Search Results (37)

To reduce the high computational cost and lengthy design cycles of traditional solid rocket motor (SRM) development, this paper proposes an efficient surrogate-assisted multi-objective optimization approach. A comprehensive performance model was first established, integrating internal ballistics, grain structural integrity, and cost estimation, to enable holistic assessment of the coupled effects of key motor components. A parametric analysis framework was then developed to automate the model, facilitating seamless data exchange and coordination among sub-models through chain coupling. Leveraging this framework, a large-scale, high-fidelity dataset was generated via uniform sampling of the design space. The Kriging surrogate model with the highest global fitting accuracy was subsequently employed to replicate the integrated model’s complex responses and reveal underlying design principles. Finally, an enhanced NSGA-III algorithm incorporating a phased hybrid crossover operator was applied to improve global search performance and guide solution evolution along the Pareto front. Applied to a specific SRM, the proposed method achieved a 4.72% increase in total impulse and a 6.73% reduction in cost compared with the initial design, while satisfying all constraints. Full article

Fin-stabilized guided rockets exhibit ballistic characteristics such as low initial velocity, high flight altitude, and long flight duration, which render their impact point accuracy and flight stability highly susceptible to the influence of wind. In this paper, the four-dimensional nonlinear angular motion equations describing the changes in attack angle and the law of axis swing of a dual-spin rocket are established, and the phase trajectory and equilibrium point stability characteristics of the nonlinear angular motion system under windy conditions are analyzed. Aiming at the problem that the equilibrium point of the angular motion system cannot be solved analytically with the change in wind speed, a phase trajectory projection sequence method based on the Poincaré cross-section and stroboscopic mapping is proposed to analyze the effect of wind on the angular motion bifurcation characteristics of a dual-spin rocket. The possible instability of angular motion caused by nonlinear aerodynamics under strong wind conditions is explored. This study is of reference significance for the launch control and aerodynamic design of guided rockets in complex environments. Full article

The article discusses the problem of a preliminary analytical method for modifying the shape of a rocket’s nose. The purpose of this method is to determine the shape that minimizes aerodynamic drag, in the context of modifying a ballistic missile to incorporate guidance systems. The traditional design process relies on numerical methods such as CFD (Computational Fluid Dynamics) or machine learning techniques; however, the method presented here can serve as a first iteration to support the design. Advanced simulation tools are often expensive and difficult to access for smaller companies, while open-source software can sometimes be unreliable, difficult to use, and incompatible with professional solutions. This can pose a challenge for businesses planning to collaborate in the future with large corporations that rely on advanced engineering tools. The proposed solution, as previously mentioned, provides a starting point for the entire design process. The approach has been shown to be sufficient from the design work. The entire process was validated during test range trials, during which rockets were launched, and the flight measurement results accurately reflected the aerodynamic properties of the missiles. In the next stages of the project, numerical methods including CFD simulations are planned to verify the analytical results and enable further aerodynamic modification of the design. Full article

Multi-thrust solid rocket motors are extensively used in tactical missiles. To effectively achieve the desired multi-thrust performance curve, firstly, the concept of modular grain is introduced. Star grain, slot grain, and end-burning grain are chosen as the fundamental templates, which can be flexibly combined to form an arbitrary multi-thrust performance curve. Secondly, a quadric approximation of the burning perimeter is derived, leading to the establishment of a governing equation for modular grain design. This equation ensures a close match between the resulting performance curve and the target one. Thirdly, the Nelder–Mead optimization algorithm is employed to maximize the propellant loading fraction and reduce the combustion chamber size. Finally, the method successfully produces single-thrust, dual-thrust, and triple-thrust grains. The results show that the relative maximum deviation between the designed and target pressure curves is less than 6.1%. Additionally, the best grain configuration is identified, which maximizes the propellant loading fraction while adhering to the throat-to-port ratio constraints. Consequently, the concept of modular grain offers a valuable approach for creating complex internal ballistic characteristics by combining simpler grain templates. This approach allows for fast, responsive motor conceptual design, prototyping, testing, and even production, thereby advancing the development of solid rocket motors in a more efficient and effective manner. Full article

