The influence of low initial temperature on detonation characteristics is of significant importance for methane-fueled propulsion and energy systems operating under low initial conditions. In this study, detonation experiments were conducted using premixed methane-oxygen mixtures over an initial temperature range of 203-283 K. Detonation pressure and propagation velocity were measured to characterize the macroscopic behavior of detonation wave, while chemical kinetics analysis based on the Zeldovich-von Neumann-D & ouml;ring (ZND) structure was performed to elucidate the underlying mechanisms. The results indicate that decreasing the initial temperature significantly increases the peak pressure, while reducing fluctuations in propagation velocity, thereby enhancing both detonation intensity and propagation stability. ZND analysis reveals that both the induction and exothermic length decrease with decreasing temperature, resulting in a more compact reaction zone. This structural evolution strengthens the coupling between the shock wave and chemical energy release, contributing to the observed increase in detonation pressure. Sensitivity analysis further demonstrates that the sensitivity coefficients of the dominant elementary reactions increased as the initial temperature decreased. In addition, the production rates of key reactive radicals become faster and more spatially concentrated at low initial temperature conditions, thereby accelerating the overall energy release process. These findings suggest that the effects of low initial temperature on detonation behavior are governed by the coupled influences of chemical kinetics and reaction zone structure. The study provides a valuable insight into detonation propagation under low-temperature environment and offers practical guidance for the design and optimization of advanced detonation propulsion systems.

An experimental study was conducted on the combustion characteristics of liquid hypergolic propellants (monomethylhydrazine (MMH)-nitrogen tetroxide (NTO)) in an annular combustion chamber equipped with 62 pairs of impinging-jet injectors. The pressure oscillation behavior of tangential combustion instability and rotating detonation, as well as their correlation mechanisms, were systematically analyzed. The experimental results indicate that under combustion instability conditions, three competing pressure oscillation modes coexist within the combustion chamber: the 1T transverse mode instability, a high-frequency pressure oscillation band, and detonation-related pressure oscillation. The observed multi-mode hybrid characteristics of the pressure oscillations are not coincidental and are hypothesized to be related to the current combustion chamber configuration. Detonation phenomena were observed in two other operating conditions with the same configuration. The transition from deflagration to detonation exhibits a step-like growth characteristic, lasting approximately 15 oscillation cycles and taking about 3.2 ms. The amplitude of the rotating detonation oscillations reaches 133 %-224 % of the mean chamber pressure, with the wave propagation velocity attaining supersonic speeds. Significant differences in pressure oscillation amplitudes are observed across different cycles of the detonation wave, without showing clear regularity. The detonation-related pressure oscillations in unstable combustion may share a common physical origin with rotating detonation, or the rotating detonation may develop from the further evolution of such pressure oscillations. There is an essential distinction between rotating detonation using liquid MMH/NTO hypergolic propellants and combustion instability in rocket engines: combustion instability arises from the coupling between heat release and the inherent acoustic modes of the combustion chamber, whereas hypergolic propellant rotating detonation combustion is not influenced by the chamber's inherent acoustic characteristics.

To investigate the influence of hydrogen ratios (0%, 10%, 20%, and 100%) on detonation propagation, the detonation characteristics of multi-component mixtures (H2/CH4/O2) were studied at cryogenic (203.15-263.15 K) and ambient (283.15 K) temperatures using a combination of experiments and chemical kinetics analysis. The experimental results revealed a linear relationship between the mixture calorific value and hydrogen ratio, expressed as hmix = 161.28XH2 + 266.56(1 -XH2). A predictive model for peak pressure was developed, enabling the estimation of the peak detonation pressure as a function of initial pressure, initial temperature, mixture density, calorific value, and adiabatic index. Chemical kinetic analysis further revealed that cryogenic temperatures shortened the induction length in the ZND model. Furthermore, increasing the hydrogen ratio accelerated chain reactions by generating highly reactive radicals, and significantly altered elementary reaction pathways. These findings provide critical insights for the safety design and optimisation of hydrogen energy systems operating at cryogenic temperature.

