Search Results (186)

Hybrid rocket engines (HREs) are increasingly considered for space transportation due to their inherent safety, cost-effectiveness, and operational flexibility. However, their diffusion flame structure promotes soot formation, leading to potentially significant emissions of black carbon (BC) directly into the stratosphere. This review provides a comprehensive analysis of soot formation and evolution in hybrid rocket combustion, covering fuel pyrolysis, polycyclic aromatic hydrocarbon (PAH) growth, particle nucleation, surface growth, and oxidation processes. The physico-chemical properties and nanostructural evolution of soot are discussed in relation to combustion conditions typical of HREs. Particular emphasis is placed on the environmental and climatic implications of BC emissions, including radiative forcing, atmospheric lifetime, and heterogeneous chemical interactions. Existing diagnostic techniques for soot measurement are critically assessed, highlighting their limitations under the extreme conditions of rocket exhaust plumes. A key outcome of this review is the identification of a major knowledge gap: although preliminary experimental emission-index measurements for HRE soot have recently become available, the database remains extremely limited, fuel- and configuration-specific, and not yet supported by standardised diagnostic protocols. Addressing this gap is essential for accurate environmental impact assessments and for the development of sustainable propulsion technologies. Full article

Hybrid rocket engines (HREs), which pair a fluid oxidizer with a solid fuel, offer safety and handling advantages, can reduce environmental impact relative to selected legacy systems, and are capable of deep throttling and restart, making them strong candidates for “green” in-space propulsion applications. However, until very recently, there has not been any flight heritage of an HRE used in a spaceflight mission. Over the past decade, the Propulsion Research Laboratory at Utah State University (PRL-USU) has matured a portfolio of HRE technologies—low-energy arc ignition, digital throttling, additively manufactured sustainable fuels, the Nytrox green oxidizer, and electroplated thruster assemblies—that together address the historical barriers to spaceflight adoption. This paper summarizes that progress, describes two flight systems built upon it, and presents a roadmap for future applications. Full article

Hybrid rocket motors are an attractive option for the upper-stages of low-cost small launchers, but are susceptible to variability in performance both in time and between firings. Moreover, key contributors to hybrid motors’ performance such as oxidizer-to-fuel ratio ($O/F$) are difficult to estimate, and by extension, to control. Four approaches were evaluated for the estimation and control of $O/F$ under system uncertainty, including through on-line estimation by an Unscented Kalman Filter (UKF). A Monte Carlo analysis was conducted of a simulated hybrid kick motor, where key sources of system uncertainty such as the characteristic velocity efficiency (${\eta }_{c^{*}}$), fuel regression coefficients, and oxidizer flow characteristics were allowed to be variable. Feedback control of $O/F$ informed by the UKF obtained 6.8% smaller control error than the best alternative approach. Yet the Monte Carlo analysis showed that among uncertainty sources considered, ${\eta }_{c^{*}}$ was the primary driver of performance variability, while $O/F$ regulation had a small influence. This was because the total and specific impulses were relatively insensitive to $O/F$ for the considered motor configuration and ranges of $O/F$ observed during the simulated burns—highlighting the importance of system uncertainty quantification when formulating performance-regulating interventions. Further, the proposed UKF observer provided data-informed estimates of combustion efficiency and propellant residuals in time, which are valuable for the planning and execution of accurate orbital insertions in a kick motor susceptible to performance uncertainty. The developed uncertainty quantification and control modeling framework can be used also during the design and assessment of other control interventions under system uncertainty. Full article

The use of liquid hydrogen (LH₂) in rocket propulsion systems is among the most critical technologies enabling highly efficient space exploration. Fuel-oxidizer combinations directly impact mission performance, safety, and sustainability. In the literature, determining which oxidizer to use to obtain an efficient combination in experimental studies is a valuable area of research. However, evaluating different alternatives according to various criteria in experimental studies is expensive, time-consuming, and quite dangerous. This study aims to provide decision-makers with analytical support through a novel multi-criteria decision-making methodology. In this context, the simple weight calculation (SIWEC) method is integrated to determine the weights of the criteria, and the mulTi-noRmalization mUlti-distance aSsessmenT (TRUST) method is integrated to evaluate the oxidizers. The results show that the most important criterion is combustion, followed by environmental sustainability, applicability, cost, and operational safety. The foremost oxidizer, according to the analyses, parametric sensitivity analysis scenarios, and comparative analyses, is overwhelmingly LH₂-liquid oxygen (LOX). In all analyses, fluorine (F₂) ranked second, followed by the FLOX mixture in third place, and ozone (O₃) in last place. In addition, sensitivity analyses based on α and β parameter variations and comparative analyses were conducted to evaluate the robustness and stability of the proposed decision-making framework. Full article

