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Al–Li alloys are considered promising materials in solid propellants owing to their higher combustion efficiency than Al particles, which exhibit incomplete combustion. However, Al–Li alloy particles cannot be directly applied in propellant systems owing to their poor compatibility. Thus, in this work, an in situ polymerized film was formed via the pretreatment of Al–Li alloy followed by polymerization with trifluoroethyl methacrylate (TFEMA) and N,N′-methylene-bis-acrylamide (MBA). The hydrothermal stability of the coated samples was improved with a mass variation rate of 5.82% on treatment with TFEMA (30%) and MBA (10%), corresponding to 116.62% of Al–Li at 60 °C and 75% humidity for 30 days. Besides, the thermostability increased based on the 19° backward skewing of the high temperature exothermic peak. Meanwhile, there were no pores and cracks in the HTPB propellant, which proved the possibility of its practical application with the compatibility of level 1 with HTPB in solid rocket motors. In addition, the Al–Li propellant exhibited a better combustion performance of 0.53 s and 1309.33 °C than the Q3 Al propellant, with 1.34 s ignition delay time and 990.24 °C burning temperature. The combustion efficiency was also improved based on the observation of combustion phenomena and residue size. Besides, the combustion mechanism was analyzed based on the characteristics of Al–Li with the micro-explosion effect. Therefore, Al–Li alloy has the potential to be used as a new generation solid propellant. [ABSTRACT FROM AUTHOR]

When a solid rocket engine is ignited, the throat lining of the nozzle is prone to chemical ablation owing to high-temperature gas erosion, resulting in thrust loss. In this paper, a coupled fluid–solid model for thermochemical ablation on the nozzle wall is established based on the multi-component Navier–Stokes equations, SST k-ω turbulence model, finite-rate chemical reaction model on the nozzle wall, variable transport properties of the nozzle material, and the heat conduction equation. Compared with the experimental data, the maximum error of the calculated ablation rate was 4.37%, validating the effectiveness of the model. Subsequently, the effects of different combustion chamber components, pressures, and temperatures on the ablation rate of the carbon–carbon (C/C) throat lining were studied. The results indicate that the temperature at the nozzle throat was the highest, resulting in the maximum ablation rate. As the Al mass fraction at the nozzle inlet increased, the thermochemical ablation rate of the nozzle decreased with a lower oxidizer mass fraction. The inlet pressure and temperature of the nozzle were positively correlated with the ablation rate, with the temperature having a more significant impact than the pressure. These findings provide theoretical guidance for the thermal protection design of rocket engine nozzles. [ABSTRACT FROM AUTHOR]

Numerical and experimental approaches are conducted to investigate the flow features and secondary combustion performance induced by different air–fuel ratios in a boron-based solid-ducted rocket engine. The results indicated that the afterburning chamber flow features become more complicated owing to the multiple nozzles of the gas injector, and a number of recirculation zones are generated. Because of this, the mixing of the fuel gas and incoming air is enhanced. When the air–fuel ratio decreases, the heat release in the afterburning chamber increases continuously, which causes the pre-combustion shock train to continue to propagate upstream in the subsonic diffuser of the inlet isolator, along with the boundary layer separation zone also moving forward, and the stability margin of the direct-connect inlet decreasing gradually. Furthermore, the direct-connect inlet works at a critical state with an air–fuel ratio of 11.5. As the mass flow rate of the fuel-rich gas rises gradually, the engine thrust gradually increases, and the number of vortexes in the afterburning chamber and the corresponding region affected by the vortexes generally decrease. Meanwhile, the mixing and combustion of the fuel-rich gas and incoming flow were not substantially changed. Additionally, the combustion efficiency and specific impulse are proportional to the air fuel ratio. [ABSTRACT FROM AUTHOR]

This paper aims to validate the design contours and result parameters of the newly developed Dual Expansion Nozzle (DEN) using the Fastran software when it enters the non-adaptation regime with the increase in altitude and variation of the Nozzle Pressure Ratio (NPR). DEN design and Computational Fluid Dynamic (CFD) application utilize the High Temperature (HT) model defined by a calorically imperfect and thermally perfect gas, providing good accuracy compared to PG model. The computational domain is decomposed into subdivisions of structured grids using the algebraic grid generator software CFD-GEOM. The mesh convergence was analyzed using three mesh resolutions: coarse, medium, and fine. A comparison was made between the two conventional nozzles, MLN and the newly developed BPN, as both are currently used in aerospace propulsion to enhance aerodynamic performance. If a rocket engine operates under strongly over-expanded conditions with ambient pressure considerably higher than the nozzle exit pressure, and by increasing NPR value, the flow separates from the wall and can lead to high side loads. These loads are reduced in DEN. The flow separation point is observed close to the exit section for the DEN in comparison with Minimum Length Nozzle (MLN) and Best Performance Nozzle (BPN) for the same NPR , resulting in low flow separation and low side load effects for DEN. This result demonstrates a small loss in the exit Mach number, leading to less loss in the thrust coefficient and a small effect of boundary layer friction for DEN. Enhanced over-expanded aerodynamics is observed in DEN compared to MLN and BPN. The application is made for air with an exit Mach number equal to 3.00. [ABSTRACT FROM AUTHOR]

This paper explores the aerodynamic behavior of a reusable launch vehicle, the Skylon spaceplane, for different angles of attack at Mach 5. The goal is to determine how aerodynamic effects manifest at various angles of attack and compare it with the theoretical data along with the determination of total time required for such computational fluid dynamics (CFD) simulations. The analysis focuses on the aerodynamics of the Skylon spaceplane and the Synergistic Air-Breathing Rocket Engine to understand their impact on single-stage reusable vehicles in the aerospace sector. This work is also crucial for paving the way for CFD simulations in future research in this sector. The total time consumed by the simulation and the possibility of using its data for other less time-consuming methods, such as convolutional neural networks, are considered. This research establishes a foundation for understanding the aerodynamic effects of specific angles of attack by comparing theoretical and simulation values. Results are produced for angles of attack of −60°, −45°, −30°, −15°, 0°, +15°, +30°, +45°, and +60° at Mach 5. The lift coefficient CL and drag coefficient CD are analyzed using the shear-stress transport k–ω model, and a comparison is made between the theoretical and simulated values. Finally, errors are calculated to understand the variation between the computed and theoretical values, as experimental data comparison is expensive and experimental data are often confidential. [ABSTRACT FROM AUTHOR]

The prediction of backflow from multi-species high density rocket engine plume at high altitude, i.e., plume gases going upstream of the vehicle in rarefied atmospheric conditions remains a challenging numerical problem. Direct Simulation Monte Carlo computations are used to assess the sensitivity of backflow to plume and atmosphere inflow properties, and ultimately to derive a semi-analytical backflow model. It is found that backflow is independent of plume density in the thermal equilibrium limit at the nozzle exit plane, which allows for huge computational cost reductions in simulations to determine the backflow model parameters. The backflow behavior also appears to be dependent on two Knudsen numbers, representing the density effect in the far field atmosphere and in the compressed region in front of the vehicle. The ability of the model to estimate the backflow of specific species is finally demonstrated on a meter sized solid rocket case. [ABSTRACT FROM AUTHOR]

