Search Results (135)

In the space sector, test benches are facilities used to perform ground testing of rocket engines with solid, liquid, or hybrid propellants. These tests typically occur during the stages of development, validation, qualification, certification, and acceptance of the rocket engine assembly. Due to the periodicity and complexity related to these stages, it may be necessary to carry out several test campaigns in order to obtain the expected results. This requires skill and knowledge from the technical team involved, since the occurrence of operational nonconformities in actuator components installed in this test bench constitutes a situation with sufficient potential to put human lives at risk and/or cause material losses. In this context, this work presents a proposal for a Graphical User Interface to be integrated into the architecture of a system that performs the operational self-testing of these actuators. The virtual resources established for the windows of the Graphical User Interface are expressive and are related to the procedures defined to perform the self-testing of actuator activation, steady-state operation, and shutdown. The validation of the functionality of this Graphical User Interface is obtained through tests carried out on a prototype that was developed considering the main blocks contained in the mentioned system architecture. The satisfactory results observed in these tests suggest that the Graphical User Interface is suitable for the purpose for which it is intended. 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

High-speed permanent magnet synchronous motors (PMSMs) used in electric pump-fed liquid rocket engines require stator shielding sleeves to prevent corrosive propellants from causing harm under cyclic pressure. However, metallic sleeves suffer significant losses due to eddy currents. Conversely, pure carbon fiber reinforced polymer (CFRP) sleeves have failed when exposed to 98% H₂O₂. Micro-CT analysis of a failed pump sleeve reveals a four-stage failure mechanism. Manufacturing defects caused matrix cracking, which propagated under pressure and thermal cycling. This progression resulted in the formation of through-thickness leakage paths, which ultimately triggered catalytic decomposition and explosion. To address these issues, an improved dual-layer sleeve is proposed, featuring a 2.5 mm PEEK 450G liner and a 2.0 mm T700S/epoxy CFRP overwrap. Finite Element Analysis (FEA) indicates peak von-Mises stresses of 86.25 MPa and 112.16 MPa, yielding Tsai–Wu safety factors of 2.9 and 1.7. Furthermore, various tests, including immersion, fatigue, burst, hydraulic, and thermal evaluations, demonstrate a burst margin of 2.37× at 7.12 MPa, with only 0.19% increase in mass. This design effectively eliminates leakage pathways while preserving zero eddy-current loss and ensuring a low weight. Full article

High-pressure liquid CO₂ jets possess the characteristics of low-temperature cooling and dry, residue-free impact, which makes this technology particularly suitable for removing hydroxyl-terminated polybutadiene (HTPB) propellant from decommissioned solid rocket motors. However, existing studies lack multi-objective optimization of impact efficiency and CO₂ consumption, which limits their engineering applications and further promotion. In this study, a high-accuracy quadratic Response Surface Methodology (RSM) relating process parameters to damaged volume was established using a Box–Behnken design (BBD) combined with three-dimensional topography scanning. A theoretical model for CO₂ consumption was developed based on the Homogeneous Equilibrium Model (HEM). On this basis, the Non-dominated Sorting Genetic Algorithm II (NSGA-II) was used to obtain the Pareto-optimal set for maximizing propellant damaged volume and minimizing CO₂ consumption. The results indicate that nozzle diameter has the most significant effect on damaged volume and exhibits a strong interaction with jet pressure. The knee-point solution gives a jet pressure of 15.35 MPa, a stand-off distance of 5 mm, and a nozzle diameter of 1.8 mm. Compared with the initial condition, this compromise condition increases the damaged volume by 72% while increasing CO₂ consumption by only 4.9%. Furthermore, the temperature in the impact zone was reduced to a minimum of −92.4 °C, with no thermal accumulation observed. These findings reveal the influence of liquid CO₂ jet process parameters on impact efficiency and CO₂ consumption, providing a theoretical basis and parameter references for its engineering application in the safe removal of propellants from decommissioned solid rocket motors. Full article

