Yuntao Guo

Ionic-liquid electrospray thrusters are promising micropropulsion devices for small spacecraft; however, excessive ion-beam divergence reduces useful axial momentum and increases particle interception by downstream electrodes. This study investigates a compact converging-grid configuration designed to reduce the maximum beam-divergence half-angle while preserving high ion-transmission efficiency. Unlike architectures employing a dedicated focusing electrode or a separate focus grid, the proposed configuration uses the downstream biased accelerator extractor itself as an ion-optical field-shaping element. A three-dimensional bidirectionally coupled electrostatic–particle model was used to account for particle-induced space-charge effects. Following preliminary screening of the rational Bezier emitter-tip geometry, a multidimensional all-combinations study comprising 15 625 configurations was performed for six geometrical and electrical variables, followed by targeted one-dimensional analyses and numerical mesh- and particle-number sensitivity assessments. The configurations were evaluated using the maximum divergence half-angle and ion-transmission efficiency as the principal beam-transport metrics, together with the trajectory-based mean effective axial specific impulse and maximum emitter-tip electric-field magnitude as complementary indicators. The selected configuration, defined by ( w = 0.56 ) , ( C 1 = 0.04 mm ) , ( C 2 = 0.05 mm ) , ( r 3 = 0.23 mm ) , ( r 4 = 0.28 mm ) , ( V emi = 3500 V ) , and ( V acc = 2800 V ) , reduced the maximum divergence half-angle from approximately ( 44 ∘ ) for the reference single-extractor configuration to approximately ( ∼ 2 5 ∘ ) , corresponding to an approximately ( 43 % ) reduction, while maintaining an ion-transmission efficiency of approximately ( 89.3 % ) . Under the investigated numerical conditions, these results indicate that the converging downstream accelerator extractor can substantially reduce the outer divergence envelope of the transmitted ion beam while preserving high ion transmission without requiring a separate focusing grid.

Transient shock evolution, pressure-wave dynamics, and shock-wave/boundary-layer interactions (SBLI) induced by moving boundaries are investigated during the opening process of a supersonic inlet. A dynamic-mesh-based numerical framework is developed to couple the rotational motion of the inlet cover with the unsteady compressible flow field. Using an integral solid rocket ramjet as a representative application, the effects of rotational inertia, spring preload, and incoming Mach number are systematically examined. The results show that inlet-cover motion is governed primarily by aerodynamic loading, with aerodynamic torque remaining substantially larger than both the spring-preload torque and gravitational moment throughout the opening process. Rotational inertia is identified as the dominant structural parameter. Although it does not alter the sequence of transient shock evolution, it changes the characteristic timescale of moving-boundary evolution and regulates the synchronization among geometric opening, shock evolution, pressure-wave propagation, and SBLI development. In contrast, spring preload only slightly modifies the opening history and has little influence on the subsequent transient flow evolution. Within the investigated range of Ma = 2.8–3.2, the incoming Mach number exhibits the strongest influence on transient shock evolution among the parameters considered. Increasing Mach number enhances aerodynamic loading and flow compressibility, accelerating moving-boundary evolution, pressure-wave propagation, and SBLI development, thereby promoting earlier shock displacement and stronger flow separation. The results further demonstrate that SBLI acts as the key intermediary mechanism linking pressure-wave propagation to transient shock evolution. Overall, transient flow evolution is governed by the coupled interaction between moving-boundary evolution and internal compressible-flow dynamics, providing new physical insight into moving-boundary compressible flows under time-varying geometric constraints.