This study investigates the evolution of spray characteristics in adjacent gas-centered swirl coaxial (GCSC) injectors, which finds application in liquid propellant rocket engines. The main objectives here are to measure the axial evolution of droplet characteristics in the spray interaction zone and understand the fundamental physics governing the spray interaction process. Experiments were conducted using air and water as the working fluids under atmospheric conditions. Utilizing the high-speed shadow imaging technique, the droplet images were captured at different axial and radial measurement stations for gas-to-liquid momentum flux ratio (M) ranging from 30 to 70. The images were processed to obtain droplet size, axial/radial components of droplet velocity, and droplet mass flux. The Mie-scattering images of the spray were acquired by laser sheet imaging to visualize the spray structure and spatial distribution of the droplets. Droplet measurements were also obtained by operating the injectors individually. Comparative analysis between the interacting and individual sprays highlighted the significant reduction in characteristic droplet size and an increase in the mean droplet velocity and local mass flux due to spray interaction. To elucidate the physical mechanisms behind the above observations, further analysis was carried out by evaluating the droplet collision, secondary atomization, and droplet dispersion in the interaction zone. Interestingly, the results highlight that, despite the intuitive notion that droplet collisions are the primary driver of the spray interaction process in the intersecting sprays, the improved secondary droplet atomization due to modification of airflow characteristics serves as the dominant factor in altering the droplet characteristics. [ABSTRACT FROM AUTHOR]

In response to the inherent limitations of traditional thrust vector measurement methods for small solid rocket motors(SRM), this paper proposes an innovative method based on flight attitude inversion. Flight measurement tests were conducted. Firstly, a six-degree-of-freedom (6-DOF) data inversion algorithm for solving the dynamic data of the three-axis thrust vector was constructed based on the Euler transformation method and the 6-DOF flight motion mathematical model of the flight measurement platform under the action of a three-axis thrust vector. Then, an unconstrained experimental platform was established, obtaining 6-DOF data during the operation of the experimental SRM through an inertial measurement unit(IMU). Finally, the thrust vector data were obtained from the recorded data. During the stable working phase, the average main thrust was 845 N, and the average thrust eccentricity angle was 0.018°, with a main thrust error of 0.6 %, thereby validating the feasibility of the measurement method. [ABSTRACT FROM AUTHOR]

Currently, there is considerable emphasis on the electric pump-fed cycle for liquid engine, primarily due to its design simplicity. However, its development is hindered by the underdeveloped state of power battery technology. Drawing inspiration from hybrid power technology used in electric vehicles and turbochargers, a hybrid pump feed system for throttleable engines is originally proposed as a promising solution. This system integrates the electric motor into the gas generator cycle, with several topologies evaluated. The parallel configuration featuring a mid-motor is selected for its compact structure, efficient power-splitting and energy recovery. Additionally, customized energy management strategies and optimization models are developed to effectively allocate power throughout the operational processes of liquid engines. A comparative analysis of four engine cycles is conducted under the typically variable-thrust mission. The results indicate that attributed to the conservation of turbo-gas and battery energy, the optimized hybrid pump achieves a reduction of 2.39 % compared to the turbopump and 7.15 % to the electric pump in total mass. Adaptability assessment further indicates that the mass advantage of the hybrid pump system is more significant during prolonged engine burning and deep throttling. Specific working conditions are found in which the system prefers electric-motor driving or regenerating turbine energy. Although energy-recovery results in the system efficiency decrease, it serves to lower energy demand of battery pack, thus easing the burden on cell thermal management and structural design. This study provides a practical design framework for hybrid pump-fed rocket engines in future variable-thrust missions. • The practical scheme of hybrid pump feed system for liquid engine is originally proposed. • Energy management strategies with respect to throttleable engine mission profile is designed. • A comparison is conducted between the hybrid pump, electric pump, and gas generation cycle. • Long-term firing and prolonged, deep throttling of thrust conditions are effective for the hybrid pump system. [ABSTRACT FROM AUTHOR]

Equivalent ratio (ER) is an important factor affecting detonation characteristics and propulsion performance of rotating detonation rocket engine (RDRE). In this paper, the effects of different equivalent ratios detonation characteristics and thrust performance of methane-oxygen RDRE were studied by 2D numerical simulation. The premixed reactants were injected through the injection holes to simulate the discrete injection of reactants on the injection panel in actual RDRE, the number of injection holes was 60 and 120. The results show that there is hybrid detonation mode (HDM), co-direction multi-wave detonation mode (CMM) and unstable detonation mode (UDM) in detonation combustion due to the influence of equivalent ratio and the number of injection holes, and the co-directional multi-wave detonation mode is beneficial to the thrust stability of RDRE. At the last, the number of detonation waves in RDRE decreases with the increase in the equivalent ratio, and the specific impulse (Isp) increases with the increase of the equivalent ratio. [ABSTRACT FROM AUTHOR]

In this paper, we construct conditional reliability in the context of stress-strength models, taking into account the relationship between applied stress and strength. The Bernstein copula approximation is used to calculate the conditional reliability values. Real data from a rocket motor experiment are considered to evaluate the conditional reliability strength of a motor case and the required stress endurance for the motor. [ABSTRACT FROM AUTHOR]

The work carried out numerical studies of the dynamics of iodine plasma formation in an inductively coupled plasma (ICP) discharge, which is the working chamber of modern plasma engines. Calculations were carried out for various values of input power. It is shown that at short times an ion-ion plasma is formed with dominant particles and , and at times from several fractions to a few milliseconds there is a transition from ion-ion to electron-ion plasma with a dominant atomic iodine ion . [ABSTRACT FROM AUTHOR]

A method for analyzing liquid ligaments of a textural atomization process is presented in this article for the case of a rocket engine type-assisted atomization process under combustion. The operating point positions the atomization process in the fiber-type regime carrying an intense textural atomization process. Multiscale in nature, the method based on image analysis associates a scale distribution with a family of ligaments, this distribution being sensitive to the number, size and shape of these ligaments. The quality of scale distributions measured by image analysis depends on the spatial resolution and the precision of area measurements of surfaces with curved boundaries but described by square pixels. Part of the work consisted of improving the method for measuring scale distributions by using a sub-pixel image analysis technique and refining the surface area measurement method. Another part directed the multiscale analysis toward the estimation of the diameter distributions of the blobs that characterize the large-scale deformation of the ligaments. The analysis describes the atomization process at a level of detail never reached. For instance, assuming that the blobs are drops in formation, the estimated diameter distribution (bimodal in the case examined here) and the number of these drops are evaluated as a function of the distance from the injector. This information indicates where the process is most intense and where it stops. Furthermore, these diameter distributions receive a mathematical expression whose parameters report clear evolutions with the distance from the injector. This shows the possibility of elaborating mathematical models appropriate for textural atomization mechanisms. [ABSTRACT FROM AUTHOR]