Nuclear thermal propulsion, which uses a reactor core as the energy source of a nuclear thermal rocket, is expected to become an effective means of deep space exploration in the future. The reactor core can be damaged by a large temperature gradient. Thus, investigating the structural distribution of its internal components and understanding its flow and heat transfer characteristics is highly important. In this study, a 19-hole hollow hexagonal prism fuel element is selected for simulation. A new type of fuel element is proposed by changing the diameter of the channels in the work material, and the heat transfer characteristics are compared and analyzed. Compared with a conventional fuel element under uniform inlet conditions, when the inlet conditions and the diameter of the channel in the work material are changed, the peak temperature inside the fuel element decreases, but the overall temperature distribution is more uniform. Along the flow direction, the temperature distribution boundary is located at y = 300–500 mm. From the inlet to this position, the temperature distribution on the axial cross-section is uniform. From this position to the outlet, the temperature difference along the radial cross-section is significantly reduced, and the temperature fluctuation at the periphery of the fuel element is significantly improved. The research results can provide a reference for the design of fuel elements.

An in-depth study of fuel injection atomization in the expanded configuration under supersonic inflow conditions is extremely important for developing air-breathing rotating detonation engines. Combinations of H2-assisted transverse jets designed to enhance penetration depth of liquid and gas jets were investigated via Euler-Lagrange simulation method. The flow field structure and the enhancement mechanism of liquid jet atomization under the gas-liquid and liquid-gas combination modes were intensely studied. The hydrogen distribution and the atomization characteristics of the transverse liquid jet under other combinations were also summarized. Results show that H2 and liquid spray distribution produces inconsistent characteristics under different combination methods. Related aggregation phenomena of liquid spray can be corrected by changing the combination method. When the static temperature of incoming flow is high enough, the gas-liquid combination mode is recommended; otherwise, the liquid-gas combination mode based on active fuel such as H2 is recommended. The gas-liquid distance has little effect on the breakup of droplets. However, the larger the gas-liquid distance, the lower the penetration depth of the liquid jet. The same time, a prominent empty flow area appears on the lower wall surface. Results also show that the recommended gas-liquid spacing is 5 mm in the gas-liquid combination mode. The breakup of droplets under the gas-liquid combination is mainly caused by the shearing effect of the downwash flow field, and the liquid-gas combination is mainly due to the shearing effect of the upwash flow field.

This study employed an impinging-jet injector with 62 impinging pairs and an annular combustion chamber to conduct experiments on rotating detonation using liquid monomethylhydrazine (MMH) and nitrogen tetroxide (NTO). The experiments successfully achieved self-sustained rotating detonation waves spontaneously formed by hypergolic propellants, and observed detonation initiation, single-wave mode, and dual-wave mode. Furthermore, numerical simulations based on the Reynolds-averaged Navier-Stokes/large eddy simulation hybrid model and the discrete phase model were performed, successfully reproducing the hypergolic rotating detonation phenomenon. Experimental results revealed low-amplitude pressure oscillations in the combustion chamber prior to rotating detonation formation. The detonation waves in liquid MMH-NTO were found to develop from low-amplitude combustion instability, with the deflagration-to-detonation transition process exhibiting progressive characteristics. Numerical simulations revealed a transient deflagration process following the detonation wave. As the upstream droplets evaporated and absorbed heat, the temperature drop weakened the deflagration. Prior to the subsequent detonation wave arrival, the deflagration intensity attenuates to its minimum level, thereby storing energy to sustain stable detonation propagation while effectively suppressing the pre-wave deflagration characteristic. Energy release in hypergolic rotating detonation concentrates in the high-density droplet zone at the combustor head, exhibiting peak combustion intensity. The detonation wave triggers abrupt pressure and temperature surges that induce rapid propellant droplet evaporation and reactant compression. The compressed reactants undergo near-constant-volume combustion with fast energy release, driving supersonic detonation wave propagation.