Hybrid Propellant Rocket Motors (HPRMs) have been advancing rapidly in recent years. These improvements are finally increasing their competitiveness in the global launch-vehicle market. However, some topics, such as the pre-combustion chamber design, still require more in-depth studies. Few studies have examined this subject. This work proposes a low-computational-cost algorithm that calculates the minimum pre-combustion chamber length, with a vaporization and feed-system coupled instability model. This type of analysis is a key tool for minimizing a vehicle’s size, weight, losses, and costs. Additionally, coupling with internal ballistics codes can be implemented. Furthermore, the results were compared with real HPRMs to verify the algorithm’s reliability. The shortened pre-chamber architecture trimmed the inert mass and reduced the feed-system pressure requirement, boosting overall propulsive energy efficiency by $\approx 8\%$ relative to conventional L*-based designs. These gains can lower stored-gas enthalpy and reduce life-cycle CO and CO₂-equivalent emissions, strengthening the case for lighter and more sustainable access-to-space technologies. Full article

The conventional grey relational analysis (GRA) demonstrates limitations in dynamic simulation data evaluation due to its failure to simultaneously account for geometric similarity among dynamic indicators and the proximity of data curve distances. This deficiency manifests as a compromised robustness in noise resistance and interference suppression, consequently leading to discrepancies between model accuracy and practical scenarios. To address these shortcomings, this paper proposes a robust grey relational analysis-based accuracy evaluation method (RGRA-AEM). The methodology incorporates the expected penetration rate to facilitate interpolation computation and employs deviation acceptability as a distance threshold indicator. By integrating the grey relational degree, mean squared deviation distance, and accuracy modeling, this approach achieves enhanced stability in accuracy assessment. It effectively mitigates the inherent weakness of traditional GRA that overemphasizes sequential curve similarity while significantly improving the anti-noise performance and interference resistance of grey relational coefficients. Experimental validation through the internal ballistic test-simulation dynamic data of a hybrid rocket motor conclusively demonstrates the superior robustness of the proposed methodology. Full article

Previous reviews by authors indicate the continuing development and improvement of thermal protective systems through the introduction of polymer nanocomposites into polymer matrix composites. These materials perform as thermal protective systems for a variety of aerospace applications, such as thermal protection systems (TPSs), solid rocket motor (SRM) nozzles, internal insulation of SRMs, leading edges of hypersonic vehicles, and missile launch structures. A summary of the most recent global technical research is presented. Polymeric resin systems continue to emphasize phenolic resins and other materials. New high-temperature organic resins based on phthalonitrile and polysiloxane are described and extend the increased temperature range of resin matrix systems. An important technical development relates to the transformation of the resin matrix, primarily phenolic resin, into an aerogel or a nanoporous material that penetrates uniformly within the reinforcing fiber configuration with a corresponding particle size of <100 nm. Furthermore, many of the current papers consider the use of low-density carbon fiber or quartz fiber in the use of low-density felts with high porosity to mimic NASA’s successful use of rigid low-density carbon/phenolic known as phenolic impregnated carbon ablator (PICA). The resulting aerogel composition with low-density non-wovens or felts possesses durability and low density and is extremely effective in providing insulation and preventing heat transfer with low thermal conductivity within the aerogel-modified thermal protective system, resulting in multiple features, such as low-density TPSs, increased thermal stability, improved mechanical properties, especially compressive strength, lower thermal conductivity, improved thermal insulation, reduced ablation recession rate and mass loss, and lower backside temperature. The utility of these TPS materials is being expanded by considering them for infrastructures and ballistics besides aerospace applications. 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

