Solid Rocket Motor Research Articles

Investigations of high-speed projectile impact on symmetric sandwich structures containing solid propellant with core perforations

The response of solid rocket motors (SRMs) to high-speed fragment impacts is crucial for their safety design and operational use in scenarios such as rocket launches and space applications. The visualized Burn to Violent Reaction (BVR) test is used to observe intense reactions induced by high-speed projectile impacts. Employing a two-stage light gas gun and optical diagnostic techniques including high-speed schlieren imaging and direct photography, the impact-induced deflagration/explosion behavior, and reaction growth behavior were investigated. The damage mechanisms of the casing and propellant samples were assessed, and the reaction growth and afterburn effects of the impact-induced fragment cloud were quantitatively analyzed. The results indicate that the ignition delay time is inversely correlated with the impact velocity, decreasing from ms to μs scale. Across a wide range of velocities (1050–2058 m/s), higher projectile velocities induce more sustained and vigorous combustion reactions within the propellant. Furthermore, increasing the propellant air gap to 7.8 cm does not trigger further reactions under the studied configurations. The reaction mechanisms are closely linked to the characteristics of the fragment cloud induced by the impact. The developed Smoothed Particle Hydrodynamics (SPH) method, incorporating material constitutive models, ignition criteria, and reaction growth models, was used to study the influence of projectile velocity on the reaction mechanisms. The simulation results were compared with experimental data, demonstrating satisfactory accuracy.

Modeling and analysis of fluid-solid coupling heat transfer and fluid flow in transonic nozzle

Solid rocket motors (SRMs) are the power units of modern aerospace transportation, and it is of great significance to predict their flow and heat transfer characteristics by numerical simulation. This numerical study is based on a self-developed code framework MHT (Multi-X Heat Transfer, where X represents the region, scale, physical field, and other aspects) and presents an effective method of analyzing fluid–solid coupling heat transfer problem in an axisymmetric supersonic nozzle. For the nozzle wall heat conduction and nozzle-in fluid flow there are two kinds of solution methods, the detached (solved separately in solid and fluid region and them coupling at the interface) and the whole field method (solve the conduction and convection simultaneously). The whole field method is more efficient but few scholars use it to solve the nozzle coupled problem. This paper introduces the numerical details of the whole field solution method, and provides a new idea for the coupling solution of the nozzle. In addition, a mixed grid system is proposed around the interface region of fluid and solid, in which the quadrilateral structured mesh is used at the fluid side, and the triangular unstructured mesh is used for the outer fluid region and the whole solid region to ensure the accuracy of the 1st-oder normal derivatives calculation. The self-developed code is validated by comparison with well-known commercial soft-ware. Transient results of temperature and velocity are presented for the instants of 1st 20 and 40 s of a practical nozzle.

Dynamic behavior of metal droplets impacting on porous surfaces

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.

Harmonizing lightweight and ablation resistance: Design and performance of multilayer composite insulation materials for solid rocket motors

AbstractThe realization of lightweight insulation materials is often accompanied by a decrease in ablation performance. Coordinating the contradiction between lightweight and ablation resistance is the key to the development of thermal protection technology. Addressing the need for high‐performance insulation materials within solid rocket motors (SRMs) combustion chambers, we designed three multilayer composite structural insulation materials and studied their properties. The results show that among these constructed multilayer composites, the bilayer structure exhibited the lowest density at a mere 0.72 g/cm3, marking a 27% reduction compared to the basic structure. Remarkably, the bilayer configuration demonstrated superior ablation resistance, showcasing a 25% decrease in the line ablation rate in contrast to the basic structure after oxygen‐acetylene test. This achievement embodies the successful harmonization of lightweight properties with ablation resistance. Furthermore, based on the performance and microstructure of the char layer, a further analysis revealed the enhancement mechanism of ablative performance by the multilayer composite structure. It was found that the synergistic effect of the compact/porous structure of the char layer grants the bilayer structure thermal insulation material optimal ablative performance. This work provides strong support for improving the performance of SRMs.Highlights Insulation materials with multilayer composite structures are designed. The materials exhibit both lightweight and superior ablation resistance. Compact/porous char layers enhance material ablation performance.

