Solid Rocket Motor Nozzle Research Articles

Numerical prediction of the two-phase flow and radiation effects on the thermal environment and ablation of solid rocket nozzle

The effects of two-phase flow and radiative heat transfer have a significant impact on the internal thermal environment and thermochemical ablation process in aluminized solid rocket motor (SRM). However, these two effects were not revealed systematically in previous studies. In this paper, a multiscale methodology combined radiative properties of molten alumina particles at microscale calculated by Mie theory and the two-phase flow and radiative heat transfer at macroscale calculated by Eulerian-Lagrangian method coupling with discrete ordinate method is constructed within aluminized SRM. The convection and radiation are coupled with the chemical kinetics to predict the thermochemical ablation and thermal environment. The influence of two-phase fluid flow and thermal radiation on the internal thermal environment and thermochemical ablation of a SRM nozzle is analyzed quantitatively. Comparisons are made among the methodologies that models the alumina particles as gas species, considers the effect of two-phase flow and considers the effects of two-phase flow and radiative heat transfer. The results show that the erosion rate and total heat flux at the inner wall in the convergent section increase noticeably when considering the effect of two-phase flow. Meanwhile, the erosion rate in the divergent section also increases and the thermal environment is dependent on the particle size and aluminum content. Radiative heat transfer considerably enhances the total heat flux, especially in the convergent section. On the contrary, its impact on the erosion rate is very small except for the inner wall with very low temperature.

Parametrical analysis of the strength of the nozzle of a solid fuel rocketer

The paper presents an approach to solving the problem of designing a solid propellant rocket engine nozzle using a design feature in the form of a carbon fiber insert. The design task is to select the optimal parameters of the plate thickness (plate shape) that provide the necessary strength and stability of the structure with a minimum weight. During the design process, a parametric analysis of a carbon fiber insert in the solid propellant rocket nozzle was carried out. By varying the thickness of the plate, an optimal design scheme is found that meets the given safety and stability factors. Parametric analysis of an insert plate made of CM includes modeling of the main weight and strength parameters: determination of the stress-strain state of the structure, values of natural frequencies, determination of the buckling margin, determination of the mass of the solid propellant rocket motor nozzle. Analysis of the bearing capacity of the solid propellant rocket motor nozzle with an insert plate made of CM was carried out using the finite element method using the SolidWorks Simulation software package. During the parametric analysis, two variants of the solid propellant rocket engine nozzle with and without an insert plate were considered. According to the results of the parametric analysis of the solid propellant rocket nozzle, its geometric dimensions were determined and the mass of the structure was minimized.

Uncertainty in Modeling Ablation Heat Transfer in Rocket Nozzles

The ablation of carbon cloth phenolic insulators used in solid rocket motor (SRM) nozzles involves highly complex phenomena that are difficult to accurately predict. Historical and even more modern ablation predictions rely heavily on anchoring to SRM test data to improve predictability and SRM reliability. Accelerated schedules, reductions in static SRM testing prior to flight, and a highly competitive global market are placing a substantial onus on computational capability and predictive uncertainty. Quantifying uncertainty in SRM nozzle ablation predictions is essential for motor reliability. This paper provides the details of a modern uncertainty quantification methodology applied to ablation predictions in carbon cloth phenolic insulators exposed to SRM nozzle environments. A particular historical test motor is used as a demonstration case. The system response quantities of interest are the erosion depth and char depth. A representative model and input uncertainty are provided, and a sensitivity analysis is performed to identify influential parameters. Uncertainties in the numerical models and inputs are propagated through a two-dimensional uncertainty quantification methodology using a Latin Hypercube Sampling approach. The results show that the primary sources of uncertainty in SRM nozzle thermal modeling are the heat transfer coefficient, incident radiation heat flux, char material thermal conductivity, virgin material density, char material density, char material specific heat, and pyrolysis gas enthalpy. Uncertainties in the predictions of nozzle insulation erosion and char for the test case are provided relative to the nozzle location at the 99 th percentile and 95th confidence interval. Uncertainty in the char depth is roughly ±28% along the entire axial length of the nozzle. Uncertainty in the erosion depth ranges from about ±45% for the entrance and throat regions to ±80% at the nozzle exit.

Research on thermal contact resistance between C/C and resin thermal protection materials for solid rocket motor nozzles considering real surface roughness

C/C composite and resin materials have been applied in rocket motor nozzles. This study aims to investigate the thermal contact resistance (TCR) between C/C composite and resin materials. Contact surfaces are scanned by surface morphological-measuring instrument. The TCR values are solved numerically. With the steady-state heat flux method, the TCR values are experimentally measured at different temperatures and pressures. The study results show that the TCR value between C/C composite and resin materials varies in the range of 1.8245 × 10–3 ∼7.5851 × 10–3 K·m2·W-1 at different pressures and temperatures. The average TCR error between numerical analysis and test results is 8%, which indicates a high accuracy. The developed TCR numerical prediction model can provide a reference for thermal-structural design and analysis of nozzles.