PEG and AP are widely used in strategic and tactical missile engines as key components of composite propellants. It remains a challenge to investigate the detailed combustion mechanism of PEG/AP due to the complex structure and complicated chemical reactions. DME, N2O and Cl-2 are the main intermediates of PEG and AP pyrolysis, respectively, which play a crucial role in PEG/AP combustion. DME/N2O is also a promising combination propellant because of its high energy content and good combustion and environmental properties. This study systematically investigates the combustion characteristics of DME, N2O and Cl-2 mixtures based on experimental measurements. The Ignition Delay Times (IDT) of DME/N2O mixtures at equivalence ratios of 0.5, 1.0, and 2.0 (N2O as the oxidant) were measured using a high-pressure shock tube at pressures of 10.0 and 20.0 bar and in the temperature range of 1250-1600 K. Besides, half of the N2O was replaced by Cl-2 to investigate its impact on the ignition characteristics of DME/N2O. The result shows that although the addition of Cl-2 reduces the activity of the fuel mixture system, the ignition activation energy required for ignition has not changed. The laminar flame speeds of DME/N2O mixtures were measured by a constant-volume reactor. The equivalence ratios ranged from 0.8 to 1.4, with N-2 content controlled at 60 %, pressure at 1.0 bar, and initial temperature at 298/333 K. The experimental results were simulated using the NUIGMech1.3 model and a constructed model adding Cl-2 related reactions to NUIGMech1.3 in this study. Sensitive and flux analyses were conducted to determine the crucial reactions for the IDT of DME/N2O and DME/N2O/Cl-2. The results indicate that the decomposition of DME generates <(C)over dot>H-3 and <(C)over dot>H3O, which is the most reactivity promoting reaction at all temperatures, and it doesn't be influenced by Cl-2 presence. Meanwhile H-atom abstraction from DME by <(H)over dot> is the most reactivity inhibiting reaction, while it shows promoting effect with the Cl-2 addition, and the H-atom abstraction reaction by O-2, which did not show significant sensitivity before the addition of Cl-2, shows the strongest inhibitory effect at this time. H-atom abstraction reactions and C-O bond dissociation are two major pathways of DME primary consumption. Although the presence of Cl-2 did not alter this macroscopic phenomenon, it had a significant impact on the flux of each pathway. Meanwhile, the addition of Cl-2 directly changed the reaction after the third stage in the DME reaction pathways, making the reaction involving Cl-2 dominant at this time. The results in the current study should be a positive contribution to the development and optimization of detailed gas-phase chemical kinetic mechanisms for PEG/AP multicomponent solid propellant.

Unsymmetrical dimethylhydrazine (UDMH) is a widely used hypergolic rocket fuel. It is one of the most commonly used fuels for attitude control engines, such as those in missiles, satellites, spacecraft, and launch vehicles. We conducted a high-level theoretical study to develop a detailed combustion kinetic mechanism for UDMH, focusing on crucial elementary reactions. We analyzed 14 H atom abstraction reactions (HAA) of UDMH, four unimolecular reactions, and five chemical activation reactions on the potential energy surface (PES) of UDMH radicals (C2H7N2). Seven reactive radicals Ḣ, ȮH, NO2, HȮ2, ĊH3, CH3Ȯ, and CH3Ȯ2 were involved. The M06-2X/6-311++G(d,p) method, along with the CCSD/cc-pVXZ (where X = T, Q) methods, was used for theoretical calculations. The rate coefficients and thermochemical property parameters for the 23 elementary reactions and related key species were calculated. Our results indicate that HAA by Ḣ, ȮH, and CH3Ȯ dominate at low temperatures (300-700 K) compared to CH3Ȯ2, HȮ2, and NO2. Some of the calculated rate coefficients were compared with previous experimental and theoretical studies, and our computed values agreed well with reference data. For the reactions on the PES, the N-N breaking reaction and the CH3NNH + ĊH3 system show great importance in the initial reaction network. The pressure dependence of all the reactions on the PES was evaluated. Regarding the thermochemical properties of UDMH and its corresponding product radicals, our standard enthalpy values deviate by about 1 kcal·mol-1 from database values.