Research on the Design and Bidirectional Work Process of Metal Diaphragms in Small Double-Pulse Solid Rocket Motors

According to the requirements of the small double-pulse solid rocket motor, a compartmentalized isolation device has been designed. This device consists of a metal diaphragm and a support frame. An experimental study and numerical simulation were used to verify the bidirectional working process of the metal diaphragm during operation of the double-pulse motor. The results show that the pressure-bearing capacity of the metal diaphragm meets the requirements under the working conditions of pulse I without affecting pulse II, because the metal diaphragm can provide insulation and flame retardancy. The metal diaphragm can be cracked in the direction of the preset V-groove in a relatively short time under the working conditions of pulse II, which allows the gas to flow to the first pulse combustion chamber normally. This indicates that the metal diaphragm can meet the requirements of bidirectional working process in dual-pulse motors.

Effects of SiO2-coated CNTs on the directional formation of SiC whiskers and improvement in the ablative resistance of polymer-matrix composites

As the development of hypersonic aerospace technology progresses, greater challenges are presented for solid rocket motors (SRMs) thermal protection, and the ablation performance of insulation materials needs to be further improved. Carbon nanotubes (CNTs) as a new type of reinforcing nano-filler, readily react with the oxidative components in the working gas during SRMs operation, limiting their excellent performance. In this study, we propose to coat the commonly used reinforcing filler, SiO2, on the surface of CNTs to suppress their susceptibility to oxidation and investigate the effects of adding CNTs, SiO2, and CNTs@SiO2 to the matrix on material properties. The results show that the addition of CNTs@SiO2 significantly improves the ablation resistance of the insulation material, with the linear ablation rate of M-@SiO2-2 being 56 % lower than that of M-SiO2-2. Based on the analysis of the material's antioxidation performance and the strength of the resulting char layer after ablation, the reasons for the improvement of ablation performance are discussed. By conducting high-temperature tube furnace tests, the composition and structure of the char layer at different temperatures are studied, and it is found that CNTs in the CNTs@SiO2 formulation can directly provide the carbon source required for the carbon thermal reduction reaction, promoting the directional growth of SiC whiskers. Based on these findings, an ablation mechanism is proposed.

Long‐Term Creep Prediction of NEPE Propellant Based on SSM Method

AbstractThroughout the entire operational lifespan of a solid rocket motor, the propellant grain experiences creep due to gravitational forces, potentially influencing the motor′s reliability. To enhance the efficiency of creep tests for solid propellants, short‐term (3000 seconds) creep tests were conducted on nitrate ester plasticized polyether (NEPE) propellant at stress levels of 0.15 MPa, 0.20 MPa, 0.25 MPa, and 0.30 MPa using both the step stress method (SSM) and the conventional constant creep test (CCT). Furthermore, mid‐term (10000 seconds) and long‐term (7 days) creep tests were conducted. Based on the traditional generalized Kelvin model, the Saint‐Venant body was incorporated to derive a visco‐elastic‐plastic model with overstress considered, and constitutive model parameters were determined through data fitting. The findings demonstrate the efficacy of the constitutive model in well characterizing propellant creep behavior. According to the model, the yield stress of NEPE propellant was determined to be 0.2302 MPa, representing a significant 64.03% reduction compared to its design strength. This insight holds relevance for structural integrity analysis and the evaluation of solid motor storage life.

Inaccurate search based sequential approximate constrained optimization method for Solid Rocket Motor design

Surrogate based constrained optimization methods are widely utilized in practical solid rocket motor (SRM) design problems. To improve the design performance, this paper proposes an inaccurate search based sequential approximate constrained optimization (ISSACO) method consisting of an efficient recursive evolution Latin hypercube design (RELHD) method and a three-stage constrained sampling method. The RELHD method divides large-sample experiment design problem into several small-sample design problems to ensure the computational efficiency and the design uniformity at the same time. The three-stage constrained sampling method employs the elite archives mechanism to balance the feasibility and the optimality in optimization. Results of the numerical cases indicate that the proposed ISSACO algorithm is competitive with other state-of-art constrained optimization method in efficiency and effectiveness. The finocyl grain design and the impulse mass ratio optimization are also conducted to make further verifications, and the ISSACO method can efficiently obtain feasible optimum and proves to be a potential method for computationally intensive solid rocket motor design problems.