Temperature Field and Thermal Stress Analysis of Rocket Engine Composite Structure Nozzles

As one of the core components of the rocket engine, the working conditions of the nozzle are very bad. Ensuring the nozzle works normally is an important prerequisite for the normal operation of the rocket engine. To ensure the safety of the solid rocket motor nozzle in use, the temperature distribution and thermal stress of the nozzle must be preliminarily analyzed. Therefore, the three-dimensional model of an L-shaped nozzle is established. The temperature field and stress field of the structure is numerically simulated by FLUENT and Static Structural. The influence of gas temperature and material thermal conductivity on temperature distribution and thermal stress of solid structures are analyzed. The results show that the throat has the highest temperature and maximum thermal stress. When the gas temperature is 2600K, the maximum temperature of the throat is 1400K, and the thermal stress is 158MPa. When the thermal conductivity of the insulation layer is 2.8 W/(m·K), the thermal stress on the throat is 170 MPa. The conclusion can provide an important reference for the thermal protection design of the nozzle.

Preliminary experimental study on combustion characteristics in a solid rocket motor nozzle based on the TDLAS system

Preliminary experimental study on combustion characteristics in a solid rocket motor nozzle based on the TDLAS system

Experimental research on the high‐temperature compression mechanical properties of carbon/phenolic composites in rapid carbonization conditions

AbstractThe structural characteristics of fiber‐reinforced plastic composites are usually determined by ablation, but bearing capacity is also a key parameter that should be clarified in aerospace engineering. Needle‐punched carbon/phenolic composites are most commonly used as the inner material of solid rocket motor nozzles, and the mechanical properties thereof are dynamic and difficult to obtain at high temperatures. A radiation‐heating mechanical testing machine with split‐test chambers was designed to test the in‐plane compression properties of the composite at different degrees or pyrolysis and temperatures. The results show that the final weight‐loss rate was 25.5%. During the loading process, the composite had a linear elastic mechanical response and brittle fracture failure, and the compressive strength was 218.9 MPa at room temperature. At different holding times, the higher the temperature, the higher the weight‐loss rate, the higher the graphitization degree of the specimen, and the lower the degree of disorder of the carbon matrix. Thus, the minimum compressive strength was found to be 23.42 MPa at 600°C, and the maximum was 60.16 MPa at 1800°C after high‐temperature pyrolysis. The present work provides a reference for the refined design of solid rocket motor nozzles.

Characterization of the Nozzle Ablation Rate Based on 3D Laser Scanning System

The nozzle of solid rocket motor (SRM) is easily ablated by high temperature, high pressure, high speed, and corrosive particles, which affects the stability of rocket flight. Therefore, the measurement and characterization of the nozzle ablation rate are helpful in providing some guidance for the design of nozzle material and structure. However, due to the high surface roughness of the composite nozzle after ablation, it is difficult to obtain an accurate ablation rate by contact measurement methods. The 3D laser scanning system is a 3D non-contact measurement technology using structured light technology, phase measurement technology, and computer vision technology. It has the advantages of non-contact, large scanning size, flexibility, and portability. In this paper, a 3D reconstruction of the ablation nozzle is carried out based on the 3D laser scanning system. Additionally, the ablation rate of the nozzle is measured without cutting the actual specimen. Furthermore, the pressure, temperature, and surface convective heat transfer coefficient trends are numerically calculated and compared with the ablation rate. Additionally, the empirical formula between ablation rate and pressure, temperature, convective heat transfer coefficient is obtained empirically by the inversion analysis method. The empirical formula can provide theoretical guidance for nozzle size design and optimization. The results show that the non-contact 3D laser scanning system is a valuable method for reconstructing the model of the ablated nozzle, and the empirical formula of ablation rate can accurately predict the ablation rate of the nozzle.