Through the method of three-dimensional numerical simulation,the rotating detonation combustion characteristics of kerosene-fueled scramjets in the range of Ma=3～7 were studied.In the flight condition of Ma=3,due to the poor effect of fuel atomization and evaporation,the detonation combustion of kerosene fuel cannot be realized.In the flight conditions of Ma=4,5,6,with the increase of Mach number,the number of wave heads increases gradually,which are single-wave,three-wave,and five-wave modes,respectively.However,the propagation speed gradually decreases.In the scramjet mode,the liquid fuel has a good effect of atomization and evaporation.Nevertheless,kerosene vapor remains in the flow field to va-rying degrees and is discharged from the combustion chamber without participating in the reaction.In the flight condition of Ma=7,the flow field will burn in a stationary detonation mode because the incoming flow is close to CJ velocity.

Thermoacoustic instability (TAI) presents a critical challenge for lean-burning combustors and rocket engines. The early detection of instability is crucial, and to address this, a data-driven prediction framework has been established for TAI in a sub-scale rocket combustor with variable chamber length. Nonlinear combustion features are generated from time series of dynamic pressure using recurrence matrices. Deep learning models are then utilized to train these features and predict the proximity of impending TAI. The performance of the proposed method is investigated through cross-validations of 12 groups of hot-fire test datasets. Remarkably, the prediction performances are in good agreement with measured experimental data, with most instabilities being predicted dozens of milliseconds in advance. This capability paves the way for the early implementation of active control systems in full-scale combustors in the future. The prediction performances are also compared with state-of-the-art TAI prediction methods.

Efficient mixing of liquid fuels and supersonic airflow is crucial for the successful detonation and stable propagation of rotating detonation engines (RDEs). In supersonic airflow, gas-assisted injection technology, using methane (CH4), effectively enhances the atomization of liquid jets. This study conducted numerical simulations to analyze the mixing process of wide-range Mach supersonic airflow and transverse liquid jet under the gas-assisted scheme to characterize the flow field structure, droplet breakup, and evaporation characteristics of preheated kerosene. Furthermore, combined with the mesh adaptive refinement technology, the effects of the incoming airflow Mach number, the number of injection positions for the auxiliary gas, and the gas auxiliary pressure drop were assessed. Specifically, the turbulent flow was simulated using the Improved Delayed Detached Eddy Simulation (IDDES) method. Additionally, the Discrete Phase Model was used to explore gas–liquid interactions. The results show: 1) Dual gas assistance can significantly improve the uniformity of droplet distribution under specific conditions. 2) The presence or absence of auxiliary gas at the liquid jet inlet notably influences droplet evaporation. 3) The presence or absence of auxiliary gas at the liquid jet inlet notably influences droplet evaporation. 4) A critical point in gas injection pressure drop at inlet 1 was identified, leading to peak total pressure loss at the combustor outlet. Additionally, the total outlet pressure loss observed under the low-speed incoming airflow condition is less than 10% of that obtained for the high-speed incoming airflow condition.

An experimental and numerical investigation of self-excited transverse combustion instabilities in a rectangular multi-injector, gaseous methane-oxygen rocket model combustor is presented. Numerical simulations based on flamelet-generated manifold and stress-blended eddy simulation are performed to reveal the dynamic features of the instabilities. Good agreement is achieved between numerical results and experimental data in the aspect of mean chamber pressure, characteristic frequency, and pressure waveforms. The nonlinear growth process (stable-rapid development-limit cycle) is clearly captured during the experiment, which is successfully predicted by numerical simulation. Nonlinear characteristics are analyzed via the phase space reconstruction method, and instantaneous flow fields are estimated to aid the interpretation of the combustion instability mechanisms. Results suggest that the linear characteristics dominate the initial growth stage of pressure oscillations. The nonlinear nature is significantly strengthened during the rapid exponential development process, indicated by the generation of harmonics. Both numerical and experimental phase space attractors vary from clutter to “Reuleaux triangle shaped” and ultimately to stretched trefoil-knot like structure, corresponding to the nonlinear growth process of combustion instabilities. The trefoil-knot like attractor is caused by the steep pressure wavefront (shock-like). Flow field analysis demonstrates that the gaseous propellants are highly disturbed by the transversely moving pressure wave. Large coherent structures of flame interactions and propellant gas stream collisions between adjacent injectors are generated, resulting in high deflection of propellants and heat release regions to both sides of the chamber. Coupling behavior of pressure oscillations between the main chamber and the oxidizer post (oxidizer manifold) is captured. A combustion instability mechanism associated with flame deflection and pulsated mass flow rate supply is proposed