Impinging injectors are extensively utilized in liquid rocket engines, characterized by a large number of paired inclined injection orifices. The diameter and axis alignment deviation of these orifices directly influence propellant flow distribution, atomization and mixing behavior, and engine operational stability. To address the challenges associated with micro-sized orifices, inclined axes, large quantities, spatial intersection, and the low detection efficiency of conventional approaches, this paper proposes a dual-line laser 3D point cloud reconstruction-based method for measuring the diameter and impact alignment deviation of injector orifices. A dual-line laser measurement system is established to capture surface point clouds on both sides of the orifice inlets. Through system calibration and point cloud registration, the 3D point cloud data of the injector orifices within a unified coordinate system are reconstructed. Cross-sectional mapping, boundary extraction, and geometric fitting techniques are applied to determine the diameter and axis parameters of the orifices, and the spatial alignment deviation of paired orifices is subsequently calculated. To validate the feasibility of the proposed method, experimental investigation is conducted on test specimens with both 8 pairs of Φ2 mm through-holes and Φ0.5 mm micro-orifices. For the Φ2 mm specimen, the diameter measurement results are compared with industrial computed tomography (CT) data, while the alignment deviation results are verified using a combination of pin gauges and coordinate measuring machine (CMM) measurements. For the Φ0.5 mm micro-orifices, both diameter and alignment deviation results are verified using a 3D coaxial line confocal sensor. After system calibration, the fitting residuals of three Φ8 mm standard spheres are all maintained within 0.08 mm. The diameter measurement results of 8 selected Φ2 mm orifices show good overall agreement with industrial CT data: the maximum absolute deviation is 22 μm, the average absolute deviation is 15 μm, the maximum relative error is 1.09%, and the average relative error is 0.74%. The diameter and alignment deviation results of Φ0.5 mm micro-orifices show good consistency with the 3D coaxial line confocal sensor: the maximum absolute deviation is 13 μm for diameter and 0.047° for alignment deviation, with maximum relative errors of 2.41% and 0.058%, respectively. The alignment deviation results of 8 pairs of Φ2 mm orifices indicate that the proposed dual-line laser method is generally consistent with the combined pin gauge and CMM approach: the maximum absolute deviation is 0.170°, the average absolute deviation is 0.125%, the maximum relative error is 0.284%, and the average relative error is 0.125%. The results demonstrate that the proposed method enables non-contact and high-efficiency measurement of the diameter and alignment angle of injector orifices in impinging injectors for both conventional Φ2 mm orifices and micro Φ0.5 mm orifices, with high measurement accuracy and promising engineering application potential, thereby providing a new technical approach for the geometric parameter inspection of multi-scale micro-injection orifices. 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

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

(1) Background: The scarcity and imbalance of real fault data significantly limit the development of data-driven fault diagnosis methods for liquid rocket engines (LREs), especially under few-shot conditions. (2) Methods: To address this issue, this study proposes a GAN-based fault data augmentation framework for multivariate LRE time-series signals and a hybrid diagnostic classifier combining convolutional neural networks (CNNs), bidirectional long short-term memory (BiLSTM), and multi-head attention (MHA). The GAN component is introduced to alleviate fault-data scarcity and class imbalance by generating additional fault-like samples, while the classifier is designed to capture local features, long-range temporal dependencies, and diagnostically informative temporal regions. (3) Results: A multidimensional evaluation based on temporal similarity, statistical consistency, and global distribution discrepancy indicates that the generated samples preserve important characteristics of the original signals under the current evaluation protocol. On the augmented LRE dataset, the proposed classifier achieved strong diagnostic performance. In addition, supplementary experiments on the public HIT aero-engine dataset further support the effectiveness of the classifier architecture, its component-wise contribution, and its behavior under imbalanced few-shot settings, while also demonstrating the value of uncertainty-aware prediction. (4) Conclusions: The results provide encouraging evidence that the proposed framework can improve LRE fault diagnosis under data-scarce conditions. However, the present findings should be interpreted within the scope of the available data and evaluation setting. More comprehensive generator-side ablation, broader external validation, and physics-oriented assessment of the generated signals are still needed before stronger conclusions can be made. Full article

The unbalanced excitation of a rotor has a significant impact on the dynamic stiffness of the bearing. Traditional unbalanced excitation force models for the calculation of bearing stiffness are usually simplified as single-directional excitation models, which cannot fully reflect the impact of unbalanced excitation of the rotor on the dynamic stiffness of the bearing. A bidirectional excitation model based on orthogonal decomposition is used in this paper and is introduced into the finite element model of the bearing based on ABAQUS. The proposed bearing mechanics model is verified through numerical software and a bearing rotor system test rig. The effects of single/bidirectional excitation models on the dynamic stiffness of bearings were compared. The variation in bearing dynamic stiffness characteristics under rotor excitation and axial load were discussed. The results show that the presented model has good consistency with experimental results (the proposed model yields a maximum stress deviation of only 2.42% compared to MESYS numerical results and a maximum dynamic stiffness difference of 9.12% against experimental data). The traditional unidirectional excitation force model can only consider the influence of excitation frequency on the dynamic stiffness of bearings. However, the unbalanced excitation force model considering bidirectional excitation can further take into account the influence of excitation amplitude on the dynamic stiffness of bearings. Under the combined effect of excitation frequency and excitation amplitude, the radial dynamic stiffness of bearings shows a quadratic nonlinear hardening trend with rotational speed. As the rotational speed increases, the contribution of axial load to the radial stiffness significantly enhances: in the low-speed zone, its influence is only approximately 8%, while in the high-speed zone, it increases to 34%. Although the modeling method formed in this paper does not take into account the thermal–fluid dynamic coupling effect of the lubricating oil film, the obtained laws can provide a basis for the dynamic design of rotor systems of actual liquid rocket engines and have certain engineering application value. Full article