Numerical simulation of particle erosion in the internal flow field of solid rocket motor nozzle using Rocstar

Solid composite propellants often contain metal particles, such as aluminum particles, and the internal flow field in a solid rocket motor is a typical gas–solid two-phase flow field. In this study, to address the problem of the erosion of particles on the wall during the working process of an engine, a particle erosion model was established based on the interaction mechanism between the particles and the wall. From the perspective of particle dynamics, the interaction mechanism between the moving particles and the wall was analyzed. The force of the particles on the wall surface was obtained, and a formula for solving the particle erosion rate was deduced by using the strength theory and the energy conservation principle. It was combined with Hertz’s contact theory to analyze the factors influencing the erosion rate. In this paper, the particle erosion experiments revealed not only the relationship between erosion rate and erosion angle but also the correlativity between erosion rate and particle concentration. In addition, the experimental results are consistent with the deduced results, which proves the reliability of the model established in this paper. The multiphysics simulation program Rocstar, developed by the Center for Simulation of Advanced Rockets at the University of Illinois, was used to perform a numerical simulation of the two-phase flow in the Jet Propulsion Laboratory nozzle. The results agreed well with the literature and the experimental data. The feasibility and accuracy of the calculation of the gas-phase and two-phase flow fields of the solid rocket motor nozzle were confirmed, laying the foundation for the numerical simulation of the entire process of the full flow field of the engine.

Study on heat transfer model of roughness wall in supersonic two-phase flow of solid rocket motor

In order to establish the heat transfer model of solid rocket motor nozzle expansion section based on roughness wall, an appropriate heat transfer model is proposed, and the heat transfer model is established by the combination of experiment and numerical calculation. Results demonstrate that: 1) with the increase of the depth of temperature measuring points in different sections, the temperature increases and the corresponding time response accelerates. With the increase of the expansion ratio of temperature measuring sections, the temperature decreases. 2) The overall trend between the calculation results and the experimental measurement results is relatively consistent, which indicates that the numerical calculation method established in this paper has certain applicability and can accurately reflect the heat transfer characteristics of solid rocket motor nozzle expansion section based on real roughness wall. 3) The establishment of heat transfer model is completed using the segmented method, and the error between Nu calculated by the proposed heat transfer model and Nu calculated by numerical calculation is controlled within 21% and 13% in the first and second segments respectively. It shows that the heat transfer model established in this paper has certain rationality and reliability.

Effect of operation pressure on heat release characteristics in solid rocket motor nozzle considering detailed chemical reaction mechanism

In order to obtain the effect of operation pressure on the heat release characteristics in the nozzle of solid rocket motor, and provide an accurate gas action boundary for the establishment of dynamic ablation model of flexible nozzle, numerical simulation is conducted to quantitatively analyze the effect of operation pressure on the heat release characteristics in the nozzle of solid rocket motor in terms of reaction increase rate and reaction progress in this paper. Results demonstrate that: 1) based on the motor configuration and propellant, the change of operation pressure has no effect on the position of the dividing point between equilibrium state and nonequilibrium state, which is located at the convergence section 0.13 m from the nozzle inlet. 2) With the increase of operation pressure, the reaction increase rate and reaction progress of temperature and H2O at the corresponding axial position in the mainstream area increase as a whole, while the reaction increase rate and reaction progress of CO2 decrease as a whole. 3) Compared with the mainstream, the reaction increase rate and reaction progress of temperature and CO2 at the wall increase as a whole, while the reaction increase rate and reaction progress of H2O decrease as a whole.

Effect of cure temperature on the thermal degradation, mechanical, microstructural and moisture absorption behavior of vacuum‐only, carbon‐fiber reinforced phenolic composites

AbstractPhenolic resin is the material of choice used for composite liners in solid rocket motor nozzles. We investigate the effect of the cure conditions (94, 116, and 155°C) on vacuum only (VO) processed composite mechanical, thermal, and microstructural changes. Although the Tg's correlated with the processing conditions, mechanical properties showed unexpected non‐linear degradation after 94°C processing. This difference is primarily due to the lower cure state of the resin, which results in poor fiber/matrix coupling and is further magnified by the reduced degree of consolidation observed in VO parts. Elevated thermal exposure was shown to generate larger tensile strains in the out of‐plane direction for the 94°C composites when compared to the other conditions. This resulted in significant increases in crack density, porosity, and delamination due to an increased volume of degradation by‐products, while the higher temperature cured composites maintained good consolidation. Even though all three cure states appeared to display equivalent moisture absorption saturation levels based on weight measurements, the 94°C cured sample was shown to partially dissolve in moisture during humidity exposure, which translated to even higher saturation levels. This larger net absorption resulted in a larger drop in Tg for the 94°C cured material and thus resulted in a more severe thermal mechanical response than the other saturated specimens. This emphasizes how a reduction in cure state will not only affect thermal and mechanical performance but will also translate to a greater degree of degradation than more highly cured specimens, when exposed to humid environments.

