 ℃ and 0.5 kHz is 0.59%/kPa, and at 20 ℃ and 3 kHz is 0.51%/kPa, with the cut-off frequency being 4.8 kHz. Therefore, the performance of the tested PC–PSP is stable and meets the engineering requirements for unsteady flow field pressure measurement tests in low-speed wind tunnels.

The phenomena of millimeter sized water droplets impinging on a thin stationary water film at multiple incidence angles are experimentally investigated in this work. The effects of water droplet impacting Weber number, incidence angle, and dimensionless water film thickness on crown characteristic parameters and splash threshold are comprehensively analyzed. The experimental results show that as the tangential impacting Weber number is constant, the height of the liquid crown increases with the increase of the incidence angle, but the evolution law of the liquid crown radius is independent. As the normal impacting Weber number is constant, the height and radius of the liquid crown increase with the decrease of the incidence angle. As the dimensionless water film thickness increases, the radius of the liquid crown top decreases, while the radius of the liquid crown bottom increases. The height of the liquid crown first increases and then decreases with the increasing dimensionless film thickness. The prediction models are summarized to describe the experimental results of the maximum liquid crown height and the evolution of the liquid crown radius over time when the dimensionless water film thickness is 0.1. The splash threshold models for the dimensionless water film thickness of 0.1 and 0.2 are also summarized, respectively.

Identifying the surface aerodynamic heat flux using the time-series temperature data collected from internal measurement points of the thermal protection system is a crucial method for acquiring the surface thermal environment of high-speed aircraft. Analyzing and quantifying the uncertainty of identification results can provide a quantitative criterion for the validity of test data and support the safety assessment of “ extreme thermal service conditions" in thermal environments. Surface heat flux identification is a typical ill-posed problem, and measurement noise can lead to explosion in high frequency, so regularization is needed to obtain stable solution. Regularization reduces the uncertainty of estimation results, but it also brings obstacles to uncertainty quantification analysis. Therefore, regularization is regarded as a low-pass filter, and the identification results are transformed into the frequency domain to analyze the uncertainty of identification results. Finally, the uncertainty sources of typical aerospace surface heat flux sensor are analyzed. The effectiveness of uncertainty quantification is verified by Monte Carlo method.

The HyTRV (Hypersonic Transition Research Vehicle) is designed for the study of hypersonic complex three-dimensional boundary layer transition, the shape of which is close to that of the real aircraft. The principal objective of this research is to gain deeper understanding of the stability characteristics and transition mechanism of the boundary layer over the HyTRV model and verify the results of numerical simulation and theoretical analysis. In the Mach 6 wind tunnel, the transition front and unstable wave evolution characteristics at the bottom of the HyTRV model, obtained by the infrared thermal imaging technology and high-frequency pressure sensors respectively, were compared with the results of one-dimensional stability analysis. At the angle of attack of 2 degrees, the instability crossflow region on the windward side, which is located after the front of transition at the centerline, shrinks and moves backwards compared to the results with zero angle of attack. Broadband high-frequency signals, containing multiple disturbances in different frequency bands, were found in the crossflow region. The interaction between various disturbances widens the frequency spectrum. Then the boundary layer appears to breakdown. The signal amplitude reaches saturation before the transition is complete.

The transition of the boundary layer from laminar to turbulent flow significantly impacts the flight stability and safety of aircraft. The aerodynamic heating of the nose of the hypersonic vehicle during flight increases the nose temperature, which affects the stability and transition characteristics of the boundary layer. In this paper, experiments were carried out using a Mach 6, 0° angle of attack conical model to explore the influence of the nose temperature and bluntness on the stability of the conical boundary layer. The wall temperature ratio (the ratio of the nose temperature to the total incoming temperature) T_w/T₀ varies between 0.6 and 1.3, while the nose bluntness R ranges from 0 to 5 mm. Measurements of boundary layer transition and unstable wave were carried out utilizing infrared thermal imaging and high frequency pulsating pressure sensor (PCB). The experimental results show that changing the nose temperature and bluntness affects boundary layer stability. For a sharp cone (R < 50 μm), the nose heating promotes the development of the second mode wave. However, for R = 0.5 mm and 2 mm cones, increasing T_w/T₀ from 0.6 to 1.1 inhibits the development of the second mode wave, while a further increase to 1.3 encourages it. For R = 5 mm cones, nose heating stabilizes the second mode wave.