The collision characteristics between high-temperature alumina droplets and the char layer in solid rocket motors are of great significance for the accuracy of slag deposition and flow-field simulations, however, the current research on the collision characteristics of the alumina droplets and char layer is still in a blank state. This study is based on the high-temperature molding method to prepare the char layer and compare the porosity with that of the char layer in solid rocket motors, indicating that the two are relatively similar in structure and can be applied to droplet impact experiments. An experimental study on the collision of alumina droplet with the char layer was conducted using a high-temperature alumina droplet impact experimental system. The experimental results show that the adhesion behavior of alumina droplets is related to the rough structure of the char layer and the high viscosity dissipation of the process of droplets impacting the char layer, and the droplets adhere during the retraction stage with violent oscillation. The rebound behavior of the droplets on the wall was characterized by "tail dragging", "spinning" and "asymmetric rebound" phenomena due to the combination of high surface tension and the pinning effect of wall roughness. A regime map of the rebound/adhesion results of droplets impact the char layers was constructed. At the same speed, droplets with smaller particle sizes are more likely to adhere to the char layer. We established a relationship between the rebound and adhesion behavior. Based on the experimental results, the relationship for the maximum spreading factor of the droplets was established, providing a theoretical basis for the in-depth understanding and study of the droplet collision process in solid rocket motors. • Adhesion behavior of droplets linked to the rough structure and viscous dissipation. • Rebound behavior of droplet occurs with different phenomenas like "trailing". • The relationship for the maximum spreading factor of the droplets was established. • A regime map of rebound/adhesion of droplets on char layers was constructed. [ABSTRACT FROM AUTHOR]

The potential of the Al-F reaction in suppressing agglomeration during propellant combustion and enhancing combustion performance is investigated by introducing fluorinated graphene as a fluorinated oxidizer. Comparative analyses of ignition combustion and agglomeration behaviors are conducted on novel composite powders and propellant samples modified with varying contents of fluorinated graphene using laser and hot wire ignition visualization systems. Characterizing parameters such as characteristic spectra, flame grayscale, ignition delay time, combustion duration, and burning rate are measured during combustion at different pressures. Additionally, agglomerated particles are collected via quenching techniques under 7 MPa pressure to explore the influence mechanism of fluorinated graphene on agglomeration near the burning surface, and a comprehensive influence mechanism is proposed. Results indicate that fluorinated graphene promotes ammonium perchlorate decomposition, accelerates oxidizing gas release, and enhances thermal conduction at the burning surface. The reaction between Al and F decreases the formation of intermediates (AlO and Al 2 O), while the interaction of F with Al and Al 2 O 3 effectively inhibits the clustering of Al particles, replacing conventional oxidation reactions and resulting in a unique micro-explosion jetting phenomenon. The introduction of 15 % fluorinated graphene concentrates most product particles around 10 μm, enhancing energy release during combustion. Overall, this composite powder containing fluorinated graphene effectively improves the combustion performance of aluminum-containing composite propellants, inhibiting Al particle agglomeration and potentially reducing specific impulse loss in solid rocket motors. • Fluorinated graphene promotes advanced decomposition and combustion. • Composite particles increase the combustion intensity of Al-based propellants. • F-O competition reaction suppresses the formation of agglomerations. • Multi-factor induced micro-explosions in agglomerations. [ABSTRACT FROM AUTHOR]

The study of the transient combustion modes is one of the key topics when considering the safety of space flights. Control of detonation onset has a dual application. First, the search for ways to prevent detonation modes in case of accidental fuel releases for fire safety issues of launch systems and the avoidance of accidents with rocket engines at a launch site and in near-Earth space. Second, the study of detonation and the possibility of using it to create propulsion systems based on detonation combustion of fuel. The paper shows the effect of the presence of spatial non-uniformities on the promotion of detonation in the chamber. Various geometries with and without obstacles and cavities are considered. It is demonstrated that the presence of obstacles accelerates the transition to the detonation process on the one hand, but on the other hand the presence of obstacles in combustion chamber could be the cause of incidental uncontrolled ignition, which ruins stable operation of an engine. The results of theoretical studies of the working cycle of the combustion chamber of a pulsed detonation engine are presented. Theoretical estimates for thrust characteristics are obtained. • The possibility of using detonation for developing new type of Space engine is demonstrated. • Different configuration of obstacles could both promote or inhibit the detonation onset. • Large distance between obstacles promotes uncontrolled self-ignition between engine cycles. • For each type of obstacles there exist optimal pitch for stable detonation onset. [ABSTRACT FROM AUTHOR]

The key problems of safety for Space missions begin with safety, reliability and effectiveness of rocket engines of different types used at different launch stages and orbit corrections. Today, the possibilities for improving chemical rocket engines of traditional types are almost completely exhausted and are limited to minor improvements in energy-mass characteristics. A qualitative leap in the development of engine building can only be achieved through the development and implementation of new types of engines. As unburned fuel in the combustion chamber is a loss of thrust for the engine, the study of droplet combustion and evaporation, in particular, the droplet lifetime, is of fundamental importance in the creation of combustion chambers using atomized liquid fuel in their operation. In this paper a quasi-stationary model, which describes the evaporation of a single droplet in a gaseous atmosphere, is presented. Since in the numerical implementation the mass flow from the liquid phase to the gas and the heat flux from the droplet to the gas are calculated based on the Peclet number and the droplet surface temperature obtained from the quasi-stationary problem, approximation formulas for these parameters are developed in this paper. As an example, the problems of evaporation of a liquid oxygen droplet in an atmosphere of gaseous hydrogen and a droplet of liquid n-decane in an atmosphere of gaseous oxygen are considered. Formulas for calculation of mass flow and heat flux from liquid phase to gas based on the solution of the droplet evaporation problem are presented. Estimates of droplets lifetime in engine are provided based on developed droplet evaporation models. • Quasi-stationary model of a single droplet evaporation in a gaseous atmosphere is presented. • Approximation formulas for Peclet number and the droplet surface temperature are developed. • Formulas for calculation of mass flow and heat flux based on the solution of the droplet evaporation problem are obtained. • Problem of a liquid oxygen droplet evaporation in an atmosphere of gaseous hydrogen is studied. • Problem of n-decane droplet evaporation in an atmosphere of gaseous oxygen is regarded. [ABSTRACT FROM AUTHOR]

The principles and methods of improvement of thrust, efficiency and cooling of rocket engines for space flight safety are of great interest for aerospace industry. The development of reliable and durable low-thrust rocket engines for space missions seems to be one of the important areas of development of the rocket and space industry. Issues related to the implementation of a new direction related to the design of propulsion systems for orientation systems of upper stages of launch vehicles using environmentally friendly fuel components are considered. A mathematical model is developed and numerical modelling of processes in the combustion chamber of a low-thrust rocket engine with gaseous fuel components (hydrogen and oxygen) is carried out. The model takes into account the presence of a cooling jacket. The mathematical model includes the effects of turbulence, combustion of a gaseous fuel mixture, kinetics of hydrogen combustion, and radiation. As a result of calculations, distributions of flow quantities in the combustion chamber and thermal loads on its walls are obtained. The influence of hydrogen mass flow rate on the efficiency of the working process and the dependence of thrust on hydrogen mass flow rate are discussed. The results of numerical modelling are compared with the data of a physical experiment obtained through fire tests of an engine model made of stainless steel on a three-dimensional printer. • Processes in combustion chamber of low-thrust rocket engine are analysed. • Model includes effects of turbulence, hydrogen combustion, radiation, cooling. • Distributions of flow quantities in combustion chamber are obtained. • Influence of hydrogen mass flow rate on engine thrust is discussed. • Results are compared with data of fire tests. [ABSTRACT FROM AUTHOR]