In a high-temperature test of the gas generator with a free-loading composite propellant, an abnormal jitter appeared in the latter part of the internal ballistic curve, whereas no such abnormality was observed in the low-temperature and normal-temperature tests. To investigate the cause, quasi-steady-state simulations of the internal flow field, as well as strength and buckling simulations of the grain, were conducted. The strength simulation revealed that the maximum stress experienced by the composite propellant during operation at 323 K is 0.7 MPa, which is lower than the ultimate stress of the grain (1.01 MPa), indicating no stress failure. The buckling simulation demonstrated that the instability arises from an imbalance of pressure on the inner and outer surfaces of the grain. In the original structure, the ventilation effect on each surface of the grain varied with the regression of the burning surface, leading to a pressure imbalance on the inner and outer surfaces of the composite propellant. Consequently, a non-ablative cladding layer was applied to ensure that the ventilation effect of each channel remains constant. The simulation demonstrated that the pressure on the surfaces of the composite propellant gradually balanced with the operation of the gas generator. Upon retesting at high temperatures, no abnormal jitter was observed in the internal ballistic curve. This indicates that maintaining a constant ventilation area for the combustion chamber and preventing changes in the ventilation effect can ensure the structural integrity of the composite propellant during operation. The working state of the composite propellant with this non-ablative cladding layer is not affected by variations in the design of the solid rocket motor. This approach enhances the adaptability and reliability of the free-loading composite propellant under different motor structures. Full article

This work proposes a new effective method to realize variable thrust through discontinuous embedded metal wires in the solid rocket motor (SRM). We aimed to study the influence of discontinuous embedded metal wires on the performance of an SRM with a single-burning-rate grain. A model based on convection heat transfer, heat conduction, and heat radiation was established to calculate the heat transfer in the discontinuous embedded metal wires in the grain, to then obtain the burning rate ratio. Most importantly, a solid rocket motor was designed to verify the feasibility of variable thrust and of the present model prediction, with the embedded silver–nickel alloy wire divided into two segments in the grain. According to the SRM ignition experiment, the silver–nickel alloy wires raised the burning rate of the grain. The pressure varied regularly with changes in the discontinuous embedded metal wires. The theoretical burning rate ratio matched the experimental result well. Based on the verified model, the effects of the burning rate, pressure exponent, burning rate ratio, and number of wires on thrust were investigated. Burning rate, burning rate ratio, and pressure exponent were found to be positively correlated with thrust ratio. The thrust ratio could reach 12.5 when the burning rate ratio was 5. The ability to adjust thrust tended to increase with an increase in the number of wires. This study also provided a method to assess whether the consecutive embedded metal wires had been broken or not. The method using discontinuous embedded metal wires in the grain was proven to be feasible to realize multi-thrusts of single-burning-rate grain, which is a new idea for the design of a multi-thrust SRM. Full article

This paper aims to study the influence of the structure of spray holes and throats on the flow field and the internal trajectory of the gas–steam ejection device. The compressible Navier–Stokes equations, discrete ordinate methods and RNG k-ε turbulence model are utilized to simulate the two-phase flow of the rocket gas with multispecies and the water sprays. The comparison between numerical results and experimental data confirms the accuracy and effectiveness of this model. The simulation analysis on the cases of the ejection process with multiple spray hole diameters, number of spray holes, total spray area, and throat diameter are conducted. The shock wave structure inside the gas–steam ejection device is examined. The simulation results show that, instead of the spray hole diameter and number, the total spray area and secondary nozzle throat diameter are the key factors that affect the flow field and internal trajectory of the gas–steam ejection device. Under the existing spray structure, the maximum number of spray holes is 300 to achieve the stability of the flow field and internal ballistic trajectory of gas–steam ejection devices. By comparing the throat diameters of multiple secondary nozzles, it was found that the minimum throat diameter of the secondary nozzles should be no less than 100 mm. The results could be valuable for the design of gas–steam ejection devices. Full article

The paper is part of the research aimed at determining if the vortex flow pancake (VFP) hybrid rocket engine is feasible as green in-space chemical propulsion. The objective of this study is to test an N₂O/HDPE VFP hybrid ignited with N₂O/C₃H₈ torch igniter. The N₂O is used in self-pressurizing mode, which results in two-phase flow and varying inlet conditions, thus better simulating real in-space behavior. The study begins with characterizing the torch igniter, followed by hot-fire ignition tests of the VFP. The results allow for the improved design of the torch igniter and VFP hybrid. The axial regression rate ballistic coefficients are reported for the N₂O/HDPE propellants in the VFP configuration. Full article

