Search Results (48)

The integrity analysis of a combustion chamber that uses Ablative Thermal Protection Systems (ATPSs) is a process that requires the analysis of the thermal and mechanical behavior of the materials involved and their interaction. A 1D thermal model for multilayered combustion chambers of hybrid rocket engines and solid rocket motors is developed, taking into consideration the thermal behavior of charring ATPSs during phase change and the capability of implementing an ablation process. A stress model is also implemented to assess the structural integrity of the combustion chamber that undergoes pressure and thermal loads. A numerical finite-difference model is used to implement analytical models and simulate the behavior of the materials. Bibliographic data and finite element analysis tools are used to evaluate and verify the models developed. Lastly, six different materials are used as a case study, and a parametric optimization is applied to obtain the minimum-mass designs using the materials selected. Full article

Liquid rocket engines occasionally experience abnormal phenomena with unclear mechanisms, causing difficulty in design improvements. To address the above issue, a data mining method that combines ante hoc explainability, post hoc explainability, and prediction accuracy is proposed. For ante hoc explainability, a feature selection method driven by data, models, and domain knowledge is established. Global sensitivity analysis of a physical model combined with expert knowledge and data correlation is utilized to establish the correlations between different types of parameters. Then a two-stage optimization approach is proposed to obtain the best feature subset and train the prediction model. For the post hoc explainability, the partial dependence plot (PDP) and SHapley Additive exPlanations (SHAP) analysis are used to discover complex patterns between input features and the dependent variable. The effectiveness of the hybrid feature selection method and its applicability under different noise combinations are validated using synthesized data from a high-fidelity simulation model of a pressurization system. Then the analysis of the causes of a large vibration phenomenon in an active engine shows that the prediction model has good accuracy, and the feature selection results have a clear mechanism and align with domain knowledge, providing both accuracy and interpretability. The proposed method shows significant potential for data mining in complex aerospace products. Full article

Accurate quantification of metallic contaminants in rocket exhaust plumes serves as a critical diagnostic indicator for engine wear monitoring. This paper develops a hybrid method combining atomic emission spectroscopy (AES) theory with a genetic algorithm (GA) optimized backpropagation (BP) network to quantify the metallic element concentrations in liquid-propellant rocket exhaust plumes. The proposed method establishes linearized intensity–concentration mapping through the introduction of a photon transmission factor, which is derived from radiative transfer theory and experimentally calibrated via AES measurement. This critical innovation decouples the inherent nonlinearities arising from self-absorption artifacts. Through the use of the transmission factor, the training dataset for the BP network is systematically constructed by performing spectral simulations of atomic emissions. Finally, the trained network is employed to predict the concentration of metallic elements from the measured atomic emission spectra. These spectra are generated by introducing a solution containing metallic elements into a CH₄-air premixed jet flame. The predictive accuracy of the method is rigorously evaluated through 32 independent experimental trials. Results show that the quantification error of metallic elements remains within 6%, and the method exhibits robust performance under conditions of spectral self-absorption, demonstrating its reliability for rocket engine health monitoring applications. Full article

Nitrous oxide is a highly suitable oxidizer for hybrid rockets due to its self-pressurizing properties, moderate cost, and high accessibility. However, its vapor pressure and density are highly dependent on ambient temperature, requiring careful consideration of temperature variations in real applications. To mitigate this issue, an oxidizer called Nytrox was produced by adding a small fraction of oxygen to bulk nitrous oxide. This modification enables the hybrid rocket propulsion system to maintain a nearly constant average thrust and total impulse across a wide range of ambient temperatures. A series of 7 s hot-fire tests of a small Nytrox/polypropylene hybrid rocket engine operating at ~60 barA of running tank pressure demonstrated a consistent average thrust of 45.3 ± 0.7 kgf and a total impulse of 307.6 ± 3.9 kgf·s within a N₂O temperature range of 5.9–22.6 °C, compared to highly varying values of the N₂O/polypropylene one within a N₂O temperature range of 10.8–29.8 °C. Furthermore, the specific impulse of the Nytrox hybrid rocket engine increases mildly with decreasing temperature because of the increasing amount of added oxygen that benefits the combustion for generating the thrust. Full article