The oxidizer system in a hybrid rocket motor needs to deliver the flow from a pressurized storage tank to multiple combustor ports. Pressure losses in the oxidizer system directly impacts combustor pressure and consequently the vehicle performance. However, oxidizer feed line designs till date have been done using simple 1D tools. Higher fidelity flow analysis methods have not been reported in the literature to identify loss generating features. Therefore, a design improvement study was carried out to identify and alleviate the impact of undesirable flow features in a typical oxidizer system design. An experimentally calibrated 3D RANS approach is applied to a typical LOx feed system which includes steps, splitters, ports, and pipes with multiple bends. These design features result in varying degrees of flow separation, secondary flows and vortical flow features and result in total pressure losses of up to 7 %. This loss means that the storage tank needs to be pressurized further to accommodate such losses and ensure combustor performance. A targeted design improvement approach that features simple, alternative, implementable solutions in the loss-generating regions is discussed. The best of these design improvements can reduce the total pressure loss to 4 %, indicating a 43 % reduction in the losses and reduced impact on storage tank design and combustor performance. Therefore, this paper demonstrates that a higher fidelity design enhancement process of the oxidizer feed system, which is often neglected in such detailed studies, can result in overall vehicle level design improvements to ensure mission targets are met effectively. • Experimentally calibrated flow analyses result in lower pressure loss oxidizer feed systems. • Diagnostic identification of pressure loss generating features in a typical flight geometry. • Targeted design treatments for traditional oxidizer feed line design elements. • Bespoke, practical, implementable solutions for improving hybrid rocket performance. • Zonal design modification approach led to pressure loss improvements of up to 4 %. [ABSTRACT FROM AUTHOR]

The benefits of using hydrogen as a fuel for rocket engines dictate the necessity of a deep understanding of possible scenarios of hydrogen ignition and subsequent combustion. Safety requirements on launch sites and hydrogen transportation and storage facilities should be elaborated based on a thorough investigation of hydrogen flame development. The paper aims to provide a detailed numerical analysis of the hot wall ignition of hydrogen-air mixtures with compositions near low flammability limits. It is shown that the development of ultra-lean flames initiated by a hot spot on the solid wall possesses several features compared to ultra-lean flames ignited by a point energy source. Three modes of flame development are observed: stable flame column with bow-shaped flame structure on top, individual flame kernel, and unstable flame column with multiple kernels generation. Effects of mixture composition, along with the impact of hot spot size, are analyzed. Column tip acceleration mechanism is determined. All of the combustion modes obtained are grouped in the diagram, providing information on the limits of various combustion modes depending on the mixture concentration and hot spot size. • Ultra-lean hydrogen-air flame initiated by a hot spot on a wall is analyzed. • Effects of mixture composition and size of igniting spot are evaluated. • Individual flame kernel, stable flame column and puffing modes are observed. • Mechanism of flame acceleration in flame column mode is identified. [ABSTRACT FROM AUTHOR]

The combustion performance of solid propellants is significantly affected by the external operating parameters of rocket motors. In this paper, the combustion response characteristics of aluminized solid propellant under acceleration conditions have been studied theoretically and experimentally. A theoretical model of unsteady burning for aluminized solid propellant, considering transient heat transfer and acceleration effects, is developed based on the Zel'dovich-Novozhilov solid-phase energy conservation. The pressure-coupled responses of the solid propellant are also measured experimentally under different normal accelerations using the T-burner and centrifuge. The model is well validated by comparison with literature and experimental results. The unsteady burning of propellants is analyzed in detail, and the effects of acceleration and propellant properties on combustion response are investigated. The results show that there is a significant difference between the transient and quasi-steady burning rate under dynamic environments. As normal acceleration increases from 0 g to 1000 g , the fluctuation amplitude of the net heat flux on the burning surface continues to decrease, the magnitude of the pressure-coupled response peak decreases significantly by 69 %, and the frequency of the response peak shifts from 140 Hz to 230 Hz. Moreover, the pressure exponent and thermal conductivity plays a positive role in the pressure-coupled response. The model presented in this paper can be helpful for predicting the differences between flight and ground test performance and evaluating the combustion stability of solid rocket motors. • An integrated theoretical model of combustion response for aluminized propellant under acceleration effects is established. • A comprehensive experimental device is designed to measure the pressure-coupled response under overload conditions. • The effects of acceleration on transient burning characteristics are deeply analyzed under different dynamic conditions. [ABSTRACT FROM AUTHOR]

Thermoacoustic instability is a persistent problem frequently observed in various types of combustors, resulting in damaging consequences. However, our understanding of the dynamics in industrial combustors undergoing thermoacoustic instability, particularly in solid rocket motors, still remains limited. Data-driven precursors for thermoacoustic instability in such systems are also unknown. In this study, we use recurrence network measures and spectral entropy to characterize the dynamics of pressure data obtained from a full-scale solid rocket motor transitioning to thermoacoustic instability and design data-driven precursors for thermoacoustic instability. We show the scale-free nature of combustion noise and that the dynamical transition from combustion noise to thermoacoustic instability can be detected using two complex network measures: the average path length and average betweenness centrality. We calculate the spectral entropy in the frequency domain and find it more sensitive to detecting the dynamical transition and computationally cheap, which is promising for flexible use as a new precursor in thermoacoustic instability prediction. Our work highlights the feasibility of employing complex network measures and spectral entropy for precursors in solid rocket motors, paving a new path for using data-driven measures to early warning of thermoacoustic instability in solid rocket motors. [ABSTRACT FROM AUTHOR]

Copyright of Journal of Harbin Institute of Technology. Social Sciences Edition / Haerbin Gongye Daxue Xuebao. Shehui Kexue Ban is the property of Harbin Institute of Technology and its content may not be copied or emailed to multiple sites or posted to a listserv without the copyright holder's express written permission. However, users may print, download, or email articles for individual use. This abstract may be abridged. No warranty is given about the accuracy of the copy. Users should refer to the original published version of the material for the full abstract. (Copyright applies to all Abstracts.)

Mitigation of combustion instability of solid propellant rocket motors is ever being attempted by several researchers as the problem is very complicated in nature. The acoustic admittance of composite solid propellants (AP/HTPB/RDX/Al) is experimentally investigated using Impedance tube technique. It is one of the indirect experimental techniques to determine the combustion response of the solid propellants besides T-burner. The governing equations based on the conservation of mass, momentum and energy (reactive flow is simplified to have only the axial temperature gradient) are linearized and solved to determine the acoustic admittance of the burning propellant. Different methods of analysis were explored to validate/compare the results obtained and are matching well. Some of the unreported details of the analysis are brought out. A novel way of utilizing the outer tube besides the inner tube as against the literature (wherein the inner tube is kept inside the outer tube) to maintain the chamber pressure is explored. A test conducted at fundamental acoustic mode (110 Hz) at 2 MPa at room temperature (303 K) is considered for the analysis. A rotary valve capable of operating up to a mean chamber pressure of 12 MPa is developed in this work to act as an acoustic driver for the experiments. The advantage of the impedance tube technique over T-burner is that the former is capable of measuring both the real and imaginary part of combustion response whereas the latter can only measure the real part. The imaginary part of the combustion response besides the real part is useful to predict the stability characteristics of the solid propellant rocket motor. [ABSTRACT FROM AUTHOR]

