The instability of combustion and low rate of fuel regression of the traditional hybrid rocket motor limit its applications. The distributed tube injector (DTI) presents a promising solution by significantly enhancing average fuel regression rates up to 8 mm/s, while maintaining stable combustion, particularly at the laboratory scale. In this study, we investigate the DTI at high fluxes to determine the characteristics of stability of combustion in hybrid rockets. We conducted tests by using motors with paraffin-based fuel and nitrous oxide in the blowdown mode. They were configured to run at fluxes of around 800 kg/(m2⋅s). Data from 13 hot-fire tests showed that pressure oscillations in the chamber did not exceed 0.5 bar, thus demonstrating that the DTI-configured hybrid motor could operate smoothly even at motor fluxes that were nearly four times higher than those of traditional hybrid motors. We also verified the capability of the DTI to improve the regression rate. The average regression rate of DTI-configured motors exceeded 6 mm/s, while classically configured hybrid motors achieved a rate of only about 1.5 mm/s at similar fluxes. The insights obtained from the tests can inform the design of hybrid motors that require a high thrust density.

Modern aeroengines require compact combustor/turbine integration with intensified cross-domain interactions. Current studies neglect turbine-induced reverse effects, contradicting physics. Conventional interface-based methods (sliding mesh/Multiple Reference Frame) enforce artificial pressure continuity at stationary/rotating interfaces, inducing numerical dissipation that attenuates reverse-propagating pressure waves and weakens turbine-to-combustor feedback. To address this, this study proposes an interface-free rotating framework with radiation modeling, capturing rotation-induced reverse-propagating pressure gradients through direct solution of global compressible flow equations. This approach simplifies mesh generation processes, achieving 16.7% computational time savings without compromising solution accuracy. The detailed analysis reveals that turbine structures constrain dilution-zone vortex development via adverse pressure gradients, driving hot streak expansion, and exacerbating temperature nonuniformity. Critically, rotation-induced pressure gradients propagate in reverse to the combustor, steepening existing pressure gradients and amplifying velocity fluctuation amplitudes. Low-speed fluctuations (less than or equal to 13,300 rpm) are dominated by swirl intensity; increasing speed attenuates swirl strength, progressively reducing fluctuations. At high speeds (greater than or equal to 19,950 rpm), turbine-induced reverse-propagating gradients constitute a dominant contribution, counteracting swirl decay. At this regime, the influence of this reverse effect traces back to the combustor front section. These findings provide critical guidance for suppressing hot streak diffusion and mitigating instability risks in combustor designs.

The rocket propulsion community is actively seeking nontoxic, stable propellants to replace hazardous hydrazine-based options. This study provides a comprehensive analysis of the hypergolic ignition and spray characteristics of a newly developed eco-friendly self-igniting fuel (named PAHyp 3) with rocket-grade hydrogen peroxide (RGHP), using an impinging jet apparatus. Using synchronized high-speed visible shadowgraph and infrared imaging, the transition from standard atomization to reactive bubbly atomization was revealed by modulating propellant reactivity. Infrared imaging aids in tracking spray breakup, mixing, and the liquid-to-gas transition of droplets, providing valuable insights. To better understand the mixing and reactive bubbly dynamics before ignition, a simplified oxidation mechanism is proposed based on infrared imaging of hypergolic and non-hypergolic samples. This work reveals key multiphase reaction behaviors, including bubble nucleation and growth, bubbly droplet burst (microexplosion), and gas-phase combustion dynamics, along with insights into the reaction kinetics. Unlike non-hypergolic propellants, hypergolic combinations exhibit unique droplet dynamics characterized by bubble nucleation, expansion, and rapid consumption via internal and external evaporation. These mechanisms result in accelerated droplet disintegration and a shorter lifespan. The findings provide insights applicable to other hydrogen-peroxide-based bipropellants as well as to conventional hydrazine-based systems.

Modern aeroengine axial turbine blades frequently encounter incidence angles that exceed 20 deg or even 60 deg in extreme cases. The parameters that exert the most significant impact on blade profile performance differ depending on the inflow angle. Hence, identifying the profile parameters with the greatest influence is crucial for optimizing performance within a desired inflow angle range. This study employs an orthogonal test design to systematically select parameter combinations for generating blade profiles in order to establish a universally applicable profile design method. Range analysis is then conducted on the generated profiles within the target inflow angle range to examine the effect of each profile parameter on the overall performance. Four parameters with the most significant impacts within different inflow angle ranges are identified. To understand the mechanisms by which these parameters influence the performance, an analysis is performed to examine the effects of profile parameter variations on local losses using the entropy generation rate, and the results confirm those of the range analysis. Finally, a profile parameter selection method based on sensitivity analysis, tailored to specific inflow angle ranges, is presented. The insights presented here can guide the design of turbine blades operating within wide-inflow-angle ranges.