The excellent atomization of the fuel in the air-breathing rotating detonation engine is extremely important to the detonation combustion. The gas-liquid combined transverse jet is one of the vital fuel injection technologies. Numerical methods were used to investigate the liquid penetration depth and auxiliary jet gas distribution. The Euler-Lagrange method was used to study the effect of different gas jets on the liquid jet. The results show that the gas jet reduces the momentum flux of the local incoming flow through the shock wave structure and increases the penetration depth of the liquid jet. Under the same incoming flow conditions, the greater the momentum flux of the gas jet, the higher the penetration depth of the liquid jet. It mainly affects the flow field area before the expansion section. The change of gas species has almost no effect on the penetration depth and atomization characteristics of liquid jet in the rear flow field.

The detonation propagation characteristics of the mixtures, 2H2 + O2+3Ar and CH4 + 2O2, were investigated. Accordingly, the mixtures were tested in round tube with inner diam-eter of D = 80 mm and annular tubes with widths of w = 25 mm, 15 mm, and 5 mm. The two mixtures represent stable mixture with regular cell pattern and unstable mixture with irregular cell pattern, respectively. Smoked foils were utilized to record cellular structure under various initial conditions. Subsequently, the length scale Ldsc was measured, which represents the length from the start of the test section to the position where the cellular structure changes drastically (the cell size obviously increases or the structure disappears). The results reveal that both mixtures can successfully propagate in round tube and annular tubes of 25 mm and 15 mm, but fail in 5 mm annular tube. The Ldsc value of 2H2 + O2 + 3Ar is higher than that of CH4 + 2O2 in 80 mm and 15 mm tubes, but it is opposite in 25 mm tube. Moreover, the relationship between Ldsc and hydraulic diameter DH was analysed. For a given tube, the values of Ldsc and Ldsc/DH increased when the initial pressure increased. And the variation trend of Ldsc and Ldsc/DH of CH4 + 2O2 is steeper. Furthermore, the mixtures 2H2 + O2 + 3Ar and CH4 + 2O2 resulted in over-driven deto-nation in 15 mm and 25 mm annular tubes, respectively. The ratio between the total re-action length (sum of the induction length and exothermic length) and the hydraulic diameter (DH/(&i + &e)) correspond to critical values of 18 for hydrogen-oxygen-argon and 6 for methane-oxygen, below which the detonation will fail. (c) 2022 Hydrogen Energy Publications LLC. Published by Elsevier Ltd. All rights reserved.

In this study, water and air were used as simulated media to investigate the spray characteristics of a gas-centered swirl coaxial (GCSC) injector under different ambient pressures through experimental, simulation, and theoretical approaches. The results showed that for the same liquid mass flow rate, the breakup length of the liquid film decreased with increasing ambient pressure, with the extent of the decrease diminishing as the ambient pressure increased. In addition, it was found that under the same ambient pressure, the larger the liquid mass flow rate, the shorter the liquid film breakup length. Moreover, for any gas–liquid mass flow ratio (GLR), the spray angle increased significantly with ambient pressure, and the liquid film suddenly expanded compared to the no-ambient pressure condition. At the same time, the breakup length of the liquid film increased. As ambient pressure increased, the spray angle increased while the breakup length decreased. When the GLR was small, self-pulsation occurred, which gradually disappeared with increasing ambient pressure. The frequency of self-pulsation decreased with increasing ambient pressure. The periodic self-pulsation of the spray was the result of the combined effects of the centrifugal force of the rotating liquid film, surface tension of the liquid film, aerodynamic force, and pressure difference inside and outside the liquid film. Overall, these findings provide insight into the spray characteristics of GCSC injectors under different ambient pressures, which can be of importance to the design and optimization of liquid rocket engines.