A human Mars mission is more challenging to astronauts than the Apollo mission because of travel time, life support requirements, and the space environment. Although plans for Mars exploration by NASA and SpaceX based on conventional rockets have been presented, there are considerations that suggest alternatives for the mid- or long-term. The purpose of this paper is to outline a fast mission enabled by advanced (nuclear) propulsion and by internationally shared technology. Whether the destination is the Mars surface or Phobos, for a chemical powered spacecraft, the round trip takes about 990 days, including a 480-day surface stay, compared to only 370 days, including a 41-day surface stay, for the nuclear-powered spacecraft assumed here. Since nuclear propulsion can provide higher speed than chemical, the radiation dose can be drastically reduced. The logistics of such a mission involve one or more cargo craft that must precede the astronauts. Ballistic entry into Mars’ atmosphere depends on accurate knowledge of its features, to date poorly known, that may result in uncertainty in landing coordinates. For a single vehicle, this is not critical, but for a human crew ballistic landing kilometers away from cargo is unacceptable: walking for anything but the shortest distance cannot be afforded with current space suits. In this context, the concept of a modest L/D maneuvering cargo glider based on the past Russian “Kliper” is recommended and developed to ensure landing within a hundred meters of each spacecraft. The crewed lander vehicle is based on the high L/D, inherently stable USAF FDL-7C/D hypersonic glider experience. In a similar approach, an exploration vehicle powered by in situ manufactured CO₂ and silane is described that can explore the Martian surface much faster and efficiently than with rovers or rocket-powered ‘hoppers’. Full article

Firing test campaigns were carried out on a 200 N thrust-class hybrid rocket engine, using gaseous oxygen as an oxidizer and a paraffin-wax-based fuel. Different fuel grain lengths were adopted to extend the fuel characterization under different operating conditions, and to evaluate rocket performances and internal ballistics in the different configurations. In addition to data collected under a 220 mm propellant grain length, two further test campaigns were carried out considering 130 mm and 70 mm grain lengths. Two different injector types were adopted in the 130 mm configuration; in particular, a showerhead injection system was used with the aim to contain high-amplitude pressure oscillations observed during some firing tests in this engine configuration. Parameters such as the chamber pressure and temperature inside the graphite nozzle, space-averaged fuel regression rate and nozzle throat diameter were measured. The results allowed for the investigation of different issues related to hybrid rockets (e.g., fuel regression rate, engine performance, nozzle ablation under different conditions). The focus was mainly directed to the nozzle heat transfer, through the reconstruction of the convective heat transfer coefficient for different tests in the 70 mm grain length engine configuration. The reconstruction took advantage of the experimental data provided by the nozzle embedded thermocouple. Then, the experimental convective heat transfer coefficient was used to validate the results from some empirical correlations. The results showed significant differences between the experimental convective heat transfer coefficients when considering tests with different oxidizer mass flow rates. Furthermore, the predictions from the empirical correlations proved to be more reliable only in cases characterized by oxidizer-rich conditions. Full article

Numerical studies intended to predict solid rocket motors anomalies are the major contributors when developing strategies to both limit expensive fire tests and to investigate and understand the physical phenomena from which anomalies can arise. This paper aims to present a mathematical–physical method to evaluate the pressure peak, namely Friedman Curl, occurring at the tail-off phase of small-scale rocket motors. Such phenomenon is linked to the grain solid particles arrangement (i.e., packing effect); indeed, those particles show a tendency to accumulate at a certain distance from the metallic case, implying a local burn rate increment and a combustion chamber pressure rise close to the tail-off phase. Comparisons between experimental and simulated combustion chamber pressure profiles are outlined to prove the effectiveness of the mathematical–physical approach. Simulations were carried out with an internal ballistic simulation tool developed by the authors of this work. Full article