This study presents a novel one-dimensional solution method to demonstrate the effects of fuel composition and channel roughness on phase-change heat transfer in spiral regenerative cooling systems. The calculated models are grounded in an experimental correlation of liquefied natural gas (LNG) flow boiling, and their accuracy is validated through ignition experiments conducted on a 1 kg/s-class thrust chamber. The experimental data shows that the physical characteristics of LNG contribute to an extended reach within the two-phase region, resulting in a calculated pressure drop that exceeds that of pure liquid methane. Variations in surface roughness influence the pressure drop by altering the frictional coefficient. Specifically, an increase in surface roughness from 2 µm to 8 µm results in a 47.8% rise in pressure drop. The proposed model demonstrates high accuracy, with deviations in the coolant temperature rise and the pressure drop being less than 9.0% and 7.6%, respectively, when compared to experimental data. The findings serve as an engineering guide for designing and optimizing heat transfer in LOX/LNG rocket engine cooling systems. Full article

This study experimentally investigated the effect of characteristic chamber length, $L^{\ast }$, on combustion efficiency and stability in CAMUI-type hybrid rockets using 70 wt% and 80 wt% hydrogen peroxide under non-catalytic spray-injection conditions. Combustion tests were conducted by systematically varying $L^{\ast }$ through changes in the nozzle throat diameter while maintaining the combustor volume constant. For both oxidizer concentrations, the characteristic exhaust velocity efficiency, ${\eta }_{c^{\ast }}$, increased with increasing $L^{\ast }$. The 70 wt% cases required a larger $L^{\ast }$ than the 80 wt% cases to achieve comparable efficiency, and flame blowoff occurred in the low-$L^{\ast }$ region. The normalized RMS pressure fluctuation was also larger for the 70 wt% cases, particularly in the low-$L^{\ast }$ region, indicating lower combustion stability. These results indicate that reducing the hydrogen peroxide concentration increases the $L^{\ast }$ required to maintain stable and efficient combustion. As a key outcome of this study, stable and efficient combustion of 70 wt% hydrogen peroxide was demonstrated without catalytic assistance when a sufficiently large $L^{\ast }$ was provided. These results demonstrate the capability of the CAMUI-type combustor to extend stable operation toward lower oxidizer concentrations and experimentally clarify the concentration-dependent $L^{\ast }$ requirement as a practical design guideline for catalyst-free hydrogen peroxide hybrid rockets. Full article

In 2024, eight Soyuz launch vehicles launched from the Baikonur Cosmodrome created 40 falling sites in Central and Northern Kazakhstan. During field work at all falling sites, the area of disturbances was assessed. The average area of zones containing large fragments was 63 m², and the average area of scatter fields of small fragments was 11,605 m². The maximum area of jet fuel spills reached 125 m², and that of hydrogen peroxide spills reached 196 m². In 2024, due to falling debris, total mechanical disturbances covered 6912 m², and total pyrogenic impact effected 1,211,453 m². At the first stage falling sites during the plant growing season, the area affected by fires and H₂O₂ spills was significantly larger, while the area of very weak mechanical disturbance was smaller (p < 0.05). During the mud season, the area with very strong mechanical impact was greater compared to winter and summer (p < 0.002). Compared to the first stage falling sites, the Soyuz abort mode falling sites were characterized by a larger area of very strong mechanical impact (p < 0.02), evidenced by deep craters, and a smaller area covered by large fragments. The data obtained can be used when planning launch dates from the Baikonur Cosmodrome to minimize negative environmental impact on the fragile and slow-recovering arid ecosystems of Central and Northern Kazakhstan. Information on disturbance areas can be used to monitor the natural recovery of affected ecosystems and to calculate the environmental damage caused by rocket and space activities in 2024. Full article

Solid rocket propellants based on hydroxyl-terminated polybutadiene (HTPB), in which HTPB acts as the polymeric binder and fuel matrix, are widely used in aerospace propulsion. During storage, transport, and service, these composite energetic materials are exposed to sustained mechanical loads as well as environmental variations, which may induce time-dependent inelastic deformation. Such creep deformation can alter the grain geometry, affect combustion stability, and reduce the structural reliability of rocket motors. In this work, room-temperature tensile creep tests were conducted on an HTPB-based solid propellant under different stress levels. Several viscoelastic and power-law constitutive models were compared, and a composite time-hardening creep model was established to describe the experimental strain–time response. The model was further implemented in Abaqus through a Fortran user subroutine for finite element simulation. The results provide a useful basis for creep deformation assessment, formulation optimization, and structural reliability analysis of HTPB-based propellants. Full article