Solid rocket motor (SRM)‐type structures are popular due to their reliability, considering that service safety during transportation can be improved by applying advanced vibration control technologies. In this study, a negative‐stiffness‐enhanced isolation system (NSeIS) with appropriately designed linear and nonlinear properties was developed to vertically isolate SRMs subjected to transportation‐ and deployment‐induced vibrations. The NSeIS design, based on the combination of a negative‐stiffness device and vertical isolator, involved a clear mechanical model, physical realization, and mechanical properties. Parametric analyses were performed on a typical SRM controlled with a linear and nonlinear NSeIS and a conventional isolation system. Subsequently, a feasible parameter domain and design recommendations were deduced. Finally, design cases for the SRM for time‐domain verification were considered. The results revealed that the NSeIS offers a flexible and enhanced isolating effect through the parallel arrangement of the negative‐stiffness device and conventional isolators. For the motor‐type structure, NSeIS ensures marked enhancements in performance and multiple levels of mitigation effects. Thus, compared with a conventional isolator with the same damping, NSeIS achieves a more substantial negative‐stiffness effect for a large displacement response range owing to its nonlinear property. NSeIS can isolate more vibration‐induced energy, thereby suppressing the interface Mises stress, which is essential for SRM‐type structures. [ABSTRACT FROM AUTHOR]

The advancement of solid rocket technology has increased demand for high‐performing propellants, particularly in the aerospace industry. Solid propellants are an essential component of rocket engines, and their stable combustion and adjustable burning rate throughout a wide pressure range significantly affect the performance of motors. The burning rate modifiers include catalysts and inhibitors. Solid rocket motors for different purposes require propellants with different burning rates. Using propellants that burn at a slower rate is beneficial for the smooth release of propellant energy, reducing the loss of energy in the process of high burning rate release and improving the endurance time of missile engines. Therefore, it is necessary to decrease the burning rate of propellant. This article summarizes the development of burning rate inhibitors (BRIs) and analyzes the mechanisms and behaviors of different BRIs, including amide‐based compounds, metal salts, and cationic surfactants, that affect the combustion of propellants. [ABSTRACT FROM AUTHOR]

To achieve low-cost and on-demand launches of microsatellites, the authors have been researching and developing a micro hybrid rocket since 2014. In 2018, a ballistic launch experiment was performed using the developed hybrid rocket, where it reached an altitude of about 6.2 km. The rocket engine had a 3D-printed solid fuel grain made of acrylonitrile butadiene styrene (ABS) resin in combination with a nitrous oxide oxidizer. The fuel grain port had a star-fractal swirl geometry in order to increase the surface area of the port, to promote the laminar–turbulent transition by increasing the friction resistance, and to give a swirling velocity component to the oxidizer flow. This overcame the hybrid rocket's drawback of a low fuel regression rate; i.e., it achieved a higher fuel gas generation rate compared with a classical port geometry. In 2021, the hybrid rocket engine was scaled up, and its total impulse was increased to over 50 kNs; it reached an altitude of 15 km. In addition to the engine, other components were also improved, such as through the incorporation of lightweight structures, low-shock separation devices, a high-reliability telemetry device, and a data logger, while keeping costs low. The rocket was launched and reached an altitude of about 10.1 km, which broke the previous Japanese altitude record of 8.3 km for hybrid rockets. This presentation will report on the developed components from the viewpoint of system design and the results of the ballistic launch experiments. [ABSTRACT FROM AUTHOR]

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During the ignition process of a solid rocket motor, the pressure changes dramatically and the ignition process is very complex as it includes multiple reactions. Successful completion of the ignition process is essential for the proper operation of solid rocket motors. However, the measurement of pressure becomes extremely challenging due to several issues such as the enormity and high cost of conducting tests on solid rocket motors. Therefore, it needs to be investigated using numerical calculations and other methods. Currently, the fundamental theories concerning the ignition process have not been fully developed. In addition, numerical simulations require significant simplifications. To address these issues, this study proposes a solid rocket motor pressure prediction method based on bidirectional long short-term memory (CBiLSTM) combined with adaptive Gaussian noise (AGN). The method utilizes experimental pressure data and simulated pressure data as inputs for co-training to predict pressure data under new operating conditions. By comparison, the AGN-CBiLSTM method has a higher prediction accuracy with a percentage error of 3.27% between the predicted and actual data. This method provides an effective way to evaluate the performance of solid rocket motors and has a wide range of applications in the aerospace field. [ABSTRACT FROM AUTHOR]

Presently, the use of lean combustion has been growing in industrial furnaces, aero-engine, and rocket engine applications due to their low $$N{O_X}$$ N O X emission and minimum maintenance requirements. On the other hand, due to the excessively low combustion reaction rate, the operation at a low fuel–air ratio leads to the sudden shutdown of the engines, which is termed lean blowout (LBO). Therefore, to operate land-based and aircraft engines at an extremely lean fuel–air mixture, control of LBO has become indispensable. In the present work, we develop several feedback in-loop control strategies using a pixel-averaged intensity ratio between the red and blue components of the flame emission. During a lean operation, based on the quantification of the proximity of the combustion dynamics to the LBO limits using this color ratio, the control strategies make the solenoid valves active to inject a secondary pilot fuel and energize the flame base. We notice that a proper choice of the color ratio as a threshold enhances the stability of both premixed and partially premixed combustion using a low pilot fuel. Besides, we test $$N{O_X}$$ N O X emission for unpiloted and piloted operations and find that the emission is only marginally affected during a pilot injection. [ABSTRACT FROM AUTHOR]

To elucidate complex flow phenomena in the solid rocket motor (SRM), it is beneficial to reduce development costs by employing computational fluid dynamics simulation, as conducting full-scale model experiments for SRM is expensive. This study focuses on numerically simulating the internal flow in the cylinder simulating the SRM geometry using the Building-Cube Method (BCM). First, the objective of this study is to develop a radial injection flow model (RIF model) for the BCM solver that can be utilized in the numerical simulation of internal flow in the cylinder simulating the SRM geometry. The numerical model geometry is based on the Ariane 5 SRM, which is a 1/30th scale axisymmetric model. The calculation results are validated through comparisons with experimental measurements, demonstrating the effectiveness of the RIF model using the immersed boundary method for accurate internal flow simulations of cylinder simulating the SRM geometry. Next, the effect of differences in wall injection conditions on the internal flow field of the cylinder simulating a solid rocket motor geometry was investigated using the BCM solver. The results indicate that the internal flow field and the vortex structure inside the cylinder simulating a SRM geometry change due to the effect of the injection velocity. [ABSTRACT FROM AUTHOR]

Nanosatellites of CubeSat type due to, i.a., minimized costs of space missions, as well as the potential large application area, have become a significant part of the space economy sector recently. The opportunity to apply miniaturized microsystem (MEMS) tools in satellite space missions further accelerates both the space and the MEMS markets, which in the coming years are considered to become inseparable. As a response to the aforementioned perspectives, this paper presents a microfluidic mixer system for biological research to be conducted onboard CubeSat nanosatellites. As a high complexity of the space systems is not desired due to the need for failure-free and remotely controlled operation, the principal concept of the work was to design an entirely passive micromixer, based on lab-on-chip technologies. For the first time, the microfluidic mixer that uses inertial force generated by rocket engines during launch to the orbit is proposed to provide an appropriate mixing of liquid samples. Such a solution not only saves the space occupied by standard pumping systems, but also reduces the energy requirements, ultimately minimizing the number of battery modules and the whole CubeSat size. The structures of the microfluidic mixers were fabricated entirely out of biocompatible resins using MultiJet 3D printing technology. To verify the functionality of the passive mixing system, optical detection consisting of the array of blue LEDs and phototransistors was applied successfully. The performance of the device was tested utilizing an experimental rocket, as a part of the Spaceport America Cup 2023 competition. [ABSTRACT FROM AUTHOR]

