Due to the continuous enhancement of pointing accuracy, detection range, and other performance metrics, reflector antennas are evolving toward high-frequency bands, increased gain, and overall improved performance. In this context, the influence of structural uncertainties, such as antenna feed and main reflector accuracy, on performance has become notably prominent. The inherent interdependence between structure and electromagnetic properties accentuates the impact of structural and electromagnetic uncertainties on electrical properties like gain and pointing of reflector antennas. Addressing the challenge of analyzing the impact of uncertainty on antenna performance and ensuring optimal functionality, this paper derives the phase interval resulting from feed position deviations in the aperture plane. Additionally, it investigates the influence of feed direction uncertainty on the amplitude of the aperture plane and provides an expression for the pattern under main reflector shape uncertainty. Building upon these analyses, an electromechanical coupling model for the structural uncertainty of the reflector antenna is established, unveiling the influence mechanism of key structural uncertainty factors on electrical performance. To illustrate these concepts, a calculation example is presented using the 110 m antenna design model from Qitai. The discussion delves into the profound impact mechanisms of feed and main reflector uncertainties on antenna gain, side lobes, pointing, and other performance aspects.

The pulsed discharge energy deposition technique is an innovative method for enhancing mixing efficiency in supersonic transverse jets. By generating transient gas bubbles to perturb bow shock structures, this approach effectively induces Richtmyer–Meshkov instabilities, amplifying vorticity generation and intensifying turbulent mixing. However, the coupling mechanisms between bubble dynamics and jet unsteadiness remain poorly understood. Under Mach 2.47 inflow conditions, this study investigates high-frequency pulsed bubble effects on jet penetration, dispersion behavior, and fluctuation evolution using high-speed schlieren imaging, planar laser scattering, and focused laser differential interferometry. The experiments capture the complete evolution of discharge-induced bubbles, from their formation to the development of hairpin vortices, and culminating in turbulent collapse. Results show that bubble-induced shock modulation weakens bow shock intensity, with reduced compressibility increasing the effective momentum flux ratio by 36.6%, corresponding to a 9.2% rise in Mach disk height and enhanced spanwise spreading. Additionally, interactions between bubbles and shock waves enhance vorticity behind the bow shock, leading to sudden vortex growth (at x=5–10δ) at the jet wake, and creating large-scale coherent structures that exceed three times the thickness of the boundary layer in vertical scale.

The Vortex Particle Method (VPM) represents continuous, incompressible, unsteady, viscous flows by using singular, independent particles represented by small concentrations of vorticity. The VPM is particularly convenient for representing viscous flow fields, which is useful for simulating vortex reconnection in an unbounded domain and flow mixing. This study aims to evaluate the accuracy of the VPM predictions using, among others, the diffusion velocity method to represent the molecular diffusion of vortex structures. The VPM can also represent the viscous diffusion of turbulent flow fields using a turbulent viscosity model. However, such models are defined up to a given constant. The dynamic computation of this constant is required for the turbulence model to encompass the diffusion of all the flow scales. A new fast dynamic technique based on the local loss of enstrophy in the flow is proposed. The coupling of the Diffusion Velocity Method with this dynamic procedure is presented. Finally, the method is validated on several known analytical flow solutions, such as the Lamb–Oseen vortex and isolated and leapfrogging vortex rings.

A nonlinear aeroelastic model for a flexible panel is expanded to consider the effect of the viscous boundary layer using a Reynolds-averaged Navier–Stokes (RANS) solver. This study presents the use of the dynamically linearized time-domain approach reduced-order model to compute the generalized aerodynamic force from a computational fluid dynamics (CFD) simulation within the nonlinear aeroelastic solver. Results obtained using the RANS approach are compared with those using an inviscid CFD solver and also the full potential flow. The flutter onset critical condition is obtained based on the limit cycle oscillation amplitudes for a range of supersonic Mach numbers. Additionally, the effect of different viscous boundary-layer thicknesses is explored. Conclusions are made based on the application of this more complete aerodynamic theory, particularly for near-transonic Mach numbers.