Conducting an in-depth study of the characteristics of a two-phase rotating detonation combustor is crucial for developing a new type of air-breathing engine that can be used for wide-range flight. This study established the first full-process air-breathing two-phase rotating detonation numerical combustor, based on published experimental data, using the Euler-Lagrangian method to simulate the complete process of fuel injection, breaking, evaporation, blending, and combustion. The simulation model analyzed the cold flow and the combustion flow field after detonation ignition, providing insight into the interaction mechanism between the two-phase detonation wave and fuel. The simulation results showed that the secondary heat release after the detonation wave was caused by the secondary combustion of unburned droplets. The working frequency of the detonation combustor was found to be 2301 Hz, and the propagation speed was 1228.89 m/s. Compared to the experiment, the simulation showed a deviation of approximately 19.5% in terms of frequency and detonation wave propagation velocity. When the Sauter-mean diameter of the droplets D-32 was around 25 mu m, the combustor was able to operate stably. The thrust generated by the engine was approximately 1009 N, and the fuel-specific impulse was about 1767.5 s. These findings have significant theoretical value in designing a two-phase air-breathing rotating detonation combustor and engine.

Abstract Purpose The purpose of this study was to examine the effects of six weeks of routine use of a novel robotic transfer device, the AgileLife Patient Transfer System, on mobility-related health outcomes, task demand, and satisfaction relative to previous transfer methods. Materials and Methods Six end users and five caregivers used the system in their homes for six weeks. Participants completed several surveys examining perceived demands related to preparing and performing a transfer and mobility-related health outcomes pre and post intervention. Participants were also asked about their satisfaction with using the technology compared to previous transfer methods. Results Both end users and caregivers reported reduction in perceived physical demand (p = 0.007) and work (p ≤ 0.038) when preparing for and performing a transfer. End users indicated that the device intervention had a positive impact, indicating some improvements to health-related quality of life as well as improved competence, adaptability, and self-esteem post-intervention. All participants were highly likely to recommend the technology to others. Conclusion The AgileLife Patient Transfer System is a promising new form of transfer technology that may improve the mobility and mobility-related health of individuals with disabilities and their caregivers in home settings. Implications for rehabilitation Robotic transfer assistance reduced physical demand and work among end users and caregivers. The robotic device had a positive impact on some quality of life outcomes after 6 weeks of use. Users were highly likely to recommend the robotic transfer device to others.

This study conducted 3D numerical simulations to investigate the impact of combustion chamber diameter on the combustion characteristics of liquid kerosene scramjet rotating detonation under high Mach number flow conditions (Ma6/28 km). After ignition, a local hotspot near the contact surface between the combustion product and fresh reactant facilitates the generation of new detonation wave heads, resulting in a co-directional multi-wave mode in the detonation combustion flow field. The higher total temperature of the incoming flow restricts the accumulation of a substantial fuel gas layer in the axial direction of the combustion chamber, resulting in a smaller fuel distribution area and a lower wave head. Increasing the inner diameter of the combustion chamber leads to an increase in the number of wave heads but a decrease in overall height. Specifically, when using diameters of 125 and 150 mm, we observed significant periodic low-frequency oscillations in the peak pressure of the detonation wave during stable propagation. The specific impulse of the fuel does not vary significantly across different combustion chamber diameters. However, when the inner diameter is 75 mm, periodic oscillations occur, which reduce thrust stability. These findings provide valuable insights into optimizing combustion chamber design and improving the efficiency and stability of liquid kerosene scramjet rotating detonation propulsion systems under high Mach number flow conditions.