The spherical assembly mechanical end-face seal, a pivotal component of the liquid rocket engine turbopump, holds direct influence over the performance of the turbopump. This paper introduces a approach for assessing the reliability of mechanical seals. The proposed method formulates a dimensionless reliability factor, R, derived from multiple sets of monitoring data collected during the operation of the seal. The health status of the seal can be evaluated based on the value of R. The computation of R is contingent on two main elements: firstly, it relies on the threshold of evaluation parameters obtained from laboratory tests, and secondly, it incorporates the weight of parameters derived from expert experience using fuzzy set theory. And then R can be calculated by substituting real-time monitoring data of the seal. The efficacy of the proposed method was substantiated through testing and verification using two sets of real-world engineering data, and was subsequently compared with methods currently employed in engineering. The results indicate that the proposed method surpasses existing methods in terms of accuracy and sensitivity. Furthermore, the data upon which it is based can be easily monitored in an engineering context, thereby enhancing its relevance and potential for widespread application in engineering. Full article

Pintle injectors offer variable thrust capability and combustion stability advantages for liquid rocket engines. This study presents experimental and numerical investigation of spray characteristics for a liquid–liquid pintle injector using water as simulant. Ten cold flow tests covering total momentum ratio (TMR) from 0.36 to 2.76 captured spray angle variations from 26° to 80°. Computational fluid dynamics (CFD) simulations using Ansys Fluent 2025 R1 with the Volume of Fluid method and dispersed interface modeling showed good agreement with experimental spray angles for TMR > 0.74 (error < 8%), but demonstrated increasing discrepancy at lower TMR values (up to 62% error at TMR = 0.36). This deviation indicates limitations of steady-state RANS models in capturing unsteady, fuel-dominated flow regimes. The experimental dataset provides validation benchmarks for CFD modeling and contributes to injector design optimization for sounding rocket applications. Full article

The paper offers a practical method for selecting an electric pump and battery pack for low-thrust liquid rocket engines. The approach combines 0D/1D modeling of hydraulic, electromechanical, and thermal subsystems in a single environment and is supplemented by sensitivity analysis, correlation analysis, and Monte Carlo simulation with N = 3000 iterations to verify the stability of the estimates. The methodology has been tested on mission profiles in the 5–50 kN thrust range and shows that the electric pump scheme is most effective at low thrusts, while an increase in thrust leads to a disproportionate increase in energy and thermal loads and narrows the scope of applicability. The determining factors remain the hydraulic efficiency of the pump η_pump and the oxidizer pressure; electrical parameters such as bus voltage and internal battery resistance have less influence and become noticeable at high power levels. Modeling confirms the stability of the obtained estimates; at a thrust of 20 kN, the spread of the battery block mass is close to ±10%. The proposed methodology provides quantitative guidelines in the early stages of design and helps to justify the scope of application of electric pump liquid rocket engines; expansion beyond low thrust requires progress in battery technologies. Full article

One of the most widely used rocket propulsion systems that uses liquid propellants is the liquid rocket engine. Due to their high density and specific impulse, liquid fuels and oxidizers are preferred. This project’s goal was to evaluate the performance of a small-scale liquid rocket engine’s pressure feed system, which generates 250 N of thrust. A thrust bed was used for testing to obtain the experimental values, since rocket engine testing was required. A bipropellant chemical propulsion system that uses kerosene II as fuel and gaseous oxygen as an oxidant was created. Using NASA CEA, a theoretical performance evaluation was also performed in addition to this experimentation. For different propellant mixture ratios, the performance metrics under equilibrium flow, including specific impulse, characteristic velocity, thrust, and nozzle exit temperature, were examined. The performance parameters, characteristic velocity, and thrust were highest for all considered chamber pressure values and were maximum at a mixture ratio of 2.25. Full article