Pressure oscillation and suppression method of large-aspect-ratio solid rocket motors

A two-dimensional large eddy simulation numerical model is proposed to study the transient vortex flow and pressure oscillation of a large-aspect-ratio solid rocket motor. The numerical model is validated through experimental data, finite element analysis and cumulative error analysis. The numerical simulations are executed to obtain the characteristics of the vortex-acoustic and pressure oscillation. The results show that the burning surface regression decreases the motor aspect ratio, increasing the corresponding natural frequency from 260 Hz to 293 Hz. The pressure oscillation phenomenon is formed due to the vortex-acoustic coupling. Decreasing the corner vortex shedding intensity shows negative effects on the dimensionless amplitude of the pressure oscillation. The head cavity without the injection can decrease the vortex-acoustic coupling level at the acoustic pressure antinode. The modified motor with head cavity can obtain a lower dimensionless oscillating pressure amplitude 0.00149 in comparison with 0.00895 of the original motor. The aspect ratio and volume of the head cavity without the injection have great effects on the pressure oscillation suppression, particularly at the low aspect ratio or large volume. The reason is that the mass in the region around the acoustic pressure antinode is extracted centrally, reducing the energy contribution to the acoustic system. With the volume increasing, the acoustic energy capacity increases.

Numerical Studies of Solid Rocket Propellant PMMA-PBAN-ALF3-Nitrocellulose-Difluoramine

Abstract The aim of the new investigation is the evaluation of solid rocket Propellant. In which performing numerical analysis of solid rocket propellant and calculating the constraints involved as derived from thrust i.e. velocity involved in the solid rocket propulsion motor usage. This paper aims to show the viability of solid rocket propellant for modern and future applications. Solid rocket motors are simple in design and fabricating. The need for a propellant that produces the same specific impulse as another chemical rocket engine. So it is needed to have a solid rocket motor with an enhanced combustion characteristic at a reasonable O/F ratio. For a healthy ecology, an economically reliable, highly effective, and environmentally concerned propellant is constantly suitable. As a superior alternative among the already used propellants in the sector, the paper suggested fuel PolyMethyl MethAcrylate (PMMA), PolyButadiene AcryloNitrile (PBAN) CoPolymer, and Aluminum hydride (AlH3), and oxidizers such as Nitrocellulose and Difluoramine. The desired O/F ratio for the technical criteria was found to result in a high combustion characteristic. With the features discovered through numerical and analytical research, we have suggested a design for the SRM that includes post-combustion, which enhances the reaction and creates the innovative elements of the existing SRM.

Numerical study of aluminum combustion with agglomerate size distribution in solid rocket motor

The aluminum agglomerate size distribution plays an important role in influencing the particle distributed combustion in motor, which subsequently affects the performance of solid rocket motor significantly. In this work, a three-phase model to describe distributed combustion of aluminum agglomerates is established based on the Eulerian-Lagrangian method. Then, the agglomerate size, including mono size and distributed size, is studied to reveal its effect on aluminum combustion and motor flow field. The simulated results indicate that the increasing agglomerate mono size observes the obvious decrease of the average temperature inside the motor combustion chamber, implying the low combustion efficiency of large agglomerate size. When considering the agglomerate size distribution, the size distribution mode and the mean size D43 determine the combustion efficiency together. In particular, even the mean size is similar, with different distribution mode, like skewed distribution, bimodal or trimodal distribution, the combustion efficiency and flow field parameters are nonnegligible different. However, when the size distribution mode is the same and the peak range is similar, the mean size D43 becomes the only and predominant factor as the mono size.