Residual propellant remaining in a nozzleless booster introduces additional mass supply and heat release during stage transition of an integrated solid rocket ramjet (ISRR), while the mechanisms governing its influence on transient backpressure development remain unclear. The present study investigates the physical mechanisms of backpressure buildup induced by residual propellant from a tandem dual-burning-rate nozzleless booster. Results show that the booster operating process is governed by the coupled effects of propellant gas generation, dynamic thermodynamic-throat evolution, and chamber-volume variation. During stage transition, residual propellant modifies the afterburner flow and combustion processes through the combined effects of residual-propellant mass addition, sustained release of fuel-rich species, particle oxidation, and aft-chamber flow confinement, which collectively accelerate pressure accumulation and govern transient backpressure evolution. When both high- and low-burning-rate residual propellants remain present (Case b), a larger reacting region extends downstream and rapidly establishes a high-temperature, high-pressure region, producing the strongest backpressure response. After depletion of the high-burning-rate residue (Case c), the reacting region contracts toward the aft chamber, resulting in a substantially weaker pressure response. Compared with the no-residual condition (Case a), the maximum and stabilized dimensionless backpressures increase by factors of 9.15 and 10.80 in Case b, and 3.67 and 4.45 in Case c, respectively. Comparison with the reference anti-backpressure value reported for a geometrically similar inlet under the same flight condition suggests that residual-propellant combustion may substantially reduce the available inlet-starting margin under comparable operating conditions. These results identify high-burning-rate residual propellant as the primary contributor to transient backpressure buildup and provide new physical insight into the coupled evolution of mass addition, combustion, and pressure-wave development during ISRR stage transition, offering guidance for transition-timing design and residual-propellant management.

The geometric configuration of a side-mounted inlet bend plays a crucial role in determining the outlet flow quality of the inlet duct and the mixing combustion organization within the afterburner of a solid rocket ramjet. In this study, three-dimensional numerical simulations were conducted to investigate the effects of bend length, turning angle, and expansion ratio on the flow characteristics at the inlet bend exit and the subsequent air/primary-gas combustion process. The coupled mechanisms by which bend geometric parameters affect flow distortion, total pressure recovery, and mixing–combustion organization were elucidated. Results indicate that under cold-flow conditions, the outlet flow quality is governed by the combined effects of centrifugal forces, pressure gradients between the inner and outer bend walls, and the evolution of Dean secondary flow structures. Larger bend lengths and expansion ratios tend to intensify low-momentum fluid accumulation and aggravate total pressure distortion, whereas variations in turning angle primarily modify transverse pressure redistribution and secondary flow transport pathways, without exhibiting a clear monotonic trend. Under reacting conditions, heat release and gas expansion significantly reconstruct the pressure field and secondary flow structures within the bend, suppressing large-scale Dean vortices and substantially improving total pressure recovery. However, the influence of geometric parameters differs markedly from that observed under cold-flow conditions. Moderate bend lengths and expansion ratios provide a more favorable compromise between pressure recovery, flow uniformity, and combustion stability. Increasing the turning angle enhances transverse transport and turbulent mixing, thereby promoting combustion efficiency, albeit at the cost of additional aerodynamic losses. Although excessively small expansion ratios facilitate rapid near-field mixing, they weaken the transverse redistribution capability required to mitigate thermally induced pressure non-uniformities, resulting in localized high-pressure cores and increased outlet flow distortion after combustion. Overall, the geometric parameters of the inlet bend exert a significant influence on the outlet flow quality and its coupling with combustion organization. Proper geometric optimization can effectively improve the integrated aerodynamic and combustion performance of solid rocket ramjet engines.

Solid rocket motors play an important role in underwater propulsion systems and trans-media vehicles. In this paper, the experimental study has been conducted on the ignition and operation process of a solid rocket motor under deep-water conditions. A variable-water-depth test system and a multi-point pressure measurement scheme for the nozzle is established to simulate the ignition start-up and operating characteristics of the motor at different water depths. Two experimental solid rocket motors are designed under different expansion states and the pressure oscillation characteristics of the motor and the separation flow in the divergent section are measured. The results show that the intensity of pressure oscillations increases with water depth, reaching a maximum at a depth of 330 m. Intermittent unchoked flow in a large expansion ratio nozzle under deep-water conditions is experimentally observed for the first time. This phenomenon causes the combustion chamber pressure to deviate from its sea-level value and exhibit sawtooth-shaped oscillations. There exists a critical depth for the occurrence of intermittent unchoked flow at the throat, which determines whether the combustion chamber pressure is affected by unstable jet behavior downstream of the nozzle throat. As the water depth increases, the flow separation line shifts closer to the throat. A lower equilibrium pressure in the combustion chamber results in a separation line position closer to the throat. At the depth of 330 m, the separation line moved upstream by 50.54 mm in high-pressure motor and 54.96 mm in low-pressure motor.