Liquid rocket engine injectors play a fundamental role in determining combustion efficiency, stability, and overall propulsion performance. This review paper provides a comprehensive analysis of liquid-injector technologies used in space propulsion systems, with emphasis on their historical evolution, atomization mechanisms, and cross-domain insights from aviation fuel injection systems. The study begins by examining the fundamental processes governing liquid jet breakup, including primary and secondary atomization, ligament formation, and droplet generation, together with the non-dimensional parameters that control these phenomena. The historical development of injector architectures -from early orifice-based and impinging designs to modern coaxial and pintle configurations—is then discussed in relation to increasing performance requirements and combustion stability challenges. A comparative perspective with aviation gas turbine injectors is introduced to highlight similarities in atomization physics and differences in operating conditions and design constraints. The paper also reviews experimental and numerical approaches used to characterize spray formation and injector performance. The results indicate that injector geometry and flow conditions strongly influence mixing efficiency, droplet size distribution, and combustion–acoustic coupling mechanisms. The study concludes that integrating cross-domain knowledge and modern design techniques is essential for advancing injector performance in next-generation propulsion systems. Full article

Hybrid rocket engines offer a compromise between safety, controllability, and performance, making them attractive for small-scale propulsion systems. However, oxidizer selection remains a critical early-stage design decision that cannot be determined solely from ideal thermodynamic metrics. This study presents a comparative analysis of three oxidizers—nitrous oxide (N₂O), gaseous oxygen (GOX), and liquid oxygen (LOX)—for a 1 kN-class hybrid rocket engine using HDPE fuel under identical operating conditions. Equilibrium combustion performance was first evaluated using NASA Chemical Equilibrium with Applications (CEA) to determine optimal oxidizer-to-fuel ratios and theoretical specific impulse. These results were subsequently refined using Rocket Propulsion Analysis (RPA) to incorporate finite combustion chamber geometry and non-ideal nozzle expansion effects. The equilibrium analysis predicts maximum specific impulses of approximately 260 s for N₂O/HDPE and nearly 300 s for oxygen-based systems. However, finite-geometry modelling indicates that practical performance is reduced by approximately 5–8%, yielding delivered specific impulses of about 275 s for GOX and 272 s for LOX. The results demonstrate that although oxygen (GOX and LOX) provides higher thermodynamic performance, the practical advantage of LOX over GOX becomes marginal at the kilonewton scale. Consequently, oxidizer selection for small hybrid engines should be treated as a system-level trade-off involving performance, infrastructure complexity, and operational safety. Full article

Hybrid rocket technologies are gaining recognition as eco-friendly alternatives to traditional propulsion systems. Utah State University’s Propulsion Research Laboratory has developed a High-Performance Green Hybrid Propulsion (HPGHP) technology, leveraging 3D-printed ABS fuel for reliable, low-energy ignition. Among tested materials, only ABS shows suitable electrical-breakdown properties for arc ignition. Unfortunately, due to the proprietary formulations in commercial ABS blends, and its limited use as a rocket-propellant, related composition and combustion data are limited. This study uses spectroscopic evaluation and bomb calorimetry to estimate material compositions, enthalpies of formation, and combustion energies for multiple commercially available 3-D print feed stock ABS types, finding minimal differences amongst the samples tested. Based on these test results, “representative” ABS properties including chemical formula, mean molecular weight, enthalpy of formation, and Higher Heating Value, is recommended. Follow-on tests with 5 alternative, commonly used, 3D-printable thermoplastic feed stocks demonstrate that ABS has significantly higher energy content. This result supports ABS’s advantages and utility as a conveniently fabricated hybrid rocket fuel. Full article

A rapid quasi-one-dimensional (quasi-1D) reacting-flow analysis code was developed for the preliminary assessment of rocket-based combined cycle engines over a broad flight envelope. The internal flow was modeled as steady and quasi-1D in a variable-area duct by solving the coupled conservation equations together with species transport, and finite-rate chemical kinetics were included to represent combustion-induced heat release and composition change. To incorporate configuration-dependent mixing effects that affect RBCC heat release evolution and thermal choking tendencies, a streamwise mixing efficiency distribution was extracted from non-reacting 3D CFD and prescribed as an input to the quasi-1D formulation to represent the progressive availability of reactable fuel along the flowpath. A mode-dependent solution strategy was established by separating the computation into scramjet mode and ramjet mode procedures with a switching criterion based on whether a sonic condition occurs within the combustor, allowing thermal choking and mode transition behavior to be addressed within a single framework. The numerical solver was implemented in Python 3.12.2 and integrated using a stiff ordinary differential equation (ODE) scheme to ensure robust convergence in the presence of reaction-induced stiffness. Verification against previously published hydrogen-fueled scramjet results reproduced the overall streamwise trends of key quantities including Mach number, pressure, temperature, and density. The developed code was then applied to an RBCC configuration under operating conditions representative of ERJ and ESJ regimes, and the quasi-1D predictions were compared with cross-section-averaged 3D RANS CFD results, showing consistent mode identification and comparable axial behavior at a level suitable for preliminary analysis with substantially reduced computational cost. Full article