An inertial particle separator (IPS) reduces sand and particulate ingestion in gas turbine engines by splitting the inlet airflow: a core leg ideally free of particles and a scavenge leg that carries contaminants. This study evaluates a novel IPS design with reduced flow turning using an axisymmetric-sector wind tunnel. Separation efficiency (η) was tested across three flow splits (β) and three standard dust types—ISO 12103-1 A2 Fine, AFRL-03, and ISO 12103-1 A4 Coarse—with a maximum uncertainty of 4%. For the ISO A4 Coarse dust, η trends matched those from prior large-scale systems. Efficiency increased with β for all dusts, with ISO A2 Fine showing a steeper η–β gradient than the others. Particle-size analysis revealed that lower-Stokes-number particles are harder to separate, especially at low β. The introduction of axisymmetry significantly affected performance, reducing η at low-flow splits compared to planar configurations. These results offer benchmark data for modeling and insight into how particle size and flow symmetry affect IPS performance.

Environmental regulations imposed stringent requirements, including the adoption of biofuels and the reduction of fuel consumption. As these measures are integrated, reliable engine operation, particularly for in-flight re-ignition and high-altitude ground ignition, must be ensured. Flame development and propagation in annular combustion chambers strongly depend on kernel characteristics generated by the igniter used and their interaction with a surrounding environment that is typically challenging to ignite. Therefore, a more detailed understanding of ignition kernel properties under adverse conditions is essential. The primary objective of the present study is to characterize ignition kernel properties using a real helicopter igniter, focusing on the characteristic times at the onset of the ignition phase during spark generation, a stage for which fine-scale data are currently lacking. Such information is crucial for accurately initializing numerical simulations dedicated to predicting ignition success or failure in the engine. To this end, several diagnostic techniques have been employed, including microcalorimetry, schlieren visualization, and filtered plasma chemiluminescence. The results indicate a decrease in energy transfer efficiency and an increase in final ignition kernel volume as initial pressure decreases. An existing simple model has been evaluated for its ability to predict the experimentally obtained results.

In this work, we investigate high-frequency thermoacoustic instabilities in a jet-stabilized gas turbine burner by means of large-eddy simulation (LES). This burner is operated at 8 bar pressure with a mixture of hydrogen and natural gas. Experimental data for the pressure oscillations in the combustion chamber are provided here to validate the simulations. The LES is conducted using a splitting scheme for solving the equations for compressible, reactive flows. To model the filtered chemical source terms in LES, two different models (namely an assumed probability density function (APDF) model and a thickened flame (TF) model) are used. Time-resolved computational data are analyzed using multiresolution proper orthogonal decomposition. It is found that mainly three longitudinal modes and one mixed transverse–longitudinal mode cause the high-frequency instabilities in the combustor. Computed frequencies of these modes agree excellently with measured frequencies. Amplitudes of the modes are best reproduced by the use of the TF model. Further analysis of the computational data obtained with the TF model reveals that heat release fluctuations caused by periodic distortion of the flame geometry and by fluctuations in equivalence ratio play an important role in the thermoacoustic feedback.

Liquid anhydrous ammonia reacts violently when placed in contact with white-fuming nitric acid (WFNA) but does not produce a flame. However, adding small amounts (∼5 wt.%) of lithium or sodium to ammonia to form a solvated electron solution makes ammonia hypergolic with WFNA. Adding small amounts of calcium (∼5 wt.%) to ammonia makes a mildly pyrophoric solution in air. The relative ignition delay between lithium–liquid-anhydrous-ammonia (LiLAA) and sodium–liquid-anhydrous-ammonia (NaLAA) solutions was experimentally determined. Kinetic calculations employing alkali metal and ammonia combustion mechanisms validated experimentally observed trends. The difference in ignition delay between lithium and sodium solutions was attributed to the strength of their respective -O and -OH bonds, which promote the formation of radicals and affect ignition delay. Thermodynamic calculations indicated that nonsolvated mixtures of lithium or calcium with ammonia increased specific impulse by several seconds in a WFNA-operated motor. However, mixtures with sodium decreased specific impulse by several seconds. Importantly, certain key thermodynamic properties of solvated electron solutions are still unknown, and combustion-related calculations are likely inaccurate. The use of these solutions in practical applications is currently very limited, as their metastable nature prevents feasible storage and would require modified engines that make the solution in situ.

This study experimentally investigates the effects of pilot injection on ignition performance for blends of conventional petroleum-derived aviation fuel (F-24) and an isobutanol-derived alcohol-to-jet (ATJ) fuel. The fuel blends are injected into a constant flow combustion chamber under various thermodynamic conditions to simulate the ignition process in compression-ignition (CI) engines. Simultaneous imaging of planar laser-induced fluorescence of formaldehyde, chemiluminescence, and schlieren are performed to analyze the detailed ignition process. Variations in fuel blend, ambient conditions, and injection strategies yield distinct ignition behaviors. Blends with higher ratios of ATJ and lower thermodynamic conditions reduce overall heat release and increase both first stage and overall ignition delays, occasionally leading to misfire. Compared to a single injection, the addition of a pilot injection to the main injection enhances overall heat release and reduces both first stage and overall ignition delays across all fuel blends and ambient conditions. This acceleration of the ignition process results from early formation of intermediate combustion radicals from the pilot mixture, elevating local temperature and promoting the chemical reaction of the main mixture. This effect is enhanced in blends with lower ATJ fractions, reducing the ignition delay up to 60% compared to single injection. These findings demonstrate that the pilot injection can expand the operational window for CI engines, especially with low cetane number fuels at lower temperature and pressure conditions.