Aluminum is widely used as a high-energy additive in aerospace propulsion, where its combustion performance can be improved by particle size refinement and alloying with elements such as Mg and Li. In this study, the ignition and combustion characteristics of aluminum particles (50 nm–30 μm) and Al–Mg, Al–Li, and Al–Mg–Li alloys were systematically investigated using thermal analysis and laser ignition. Key parameters, including ignition energy, ignition delay, combustion temperature, and spectral features, were evaluated. The results show that both particle size reduction and alloying significantly enhance aluminum reactivity. Alloy particles exhibit lower ignition thresholds, shorter ignition delays, and higher combustion intensities, primarily due to micro-explosion-induced fragmentation. A theoretical model coupling laser intensity, ignition temperature, particle size, ignition delay, and ignition energy was developed, showing good agreement with experimental data. Spectral and combustion wave temperature analyses further confirm that smaller particles and alloying improve combustion efficiency under identical conditions. Post-combustion microstructural analysis reveals cracks and openings in alloy particles, facilitating oxygen transport to the particle interior. This behavior is governed by the combined effects of particle size, melting point, and micro-explosions. These findings provide mechanistic insight and experimental support for optimizing aluminum particle design and alloy composition in propulsion applications.

Combustion efficiency is a key performance metric for solid fuel ramjets (SFRJs) and is significantly affected by the combustion organization within the afterburning chamber. To enhance the combustion of pyrolysis fuel with air, a diaphragm structure is commonly utilized to control the flow field and to improve mixing. This paper utilizes numerical simulations to analyze the flow and combustion characteristics of SFRJs with various diaphragm heights and proposes a recommended design height range for this scale ramjet. The results indicate that diaphragms can effectively promote fuel–air mixing and improve combustion efficiency by inducing recirculation zones and increasing turbulence intensity. There is an upper limit to the improvement in combustion performance with increasing diaphragm height (≤0.873). Excessively high diaphragms reduce the effective flow area, increase flow velocity, and shorten the residence time of reactants, thereby limiting further enhancement. Based on the performance gain coefficient analysis, the recommended design range is identified as Hd/Rc (ratio of diaphragm height to combustion chamber radius) = 6/25 to 8/25. In addition, the diaphragm height shows minimal effect on the average regression rate, confirming its primary role as a flow control structure rather than a direct driver of fuel consumption. This study provides theoretical support and engineering guidance for the design of flow control structures in SFRJs.

The stage transition process of solid propellant ducted rockets represents a critical phase to ensure the successful initiation and stable operation of the ramjet mode. This transition, from the booster phase to the ramjet phase, requires a rapid and stable flow field adaptation within a short time frame. In this study, unsteady numerical simulations employing dynamic mesh techniques are conducted to investigate the influence of varying inlet cover opening durations and transition Mach numbers on the internal flow characteristics of the inlet during stage transition. The results reveal that the interaction between shock waves and boundary layers governs the evolution of the internal flow structure and pressure distribution. The duration of inlet cover opening is shown to significantly affect shock strength, flow reattachment behavior, and pressure fluctuation intensity. Prolonged opening durations lead to an extended separation region near the inlet lip, which inhibits inlet restart and degrades transition performance. The optimal inlet cover opening duration for the given inlet configuration is identified to be 10 ms. Furthermore, the timing of the outlet cover opening also plays a crucial role in the stage transition process. Opening the outlet cover during the early to mid phase of the filling process effectively shortens the transition duration and enhances total pressure recovery, thereby facilitating a more rapid and stable establishment of ramjet operation. This study provides theoretical insight into the aerodynamic mechanisms underlying the stage transition and offers valuable guidance for optimizing the performance of solid propellant ducted rockets.