The potential for throttle control of hybrid rocket systems has long been known as a potential advantage for a variety of applications. Because only a single flow path is controlled, theoretically, hybrids should be significantly easier to throttle than bipropellant systems. Unfortunately, the slow response times and nonlinearity of traditional position-control valves have limited practical applications of hybrid throttling. This paper presents an alternative throttling system where the oxidizer flow path is broken into multiple streams, with each flow path controlled by a solenoid-operated on/off valve. The parallel paths allow significantly faster and more precise control than can be achieved using a single position-control valve. The achievable thrust levels are limited only by the size and number of components in the valve cascade. The 8-bit digital throttle system, developed by Utah State University's Propulsion Research Lab, uses commercial, off-the-shelf components. The throttle system was tested using a 200-N hybrid rocket motor, burning gaseous oxygen, and ABS plastic as propellants. The testing campaign of more than 50 hot fires has demonstrated multiple profiles, including deep throttle ramps, multistep boxcars, and sine waves at frequencies varying from 0.25 to 1-Hz. Comparisons to analytical models are also presented, showing good agreement. Fourier-transform spectra demonstrating the total-system, frequency response are also presented. [ABSTRACT FROM AUTHOR]

The pre-cooler is a key component of the pre-cooled turbine combined cycle engine, and its performance significantly impacts the overall engine performance. To clarify the flow and heat transfer characteristics of the pre-cooler and the effects of its key structural parameters on engine performance, the pre-cooler of the SABRE engine (Synergetic Air-Breathing Rocket Engine) was analyzed using numerical simulation methods to investigate the influences of air crossflow tube bundles and tube spacing on pre-cooler performance. The results indicate that increasing the number of air crossflow tubes significantly enhances heat transfer capacity; however, it also leads to an increase in the total pressure drop. Specifically, as the number of air crossflow tubes increases from 24 to 48, the overall heat transfer capacity improves by 42.1%, while the total pressure loss coefficient nearly doubles. Additionally, increasing tube spacing reduces the overall pressure drop, but this comes at the cost of decreasing heat transfer capacity and structural compactness. When the total pressure loss coefficient was reduced by approximately 29.8%, the overall heat transfer capacity decreased by 4.9%. Notably, the impact of tube spacing on flow resistance is greater than its effect on heat transfer, suggesting that the total pressure loss can be minimized by optimizing tube spacing. Therefore, both performance and structural integrity must be considered in pre-cooler design. Finally, selecting appropriate structural parameters based on operating conditions is essential to optimize heat transfer efficiency and overall design quality. [ABSTRACT FROM AUTHOR]

Studies on electric turbopumps for liquid rocket engines have recently increased. However, it has not been used in large-scale engines for heavy launch vehicles because the mass of power electronic devices, such as batteries and motors, has become too large. This paper proposes partial electric rocket engine cycles, in which all propellants are pressurized by conventional gas-turbine-driven turbopumps; however, an additional electric pump is used only to increase the pressure of the turbine inlet flow. It was found that the turbine flow rate was reduced and the total specific impulse increased using a reasonable sized electric turbopump for engines based on the expander bleed cycle. As a result, the launch capacity of a launch vehicle increases, even when accepting the extra mass of the electric turbopump system. The design feasibility of the electric turbopump was discussed, and it was concluded that such an electric pump system is feasible for upper-stage rocket engines using the current level of power electronics devices. [ABSTRACT FROM AUTHOR]

In this paper, the mechanism of rotating detonation is analytically discussed using a two-dimensional sheet model. Two ratios are employed in this discussion: the ratio of the sonic point width to the detonation front width, and the ratio of the effective mixture injection area to the injection area. In the rotating detonation, the unconfined boundary can increase the width at the sonic point and decrease the detonation wave speed. Although the high detonation pressure hinders mixture injection, effectively preventing some injectors from functioning, the high pressure acting on the injection end wall produces thrust. Mass, momentum, energy, and angular momentum conservations are used to determine these ratios. The calculated results are in reasonable agreement with past experimental and numerical findings. The present model succeeds to clarify the features, parameter relationships, and overall mechanism of the rotating detonation analytically and to specify flow field of the rotating detonation under given boundary conditions, e.g., the velocity deficit and the effective injection area ratio. The specific impulse of a rotating detonation engine was lower than that of an ordinary rocket engine due to the lower combustion gas pressure when the combustion gas expanded to 1 atm. The thrust coefficient and the specific impulse of an air-breathing rotating detonation engine were shown to be lower than those of a ramjet engine, respectively, primarily because of the smaller airflow rate into the engine. [ABSTRACT FROM AUTHOR]

The rapid development of space launch vehicle has put forward increasingly strict requirements for engine startup characteristics and thrust regulation capabilities. The flow regulator is a critical component in the operation of liquid rocket engines, particularly due to its role in precise thrust control and self-adaptation to perturbations. This study focuses on the complex fluid–structure interaction phenomena present in liquid rocket engine flow regulators. A three-dimensional (3D) transient numerical model was constructed using computational fluid dynamics and dynamic grid methods to investigate the response mechanism of the regulator under various disturbances. After experimental verification, the model error is less than 3%. The results demonstrate that under sinusoidal pressure perturbations (10–50 Hz), the flow oscillation frequency and amplitude are directly proportional to the perturbation frequency. The amplitude of step pressure perturbations is approximately linearly related to the amount of mass flow rate overshoot. There existed an inverse relationship between vortex size and intensity and the magnitude of inlet perturbations. The characteristics of the flow field are closely coupled with the mass flow rate, and post-perturbation recovery depends on the restoration of the flow field. Furthermore, the existence of friction will cause the sliding sleeve to experience a stagnation phase, and the greater the friction, the longer the stagnation phase. These findings provide valuable insight into the intricate dynamics of liquid rocket engine flow regulators and contribute to the design and optimization of future liquid rocket engines. [ABSTRACT FROM AUTHOR]

The rotating detonation rocket engine (RDRE) fueled by hydrogen/oxygen propellant represents a promising propulsion technology due to its high thermodynamic efficiency and propellant superior specific impulse. The rotating detonation wave (RDW) must propagate in a specific propagation mode while maintaining the self-sustaining state to ensure stable operation. An experimental system of hydrogen/oxygen fueled RDRE was developed in the present study. The operation of RDRE and propagation mode of RDW were investigated under atmospheric pressure conditions, and both hollow and annular combustors were tested. The high-frequency pressure fluctuations in the RDRE were measured by the dynamic pressure transducer, while a high-speed camera was used to capture images of flame luminescence at the rear end of the RDRE. The experimental results showed that the RDW could be initiated and reached a self-sustaining propagation state with hydrogen/oxygen propellant in the hollow and annular RDRE. A single-wave mode, a two-wave co-rotating mode, and a three-wave co-rotating mode were visualized under different conditions. With the increase in the equivalence ratio, the number of rotating detonation fronts decreased, and the variations in the RDW propagation modes were consistent in the hollow and annular RDRE. However, when the equivalence ratio exceeds 1.2, the propagation velocity decreases sharply in the annular combustor, while in the hollow combustor the RDW propagates stably, revealing a higher upper limit for the equivalence ratio. Also, the dominant frequency distribution was more concentrated in the hollow combustor. The findings provide valuable insight into the variations in detonation modes related to the equivalence ratio and combustor configuration. [ABSTRACT FROM AUTHOR]