Prediction of Solid Rocket Motor Performance Based on Deep Learning and Ignition Experimental Data

Prediction of Solid Rocket Motor Performance Based on Deep Learning and Ignition Experimental Data

A fabrication of universal multifunctional coating with excellent adhesion on the surface of solid rocket motor insulation materials through a surface activation strategy based on multiple interaction forces

In this work, a novel universal multifunctional coating for insulation materials was prepared by using polyethyleneimine as the coating substrate, low-temperature molten glass powders and MXene as additives. Meanwhile, the universal coating was tightly adhered to the surface of the insulation materials through a surface activation strategy based on the synergistic effects of hydrogen bonding, electrostatic adsorption, and chemical bonding. The shear strength data showed that the adhesion between the universal coating and the three types of insulation material were higher than that of most commercial adhesives. In the case of EPDM, for example, the adhesion of the coating is 4.17 (epoxy glue), 1.04 (chemlok) and 1.67 (Neoprene) times that of commercial adhesives, respectively. In addition, the universal coating showed good heat resistance and thermal insulation, and formed a complete and dense char on the surface after exposure to flame. The maximum mass loss temperatures of the three coated insulation materials were 2.8, 6.5, and 9.3°C higher than the original samples, respectively. Thermal insulation tests showed that the coated insulation materials reduced the average top surface temperature by 29.68%, 40.84% and 29.06%, respectively, compared to the original samples. The results of anti-migration tests showed that the prepared universal coatings displayed generally superior anti-migration properties to the insulation materials, with equilibrium migration concentrations reduced by up to more than 80% compared to the original samples, which is better than the previous products of the same type. In addition, the application potential of the advanced multifunctional insulation materials obtained from the preparation was evaluated, resulting in an increase of more than 35% in peel strength and more than 20% in shear strength for all insulation materials with propellants. This work provides a simple and environmentally friendly generalized strategy to develop high performance insulation materials for aerospace fields.

Numerical simulation study of aluminum particle evaporation combustion process based on SPH method

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.

First applications of ultrasound technology in solid rocket propellant combustion promotion

The regulation of propellant combustion using ultrasonic waves is proposed. Under ultrasonic frequencies of 25–40 kHz, we systematically examined the combustion characteristics of Al particles, ammonium perchlorate (AP)/hydroxyl-terminated polybutadiene (HTPB) propellants, and Al-containing Al/AP/HTPB propellants. Ultrasonic treatment increased the ignition delay time of the Al particles by 48.3 %. However, it increased the burning rate of the AP/HTPB propellants by up to 26.1 % and decreased their ignition delay by 39.3 %. As the ultrasonic frequency increased, the burning rate of the solid Al/AP/HTPB propellants increased by 22.5 %, and the degree of Al agglomeration decreased, resulting in a 24 % decrease in the size of the condensed-phase combustion products. The action mechanism was analyzed in terms of the effects of ultrasonic waves on Al droplets and combustion flames. The introduction of ultrasonic waves split the Al droplets near the burning surface into smaller particles, which affected combustion efficiency. Bringing the diffusion flame closer to the burning surface affected the burning rate. These findings demonstrate that the combustion and agglomeration characteristics of solid propellants can be modified using ultrasonic techniques, providing a new method for controlling the thrust of solid rocket motors.

Evaluation of the synergistic effect of magnetic nanoparticles and Ferrocene-Based 1,2,3-Triazolyl compounds as burning rate catalysts for solid rocket Motors

This work reports the synthesis, characterization, and application of promising nanocatalysts on the thermal decomposition of ammonium perchlorate (AP). Magnetite nanoparticles were functionalized with Ferrocene-derivative groups through chemical and physical methods to obtain these composite materials. The characterization was done using TEM, FTIR, VSM, and TGA techniques. The composite material MNP-Fc2I causes a decrease in the high-temperature decomposition (HTD) of AP to 357 °C. On the other hand, the MNP-Fc1 obtained the lowest activation energy (138 kJ·mol−1), indicating that the reaction may initiate at a relatively high temperature due to the need to overcome the activation energy barrier; once it begins, it proceeds with reduced energy requirements. Furthermore, performing ion exchange with energetic anions results in a decrease in the decomposition temperature of AP for the MNP-Fc1(N3) and MNP-Fc1(DCA) catalysts.