Features of a normal shock/turbulent-boundary-layer interaction at Mach 1.54 are examined using large-eddy simulation. The geometry used consists of an internal passage divided into an upper unchoked section and a lower choked section by a shock-holding plate. This geometry and incoming Mach number produce a nearly normal shock and lambda-foot structure on the lower wall, leading to separation in the incoming, Reθ=1895, turbulent boundary layer. The shock structure is slightly detached from the leading edge of the shock-holding plate and exhibits unsteady but statistically stationary oscillations in the streamwise direction. Unsteady and time-averaged quantities are then examined upstream, inside, and downstream of the interaction region. The interaction of the front foot of the λ-shock with the turbulent boundary layer induces strong separation. The characteristic low-frequency oscillation in the main trunk of the shock is also present in the spectra of wall pressure fluctuations, turbulent kinetic energy, and instantaneous separation location. The shock-induced separation amplifies and redistributes the turbulent Reynolds stresses from the incident boundary layer to the separated shear layer.

In numerical combustion research, accurate chemical reaction mechanisms for heavy hydrocarbons are essential to investigate the flame dynamics of sustainable aviation fuels (SAF) relative to conventional jet fuels. However, accounting for hundreds or thousands of reactions makes finite-rate chemistry (FRC) large-eddy simulations (LES) impractical for realistic engineering scenarios. Here, a multidimensional chemistry coordinate mapping (CCM) approach is employed to reduce the computational expense of FRC-LES for jet fuel combustion. In the CCM methodology, the flow transport equations are directly integrated (DI) in the computational cells in physical space, whereas the chemical reactions are integrated in a phase space made up of a few principal variables, improving simulation efficiency. In a turbulent premixed bluff-body burner, the LES-CCM method is compared with experimental data and LES-DI results for conventional fuels Jet A (or A2) and JP5 (or A3) and alternative synthetic fuels, referred to as C1 and C5. The skeletal HyChem reaction mechanisms are utilized, ranging from 40 to 50 species and 200 to 300 reactions. A satisfying compromise is achieved between the accuracy of the results and the speedup factor, on average around three, depending on the CCM phase space dimension and the size of the reaction mechanism.

The nonlinear interactions between the self-excited and forcing frequencies in the open-loop control of thermoacoustic oscillations remain insufficiently understood. In this experimental study, we investigate the nonlinear forced response of a self-excited hydrogen-enriched swirling flame. First, a theoretical analysis is performed based on the perturbed G equation and reveals that changes in the flame’s response at the self-excited frequency arise not only from the self-excited mode itself but also from nonlinear interactions between the self-excited and forced modes. To quantify these contributions, data-driven bispectral mode decomposition is then applied, which uncovers that interactions involving the sum frequency and the forcing frequency, as well as those between the second harmonic and the self-excited frequency, amplify the self-excited mode. In contrast, the interaction between the difference frequency and the forcing frequency has a suppressive effect on it. At last, the flame’s forced response is analyzed spatially, and the flame tip is identified as the region most receptive to external forcing, whereas the flame root remains relatively insensitive. This study advances the understanding of nonlinear flame dynamics under dual-frequency excitation, providing new physical insights for optimizing open-loop control strategies, particularly the selection of forcing frequencies.

Shock-induced combustion in a direct-connected combustor was experimentally and numerically investigated at Mach 10 flight conditions. This study aimed to reveal the establishment of shock-induced combustion and its transition to partially premixed-dominated combustion in practical combustor flows, highlighting the critical effects of fuel injection and shock wave/boundary-layer interaction on shock-induced combustion stability. The combustor consisted of a constant-area section, a diverging section, and two shock-generating wedges symmetrically mounted on the top and bottom walls of the combustor. Hydrogen fuel was injected through two symmetrical fuel injectors. Each fuel injector had three identical struts equally spaced along the spanwise direction. Oblique shock waves were generated through wedges at the end of the combustor diverging section. Wall pressure and hydroxyl radical (OH*) flame chemiluminescence imaging were used to characterize the transient phenomenon of flame stabilization and upstream propagation. Large-eddy simulations were conducted to analyze the unsteady flame propagating processes in association with experimental results. The nonuniformity of prewave parameters induced by strut injection, coupled with the exacerbation of boundary-layer separation caused by shock-induced combustion, led to combustion instability and engine performance deterioration.