The engine system requirements for different engine cycles significantly influence the design of the mixing head. A literature review of fuel-injection technology for hydrogen and methane is presented. The literature review aimed to answer proposed questions specific to the liquid rocket engine fuel injector design. The current review methodology accounts for the engine system effect. Thus, a comprehensive literature review of the working principles of startup-staged-combustion-cycle engines based on original patents is provided. At the end of the review, the research gaps and suggestions for further work are summarised. At high mass flow rate and injection pressure in the supercritical regime (>50 MPa), experience is limited to the staged-combustion cycle developed in Russia and the US. It is necessary to consider a fluid-dynamic heat transfer coupling study for the multi-injection element design in the supercritical state. Cryogenic spray atomisation experiments need to be designed with research significance in mind. It is still needed to study how the similarity of the spray flow field to the combustion performance affects a liquid rocket engine problem. Moreover, scaling stoichiometric mixing theory needs to be expanded to different injector types, such as tricoaxial and pintle injectors, to validate the correlation between the non-reactive mixing length and flame length. Full article

In the field of study of fluid–structure interaction (FSI), sloshing dynamics play a crucial role in various engineering applications, from aerospace to civil infrastructure. Finite Volume (FV)-based Computational Fluid Dynamics (CFD) methods for modeling free surface flows like sloshing are computationally expensive, particularly because high-resolution dynamic transient simulations are required. Moreover, FSI effects are usually considered by coupling different solvers for the fluid and the structural domain, respectively, thus adding to the computational burden due to the various steps of data transfer, interpolation, and mesh adaptation needed to obtain accurate results. On the other hand, reduced-order models of sloshing effects are usually obtained by tuning equivalent mechanical models, which often neglect more complex geometries and imperfections. To address this challenge, the use of acoustic finite elements, as an alternative approach for modeling free surface flows interacting with flexible structures, has been proposed previously. Such elements are defined with the sole dynamic pressure as the nodal degree of freedom; therefore, such methods can significantly accelerate simulations to predict sloshing-induced forces and pressure distribution, taking into account the actual geometry of the structure. Due to the reduced computational time, FSI analysis with acoustic elements can serve as a viable tool for control systems and design optimization. Potential applications of this approach include structural analysis of anti-sloshing devices in rocket propellant tanks, control systems for enhanced launch stability, and seismic safety assessment of liquid storage tanks, as well as slosh-induced wall load evaluation in the fuel and water reservoir, transportation, and energy systems. Validation of FSI effects is conducted against results from partitioned two-way coupled fluid–structural simulations. The simplified frequency-prediction model was reliable for practical flexibility ranges. Overall, this work deepens our understanding of how baffle characteristics influence slosh mitigation, offering valuable guidance for anti-sloshing device engineering. Full article

In this study, the feasibility of electrospraying as an alternative processing technique for the preparation of composite solid rocket propellants (SRPs) was investigated. The main objective was to improve microstructural homogeneity and interfacial contact between the oxidizer, energetic additive, and metallic fuel without altering the chemical composition of the formulation. Additionally, porous electrosprayed SRP formulations were prepared to examine the influence of controlled porosity on thermal decomposition behavior. The prepared materials were characterized using scanning electron microscopy combined with energy-dispersive X-ray spectroscopy (SEM/EDS) to assess microstructural features and component distribution. Thermal decomposition behavior and kinetic parameters were evaluated using simultaneous DSC/TG analysis conducted at multiple heating rates. Safety-related properties were assessed through friction sensitivity testing, while post-decomposition solid residues were analyzed using SEM/EDS and X-ray diffraction. The results show that electrospraying improves structural homogeneity, reduces solid residue formation after thermal decomposition, and decreases apparent activation energy, while maintaining unchanged friction sensitivity. These findings demonstrate the potential of electrospraying as a physical processing route for tailoring the microstructure and thermal behavior of composite solid rocket propellants. Full article