Hybrid Rocket Engines (HREs) combine the advantages of solid and liquid propellants, offering thrust control, simplicity, safety, and cost efficiency. Part of the research on this rocket architecture focuses on optimising combustion chamber design to enhance performance, a process traditionally reliant on time-consuming experimental adjustments to chamber lengths. In this study, two configurations of HREs were designed and tested. The tests aimed to study the impact of post-chamber lengths on rocket engine performance by experimental firings on a laid-back test engine. This study focused on designing, manufacturing, and testing a laid-back hybrid engine with two chamber configurations. The engine features a small combustion chamber, an L-shaped mount, a spark ignition, and nitrogen purging. Data acquisition includes thermocouples, pressure transducers, and a load cell for thrust measurement. Our experimental findings provide insights into thrust, temperature gradients, pressure, and plume characteristics. A non-linear regression model derived from the experimental data established an empirical relationship between performance and chamber lengths, offering a foundation for further combustion flow studies. The post-chamber length positively impacted the engine thrust performance by 2.7%. Conversely, the pre-chamber length negatively impacted the performance by 1.3%. Further data collection could assist in refining the empirical relation and identifying key threshold values. Full article

This paper presents the design and experimental evaluation of a throttleable pintle-centrifugal injector system tailored for hybrid rocket engines, aimed at improving combustion efficiency and enabling precise throttling control. The novel injector system combines the principles of swirl injection and pintle-based throttling, offering fine adjustment of oxidizer flow rates to optimize combustion dynamics. Cold-flow experiments using deionized water were conducted to assess the injector’s performance across a range of flow rates and pintle strokes. Results demonstrate that the pintle stroke effectively regulates injection pressure drop and atomization characteristics, with significant improvements observed in spray cone angle and droplet size distribution. The injector system achieved a pressure drop variation ratio of 4.162 at a flow rate adjustment ratio of 6.841, indicating a strong capacity for deep throttling. These findings highlight the potential of the pintle-centrifugal injector to enhance the performance and adaptability of hybrid rocket motors, offering promising applications in modern aerospace propulsion systems. Full article

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 (I_sp) increases with the increase of the equivalent ratio. Full article

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. Full article

Engine fault diagnosis is a critical task in automotive aftermarket management. Developing appropriate fault-labeled datasets can be challenging due to nonlinearity variations and divergence in feature distribution among different engine kinds or operating scenarios. To solve this task, this study experimentally measures audio emission signals from compression ignition engines in different vehicles, simulating injector failures, intake hose failures, and absence of failures. Based on these faults, a hybrid approach is applied to classify different conditions that help the planning and decision-making of the automobile industry. The proposed hybrid approach combines the wavelet packet transform (WPT), Markov blanket feature selection, random convolutional kernel transform (ROCKET), tree-structured Parzen estimator (TPE) for hyperparameters tuning, and ten machine learning (ML) classifiers, such as ridge regression, quadratic discriminant analysis (QDA), naive Bayes, k-nearest neighbors (k-NN), support vector machine (SVM), multilayer perceptron (MLP), random forest (RF), extra trees (ET), gradient boosting machine (GBM), and LightGBM. The audio data are broken down into sub-time series with various frequencies and resolutions using the WPT. These data are subsequently utilized as input for obtaining an informative feature subset using a Markov blanket-based selection method. This feature subset is then fed into the ROCKET method, which is paired with ML classifiers, and tuned using Optuna using the TPE approach. The generalization performance applying the proposed hybrid approach outperforms other standard ML classifiers. Full article