During the operation of solid rocket motors, the behavior of condensed particles impacting the wall will have a remarkable influence on the structure and performance of the engine. Especially when the aircraft is under overload flight conditions, the condensed particles will form a local high concentration particle flow under the action of inertia force, continuously scouring the surface of the insulation layer, seriously affecting the thermal protection structure and the work safety of the engine. Therefore, it is an essential issue to master the behavior mode of the condensate particle impinging the wall and clarify its dynamic characteristics and evolutionary mechanism. In this paper, the dynamic behavior of aluminum droplets impacting on the porous surface is experimentally investigated by preparing the porous wall, the influence mechanism of the porous structure on the spreading process of aluminum droplets is clarified, and the effects of the droplet's initial parameters as well as surface environment are analyzed. Combined with the fluid of volume method, the flow process of droplets on the porous surface is simulated. With the variation of the dimensionless parameters M and N, the main behavior patterns of the droplets obtained so far are rebound, adhesion, partial rebound, partial adhesion, and porous seepage. The presence of pore structure enhances the hydrophobicity of the wall and makes the droplets more easily broken during spreading. When the droplet initial energy is certain and the wall structure conditions change, there is a strong competitive relationship between its spreading and penetration. When the droplet initial energy is increased, its spreading and penetration strengths are significantly increased. The research results can provide a reference for the erosion process of condensed-phase droplets impinging on the char layer and provide theoretical basis and data support for the design and optimization of SRMs. [ABSTRACT FROM AUTHOR]

Turbine blades recovered from the Apollo Saturn V rocket F-1 engines were examined to determine an appropriate conservation protocol. Significant corrosion damage was observed in the turbine blades which appear to be made of a nickel based γ–γ′ superalloy. Pitting corrosion appears to have breached the surface of the turbine blades, and subsequently a form of dealloying corrosion preferentially attacked the γ′ phase. This corrosion left behind a thin network of interconnected γ phase, causing a severe loss of density of the blades and fragility of the blades. The particular alloy used for these turbine blades does not appear to be a known production alloy and may have been developed specifically for use in the F-1 rocket engines, with an increased concentration of refractory (Mo, Nb) elemental additions. The analytical results helped conservators determine a suitable treatment protocol for more than 400 blades and 100 fragments from four recovered turbines. [ABSTRACT FROM AUTHOR]

The flexible control in burning rate and stable combustion of solid propellants over a broad range of pressures essentially determine the performance of solid rocket motors. The nanoscale metals and their oxides are usually utilized as the traditional burn rate modifiers in adjusting the combustion behaviors of solid propellants, but the energy characteristics of propellants are greatly compromised owing to the inert nature of additives used. In this paper, a series of metal complexes of triaminoguanidine-glyoxal polymer (TAGP-Ms) with high nitrogen content were synthesized and employed as energetic burn rate inhibitors in tunning the combustion properties of composite modified double-based (CMDB) propellants. The influences of prepared TAGP-Ms on the thermal decomposition and combustion of CMDB propellants were comprehensively investigated based on thermal analysis technique and combustion characterization methods. It was found that the peak temperatures of the second exotherm displayed for the propellants containing TAGP-Ms inhibitors were significantly increased by more than 30°C (except for TAGP-Cu), indicating the TAGP-Ms is capable of suppressing the thermal decomposition of CMDB propellants. More importantly, the heat releases of CMDB propellants were substantially enhanced by 16%~34% determined by DSC analysis. In comparison to the blank reference, the energetic TAGP-Ms inhibitors could promote the heat of combustion for the involved CMDB propellants by 5%~9%. The burning rates of CMDB propellants were notably reduced over a wide range of pressures by adding TAGP-Ms, whereas the calculated pressure exponents at pressures ranging from 10 to ~ 22 MPa appear to be relatively high due to the accelerated HMX decomposition. The measured combustion wave temperatures suggest that the CMDB propellants with TAGP-Ms undergo a three-stage combustion process with the highest temperatures in the range of 2160 ~ 2230°C. [ABSTRACT FROM AUTHOR]

The development and application of numerical simulation to the study of gas-dynamic processes occurring in solid-propellant rocket motors (SPRMs) is discussed. A characteristic feature of internal flows in the SPRM channels and nozzles is the presence of the condensed phase of nonspherical particles. Mathematical problems in this area feature the simultaneous occurrence of processes on many time and spatial scales, which describe the formation of agglomerate particles, their combustion, and transport in a flow of combustion products in internal channels and nozzles. A multilevel multiscale technique that combines models describing the state of the system at the micro-, meso-, and macroscales is the approach used to solve these problems. An overview of models varying in complexity and level of detail is given. The construction of multiscale models is considered in relation to the simulation of two-phase flows with metal-oxide agglomerates formed in the propellant channel and representing drops of molten metal with oxide particles attached to their surface. The options of the developed approach are demonstrated by the calculations of flows of combustion products containing agglomerate particles in the channels and nozzles of propulsion systems. [ABSTRACT FROM AUTHOR]

A parametric study in terms of a global model has been performed on the kinetic processes in iodine plasma generated by electrodeless inductively coupled radiofrequency and microwave discharges. The formation dynamics of the plasma component composition has been obtained for various regimes. It has been shown that ion–ion plasma is formed at short times and a transition from ion–ion to electron–ion plasma occurs at times from several fractions to several milliseconds. The model made it possible to determine the most optimal regimes of iodine plasma generation in modern electric rocket engines. [ABSTRACT FROM AUTHOR]

A new system design tool for the simulation of a full hybrid rocket engine is developed in this article. This tool enables to simulate the behaviour of the engine at different conditions/configurations in several minutes from a desktop computer by keeping a balance between accuracy and computation time. This makes its use especially attractive for the pre-design phases of the engine. The main components of the hybrid rocket engine from our laboratory-characterized by a catalytic injection of the oxidizer-, are modelled and implemented in the tool: the feed/injection sub-system through 0-D models of a mass flow rate regulator and a catalyst; the combustion chamber, with a 1.5-D model; and the nozzle through a 1-D model. An iterative method is employed to reach the convergence of pressure in the combustion chamber between these sub-systems. The corresponding set of equations is solved by a Newton–Raphson technique. Seven experiments performed on our lab-scale engine were used to validate the system design tool. In five of the simulations, the relative differences of the main physical quantities with the experiment were below 30%, being the largest errors found in the fuel regression rate and mixture ratio. The best agreements with the experiments were retrieved for the cases with the largest oxidizer mass fluxes (above 230 kg/ m 2 /s) and mixture ratios closest to stoichiometry, defining thus, the range of applicability of the system design tool. The results presented in this article were issued of a Ph.D. thesis. • A 1.5-D combustion chamber model of a hybrid rocket engine was developed and validated. • 0-D/1-D mass flow regulator, catalyst and nozzle models were developed and validated. • A system design tool of a hybrid rocket engine implementing these models was developed. • The tool is adapted to pre-design phases, with a few minutes of computation time. • Errors between 8%–40% in fuel regression rate and below 10% in thrust and chamber pressure. [ABSTRACT FROM AUTHOR]