Experimental Study on Submerged Nozzle Damping Characteristics of Solid Rocket Motor

Acoustic instabilities in solid rocket motors (SRMs) can lead to severe performance deterioration and structural damage. Nozzle damping accounts for the main acoustic dissipation source, and it is highly dependent on geometric parameters and operating conditions. This study experimentally investigated the acoustic damping characteristics of submerged nozzles in SRMs, focusing on the effects of submerged cavity dimensions, nozzle convergent angle, throat-to-port area ratio, and mean pressure variations on the longitudinal instability. The steady-state wave decay method was used to quantify the acoustic damping, and a designed rotary valve system was employed to introduce periodic pressure oscillations in the high-pressure combustion chamber. The results revealed that a larger submerged cavity would reduce the nozzle damping efficiency, with the elimination of the submerged cavity enhancing the nozzle decay coefficient magnitude by 41.9%. Furthermore, increasing the nozzle convergent angle was found to amplify acoustic wave reflection, thereby diminishing damping performance. A linear inverse relationship was observed between the throat-to-port area ratio and the decay coefficient, with a 125% increase in the ratio resulting in a 24.3% reduction in the decay coefficient. Interestingly, despite the formation of complex vortices in the submerged cavity, the mean pressure variation presented negligible effects on acoustic damping characteristics, and its damping performance is similar to a simple nozzle without a cavity. These findings provide valuable experimental data for predicting the stability of a solid rocket motor with a submerged nozzle and offer insights into the optimization of submerged nozzle designs for higher acoustic damping in SRMs.

Experimental investigation of solid propellants combustion response to pressure oscillations using advanced high-resolution diagnostics

The combustion response of solid propellants plays a crucial role in determining the combustion stability of solid rocket motors (SRMs). This study employs multiple advanced combustion diagnostic techniques, including high-speed photomicrography, two-color pyrometry, and laser shadowgraphy, to simultaneously capture mesoscale flame dynamics and burning surface regression characteristics of AP/HTPB composite propellants under pressure oscillations. This experimental configuration enables the simultaneous achievement of high temporal resolution (0.25 ms) and spatial resolution (4.9μm/pixel). Pressure oscillations are generated by loudspeakers, cover a range of frequencies (168.9 Hz to 506.7 Hz) and amplitudes up to approximately 0.44% of the equilibrium pressure. The investigated propellants exhibit varying oxidizer characteristics, including coarse (330μm) and fine (150μm) AP particle sizes, as well as different AP mass fractions (81% and 86%). Results reveal that finer oxidizer particles and higher oxidizer content both result in an increased mean steady burning rate, with the former exhibiting a more pronounced influence on the burning rate increment. Under pressure oscillations, the mean burning rate increases with increasing pressure oscillation amplitude, accompanied by a reduction in flame height, an expansion in mean flame swing angle, and an elevation in mean flame temperature. The observed frequency dependency indicates a stronger combustion response when the oscillation frequency aligns with the characteristic frequency of propellant, which is closely associated with the mean burning rate, irrespective of the amplitude of pressure oscillation. This frequency-dependent phenomenon results in the highest combustion response, characterized by a maximum mean burning rate increment exceeding 165%. The innovative experimental platform employed in this study demonstrates the capability to simultaneously capture various flame fluctuation characteristics, such as flame swinging, flame temperature, and burning rate fluctuation in a single test. This facilitates a more thorough and comprehensive analysis of combustion responses in solid propellants under pressure oscillations.

Collision probability evaluation of SRM slag during orbital lifetime

At the end of combustion, a solid rocket motor (SRM) in orbit emits combustion products with relatively large particle sizes (slag). ISO 24113:2023 “Space debris mitigation requirements” defines criteria to limit debris emission 1 mm or larger in the LEO protected region. JAXA's “Space debris mitigation standard,” JMR-003E, sets out the same requirements. Hence, JAXA has begun an improvement study on SRM slag emissions as it develops the next generation of SRMs to conserve the orbital environment by significantly reducing SRM slag larger than 1 mm. However, ensuring the complete prevention of larger SRM slag emission is technically challenging. Thus, we are alternatively working to establish a risk evaluation method and criteria based on the collision probability (Pc) of SRM slag during its orbital lifetime. Based on ground firing tests, this study estimated the number and particle size of SRM slag in orbit emitted during the Epsilon and Epsilon S launch missions. The Pc per mission during the orbital lifetime was calculated and provisionally set to 10–3 for currently operational spacecraft. This paper discusses its acceptability to JAXA.

Theoretical and experimental study on combustion response of aluminized solid propellant under acceleration effects

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.