The trichromatic mask single-color camera 3D PIV technology uses an optical trichromatic mask element to modulate the imaging light path of a color camera, combining color information with perspective information. This enables the RGB channels of the color-sensitive chip to record images of tracer particles from three perspectives, achieving single-camera 3D flow field measurements. Color cameras typically use a single sensor with a Bayer mask to capture color images. To obtain complete three-channel information of tracer particles, the missing color components in the Bayer image must be restored through demosaicing to extract the three-perspective images. This paper employs High Quality Linear Interpolation (HQLI) algorithms, Gradient Based Threshold Free (GBTF) algorithms, and deep learning algorithms based on U-Net++ neural networks to demosaic particle Bayer images. The quality of three-perspective image extraction by these three algorithms is evaluated using image evaluation metrics and the impact on particle reconstruction quality Q. Simulation experiments on artificially synthesized Gaussian vortex 3D flow fields are conducted to analyze the impact of these algorithms on measurement error. Zero-Net Mass Flux (ZNMF) jet experiments are utilized to study the demosaicing of experimental particle images and analyze the results of transient velocity fields. The results demonstrate that, compared to traditional algorithms (HQLI, GBTF), the deep learning algorithm based on U-Net++ neural networks can more effectively extract three-perspective images of particles, thereby reducing measurement error.

To address the need for measuring soot concentration distributions in complex and spatially constrained combustion flow fields, such as the outlet of aero-engine combustors, this study developed the Two-color-Scheimpflug Laser-Induced Incandescence (2C-Scheimpflug LII) technique and a corresponding measurement system, which obtains soot volume fraction distribution images by employing an oblique imaging method. The technique and measurement system validation experiments were first conducted under specific conditions using three typical ethylene/air flames (McKenna, Gülder, and Santoro). Comparisons with measurement results obtained by other researchers under the same conditions demonstrate the feasibility of the 2C-Scheimpflug LII system. The test results in standard flames show that the lower detection limit of soot volume fraction measurements of our system is approximately 2.0 × 10⁻⁹. Further, the 2C-Scheimpflug LII technique was applied to measure the spatial distribution of soot volume fractions at the transverse cross-section of the outlet and the longitudinal section of a single-sector, dual-swirler aero-engine model combustor. The preliminary results show that the soot volume fraction distribution at the combustor outlet exhibits high instantaneous variability and randomness, while within the combustor, the soot volume fraction displays a V-shaped distribution.

Trichromatic mask particle image velocimetry employs optical elements to alter imaging paths, integrating color and perspective information for single-camera three-dimensional flow velocity measurements. Due to the small imaging viewing angle between three perspectives, conventional algebraic reconstruction techniques yield elongated particle distributions and low directional resolution along the imaging depth direction. To mitigate elongation effects and enhance three-dimensional particle reconstruction quality, convolutional neural networks (CNNs) are applied. Utilizing the AIPR (Artificial Intelligence Particle Reconstruction) architecture, an algorithm tailored for small viewing angle in trichromatic mask PIV is developed. Evaluation involves synthetic image data and artificial flow fields varying in different viewing angles and particle concentration. Results demonstrate that compared to traditional algebraic methods, the proposed method improves reconstruction quality by 30% at the particle concentration of 0.05 ppp. Additionally, it achieves a reconstruction speed of up to 2.8 times faster than traditional methods, enabling the rapid three-dimensional reconstruction of the particle field and enhancing axial reconstruction quality.

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.