Production of rocket propellants from lunar resources has the potential to significantly reduce the cost of space exploration. Recent research efforts in this area were focused on the extraction of water from icy regolith for conversion into hydrogen and oxygen, a highly efficient rocket bipropellant. However, water is available only in polar regions of the Moon, and its extraction is a challenge. The present paper aims to assess the feasibility of using propellant components that can be obtained from lunar regolith, specifically oxygen, metal alloys, and sulfur. Thermodynamic performance characteristics of rocket engines using these components were calculated over wide ranges of oxidizer-to-fuel mass ratios. It has been shown that the fuel obtained by extraction of oxygen from regolith, i.e., primarily a mixture of metal alloys, exhibits a relatively high specific impulse of up to 250 s. The use of fuel-lean propellants significantly decreases the temperatures, which facilitates cooling and potentially reduces the deposition of condensed products in the engine; at the same time, the expected decrease in the specific impulse is less pronounced. The use of sulfur in rocket engines is less promising from a thermodynamic point of view, but it enables engine designs without a need for feeding metal alloy powders. Among different designs of sulfur-based engines, a hybrid rocket (fuel: metal alloys mixed with sulfur, oxidizer: liquid oxygen) appears to be the most promising. • Rocket propellants can be directly sourced from lunar regolith. • Such regolith-derived fuels can produce a peak specific impulse of up to 250 s. • Sulfur can be used as either an oxidizer or additive to the regolith- derived fuel. • Different configurations for rocket engines using these fuels are presented. [ABSTRACT FROM AUTHOR]

A large variety of metal fuels is usually added to the modern solid rocket propellants to improve the propellant energy and motor specific impulse and to suppress high frequency unstable combustion. Among them, aluminum is the most common additive. The combustion process of aluminum can significantly affect the combustion characteristics of a solid rocket motor. In this paper, the SPH (Smoothed Particle Hydrodynamics) method is used to simulate the combustion process of aluminum particles. First, the SPH discrete equations with an evaporation combustion model are derived. On this basis, the evaporation combustion process of aluminum particles is numerically simulated. The results show that when aluminum particles are heated and evaporated in a static flow field, an alumina shell will be formed on the surface, and further thermal expansion will cause the alumina shell to break. The molten aluminum will spray out, and an aluminum cap will be formed on the surface. The "microburst process" will be similar to the gel droplet. In the convective environment, the flame structure of aluminum particles will be obviously peach shaped, which will wrap aluminum particles at the bottom. The simulation results are consistent with the experimental results. [ABSTRACT FROM AUTHOR]

Supercritical fluids exist widely in nature and have enduringly attracted scientific and industrial interest. In power systems like liquid rocket engines, fluids undergo the trans-critical process transferred from the subcritical state to the supercritical state, and the phase change process exhibits different features distinguished from subcritical evaporation. In this work, we conducted a series of molecular dynamics studies on the behavior of methanol (MeOH), n-heptane (C7), and binary C7/MeOH droplets under supercritical nitrogen environments. The emphasis is on clarifying the transient characteristics and physical origins of the trans-critical evolution of droplets. During the trans-critical process, droplets are found to experience an unstable period without a spherical shape, where the droplet diameter no longer decreases, violating the traditional d2-law rule. The occurrence of nonspherical droplets is related to the microstructural behavior of trans-critical droplets. Two types of microscopic structures within the droplet are identified: large-scale thermally induced clusters for long-chain C7 and hydrogen-bond connected network-like structures for MeOH, which contains hydroxyl (–OH) groups. Based on these findings, the mechanism behind the evolution of trans-critical droplets is illustrated. Finally, we determine the boundary of ambient conditions in the form of dimensionless expressions T r − 1 = a (p r − 1) − b , which dictate whether droplets can maintain a spherical shape during the trans-critical process. [ABSTRACT FROM AUTHOR]

Swirling steam jet condensation holds significant applications in industrial processes such as nuclear safety and gas–liquid mixing in the oxygen transmission pipeline of the liquid rocket engine. However, due to its involvement with complex flow and phase-change heat transfer, the application and optimization of related condensation technologies still face challenges. Therefore, this paper aims to investigate the condensation characteristics of the swirling steam jet by numerous experiments. The steam mass flux is 15–45 kg/(m2·s), and the water temperature ranges from 40 to 85 °C. A novel X-type swirl pressure nozzle is selected to achieve the swirling flow of the steam jet. A comparative analysis is conducted on the interface behavior and evolution of condensation parameters of the non-swirling and swirling steam jets during condensation processes. Results show that the swirling jet condensation includes three flow patterns, namely, chugging regime, smooth grown bubble regime, and rough grown bubble regime. Compared with the non-swirling steam jet condensation, swirling steam jets exhibit a 10.36% increase in the smooth grown bubble regime region and a 14.63% decrease in the rough grown bubble regime. Swirling bubble morphology evolves steadily, and the surface is smoother and more rounded. Simultaneously, irregular deformation behaviors can also occur in the swirling bubble condensation process, such as spiral growth of jet bulge, neck torsion, and the corolla pattern. This deformation helps to increase the contact area and prolongs the bubble lifetime, allowing for more adequate heat transfer at the steam–water interface. The swirling motion of the steam jet will reduce the bubble collapse frequency. As the water temperature rises from 60 to 80 °C, the bubble condensation rate and collapse frequency decrease. The bubble radius increases and the condensation time is extended. With the increasing steam mass flux, the collapse frequency gradually increases. The condensation rate and the bubble radius vary nonlinearly. At the higher steam mass flux, the swirling motion can effectively release the heat that accumulates inside the bubble after reaching the condensation equilibrium state. [ABSTRACT FROM AUTHOR]

Titanium alloy, Ti–6Al–2.2Mo–1.4Cr–0.4Fe–0.3Si (BT3-1), is a two phase α + β alloy developed for applications in rocket engines, gas turbine engines, and aircraft frames for service up to a temperature of 450 °C. The hot workability of this alloy has been studied through isothermal hot compression testing in the temperature and strain rate (ε ˙) range of 800 °C to 1000 °C and 10−3 to 10 s−1, respectively, in a thermomechanical simulator. Processing maps using dynamic material model has been generated and different regions of the map were correlated with microstructural observations. The flow stress data were fitted in Arrhenius strain-compensated model and constitutive equations were developed. Optical microstructures revealed elongated grains, kinking of α phase, flow localisation, and adiabatic shear bands at lower temperatures. Super-plasticity was found to be operative at low temperature of 850 °C and ε ˙ 10−3 s−1, whereas dynamic recrystallization (DRX) was dominating at high temperatures of 950 °C to 1000 °C and ε ˙ of 10−3 s−1. Finite element analysis showed the flow localization in the unstable regions of processing map. Enhanced hot workability was achieved above 950°C in the ε ˙ of 10−2−10−3 s−1 due to initiation of DRX in view of an increase in the β phase fraction. [ABSTRACT FROM AUTHOR]