In a worldwide scenario which sees an increasing number of small satellite launches, novel mission concepts may be unlocked providing the spacecrafts with the very precise and rapid maneuvering capability that electric thrusters cannot guarantee. In this context, chemical thrusters appear to be a possible solution. This work aimed to experimentally study and solve the problem of ignition for 10 N hybrid rockets based on hydrogen peroxide. Firstly, the study analyzed the performance of a monopropellant engine capable of functioning as a hybrid injection system. In particular, the effects of the liquid mass injected, the initial temperature, and the supply pressure on the pulsed engine performance were experimentally investigated. The injected mass showed a greater impact on the performance with respect to the starting chamber temperature and injection pressure. This thruster also showed a good potential for space applications. In the second part of the work, the objective was to find an ignition procedure that reduced propellant consumption and eliminated the need for a glow plug. This is important because the electrical power consumption in real applications significantly affects other subsystems and is undesirable for chemical engines. Different ignition procedures were tested to emphasize their respective advantages and disadvantages, and the findings indicated that the concept of pulsed preheating is feasible with only a small amount of propellant consumption, while substantially decreasing the ignition duration from approximately 45 min to a maximum of just 3 min. Finally, similar ignition procedures were adopted using different fuels. The results showed that PVC and ABS, under the same operating conditions, ignite more easily than HDPE, which requires an oxidizer consumption approximately double that of the other two fuels. Considerations about the effect of chamber pressure and oxidizer mass flow rate on engine ignition were also included. Full article

This paper extensively reviews hybrid rocket propulsion-related activities from combustion engine designs to launch tests. Starting with a brief review of rocket propulsion development history, a comparison among the three bi-propellant rocket propulsion approaches, and hybrid rocket engine design guidelines, a very thorough review related to hybrid rocket propulsion and its applications is presented in this paper. In addition to propellant choice, engine design also affects the hybrid rocket performance and, therefore, a variety of engine designs, considering, e.g., fuel geometry, swirl injection, ignition designs, and some innovative flow-channel designs are also explored. Furthermore, many fundamental studies on increasing hybrid rocket engine performances, such as regression rate enhancement, mixing enhancement, and combustion optimization, are also reviewed. Many problems that will be encountered for practical applications are also reviewed and discussed, including the O/F ratio shift, low-frequency instability, and scale-up methods. For hybrid rocket engine applications in the future, advanced capabilities and lightweight design of the hybrid rocket engine, such as throttling capability, thrust vectoring control concept, insulation materials, 3D-printing manufacturing technologies, and flight demonstrations, are also included. Finally, some active hybrid rocket research teams and their plans for flight activities are briefly introduced. Full article

This study investigated combustion characteristics of composite fuel grains designed based on a modular fuel unit strategy. The modular fuel unit comprised a periodical helical structure with nine acrylonitrile–butadiene–styrene helical blades. A paraffin-based fuel was embedded between adjacent blades. Two modifications of the helical structure framework were researched. One mirrored the helical blades, and the other periodically extended the helical blades by perforation. A laboratory-scale hybrid rocket engine was used to investigate combustion characteristics of the fuel grains at an oxygen mass flux of 2.1–6.0 g/(s·cm²). Compared with the composite fuel grain with periodically extended helical blades, the modified composite fuel grains exhibited higher regression rates and a faster rise of regression rates as the oxygen mass flux increased. At an oxygen mass flux of 6.0 g/(s·cm²), the regression rate of the composite fuel grains with perforation and mirrored helical blades increased by 8.0% and 14.1%, respectively. The oxygen-to-fuel distribution of the composite fuel grain with mirrored helical blades was more concentrated, and its combustion efficiency was stable. Flame structure characteristics in the combustion chamber were visualized using a radiation imaging technique. A rapid increase in flame thickness of the composite fuel grains based on the modular unit was observed, which was consistent with their high regression rates. A simplified numerical simulation was carried out to elucidate the mechanism of the modified modular units on performance enhancement of the composite hybrid rocket grains. Full article

This study focuses on an innovative green propellant based on paraffin, stearic acid, and coal, used in hybrid rocket engines. Additionally, lab-scale firing tests were conducted using a hybrid rocket motor with gaseous oxygen as the oxidizer, utilizing paraffin-based fuels containing stearic acid and coal. The mechanical performance results revealed that the addition of stearic acid and coal improved the mechanical properties of paraffin-based fuel, including tensile, compression, and flexural strength, under both ambient and sub-zero temperatures (−21 °C). Macrostructural and microstructural examinations, conducted through optical and scanning electron microscopy (SEM), highlighted its resilience, despite minimal imperfections such as impurities and micro-voids. These characteristics could be attributed to factors such as raw material composition and the manufacturing process. Following the mechanical tests, the second stage involved conducting a firing test on a hybrid rocket motor using the new propellant and gaseous oxygen. A numerical simulation was carried out using ProPEP software to identify the optimal oxidant-to-fuel ratio for the maximum specific impulse. Following simulations, it was observed that the specific impulse for the paraffin and for the new propellant differs very little at each oxidant-to-fuel (O/F) ratio. It is noticeable that the maximum specific impulse is achieved for both propellants around the O/F value of 2.2. It was observed that no hazardous substances were present, unlike in traditional solid propellants based on ammonium perchlorate or aluminum. Consequently, there are no traces of chlorine, ammonia, or aluminum-based compounds after combustion. The resulting components for the simulated motor include H₂, H₂O, O₂, CO₂, CO, and other combinations in insignificant percentages. It is worth noting that the CO concentration decreases with an increase in the O/F ratio for both propellants, and the differences between concentrations are negligible. Additionally, the CO₂ concentration peaks at an O/F ratio of around 4.7. The test proceeded under normal conditions, without compromising the integrity of the test stand and the motor. These findings position the developed propellant as a promising candidate for applications in low-temperature hybrid rocket technology and pave the way for future advancements. Full article

The performance of hybrid rocket engines is significantly influenced by the fuel geometry. Burnback simulations, to determine the fuel surface and fluid volume, are therefore an important tool for preliminary design. This work presents a method for the simulation of spatially constant burn-ups on arbitrary geometries. An implicit surface definition by means of a signed distance function is used to represent the fluid volume and the fuel block on tetrahedral meshes. Two methods each are used to determine the fluid volume and the burning surface. The first method is based on a direct integration of the signed distance function with the Heaviside function or the Dirac delta distribution, respectively. The second method linearly interpolates the position of an isosurface and thus reconstructs the fuel surface. Both methods are compared and validated with analytical results of four example geometries. Both calculations of the fluid volume and the calculation of the surface content with the interpolation method are characterized as first-order methods. With practicable mesh resolutions of one million computational cells, errors below two percent can be achieved. With the interpolation method, numerical meshes can also be exported for any time points of the burn. Finally, the application of the program to the fuel geometry of the Viserion hybrid rocket engine is demonstrated. Full article

The recent growth in maturity of paraffin-based hybrid propulsion systems reassesses the possibility to design an alternative Mars Ascent Vehicle (MAV) propelled by a European hybrid motor. As part of the Mars Sample Return (MSR) campaign, a Hybrid MAV would present potential advantages over the existent solid concept funded by NASA through offering increased performance, higher thermal resilience, and lower Gross Lift-Off Mass (GLOM). This study looks at the preliminary design of a two-stage European MAV equipped with HyImpulse’s hybrid engine called the Hyplox10. This Hybrid MAV utilizes the advantages inherent to this type of propulsion to propose an alternative MAV concept. After a careful analysis of previous MAV architectures from the literature, the vehicle is sized with all its components such as the propellant tanks and nozzle, and the configuration of the rocket is established. A detailed design of the primary structure is addressed. This is followed by a Finite Element Analysis (FEA), evaluating the structural integrity under the challenging conditions of Entry, Descent, and Landing (EDL) on Mars, considering both static and dynamic analyses. The outcome is a Hybrid MAV design that demonstrates feasibility and resilience in the harsh Martian environment, boasting a GLOM of less than 300 kg. Full article