A Numerical Simulation Study for A Dual Thrust Solid Propellant Rocket Motor Nozzle

CFD simulation through the rocket motor nozzle is an important issue as it gives a better understanding of the flow behavior and shows the evolution of the parameters along with the nozzle. CFD simulations using ANSYS FLUENT were conducted to model the flow through an experimental dual thrust solid rocket motor nozzle at two pressure values representing the two pressure levels. A two-dimensional axisymmetric numerical simulation was carried out using a pressure-based solver to solve the governing equation and the standard k-ε model to model the turbulence through the motor nozzle. The hot gas properties used in the simulation model were calculated using the CEA software (chemical equilibrium with applications). The parameters marched out from simulations were used to calculate the thrust values for each case. The results calculated through the numerical models compared with the experimental measurements and showed an acceptable level of accuracy.

Coupled Thermo-Structural Analysis Model of Solid Rocket Motor Nozzle considering the Variation of Friction Coefficient under Operating Conditions

Multi-interface is a typical feature of solid rocket motor nozzles, and the interface evolution law is an important aspect to analyze the thermo-structural response of solid rocket motor nozzles. In this paper, based on the friction coefficient test at different temperatures, a coupled thermo-structural analysis model of solid rocket motor nozzle considering the variation of friction coefficient under operating conditions was established. Firstly, the friction coefficient was tested to model the variation at different temperatures. Then, by adopting structure gap, variable friction coefficient, thermal contact resistance, and friction heat production, a strongly coupled non-linear model was established. Simulations using the non-linear model and traditional model were performed, in which the stress of the throat insert was within the required stress range of the material. The ground firing test result demonstrated the validity of the analysis model, and the non-linear model was in a better agreement with the firing test than the traditional model. Therefore, it can be concluded that with a more specific friction coefficient to represent the friction behaviour of the different parts of the nozzle, the strongly coupled non-linear model established in this paper can reflect the essence of thermo-structural response of solid rocket motor nozzle.

Exploration of a new affordable thermal protection system utilizing 2.5D silica/polysiloxane composite

Ablative materials are used as thermal protection systems (TPS) for reentry vehicles and solid rocket motor (SRM) nozzle applications. Phenolic and cyanate ester are the state-of-the-art (SOTA) resin systems used in many of the ablative composites today, including MX-2600 (silica/phenolic) from Cytec Solvay Group. While these ablatives have worked well, more demanding requirements drive the need for affordable lightweight advanced composites capable of handling high heat fluxes with minimal mass loss. These advanced ablative composites result in lighter reentry heat shields and solid rocket motors, increasing payload capabilities of spacecraft and rockets. Molding compound made of aerospace grade 99% SiO2 fabric and polysiloxane resin showed considerable improvement over MX-2600 in ablation properties in recent studies. In order to meet increased mechanical strength demands, NASA recently developed an ablative composite using a 3D quartz woven/cyanate ester composite material designed for the Orion spacecraft. While 3D woven composites provide excellent out-of-plane mechanical and ablation properties, they are very expensive, which limits their application. This research explores needle-punched silica fabric, sometimes referred to as 2.5D, which provides similar out-of-plane mechanical benefits to 3D woven composites in a more flexible VARTM manufacturing process at a much lower cost. The needle-punched silica fabric was infiltrated with polysiloxane resin and mechanical tests were performed. The needle-punched composites showed an increase of 181% in flexural strength, 27% in interlaminar shear strength, 2% in tensile strength, and 13% in compressive strength. In aerothermal ablation tests, the 2.5D out-performed the 2D laminate in char yield, mass loss, and recession rate; and in char yield and mass loss (%), the 2.5D out-performed the industry standard MX-2600 molding compound. The increased out-of-plane strength and char yield make it a promising and affordable ablative candidate for ablation performance with enhanced mechanical properties.

Numerical Simulation Coupling on Fluid Flow and Heat Transfer in Nozzle and Equipment Cabin of Solid Rocket Motor during Operation Process

Driven by the goal of studying the temperature changes in the nozzle and equipment cabin of solid rocket motor during the whole process of ignition starting, working, cooling and decompression, this paper uses the fluid dynamics software which uses the RNG k-ε turbulence model to simulate the mixing of the fuel and oxidant and the Sutherland law to describe the viscosity of gas and air, and the controlling equation is the two-dimensional axisymmetric viscous compressible URANS equation. In this paper, the temperature field of the nozzle and equipment cabin in the working process of solid rocket motor during 0 ∼ 60s is numerically simulated, and the ground experiment is carried out. The results show that (1) the numerical simulation results are in good agreement with the ground test results, and the absolute value of the maximum error appears at the time of 60s, which is 5.9k. indicating that the numerical calculation method and heat source equivalent method established in this paper are suitable for the temperature field simulation of composite structure nozzle and equipment cabin, (2) the unsmooth surface of the nozzle formed by high temperature ablation makes the flow situation in the divergent section of the nozzle complex and affects the heat transfer coefficient of the wall contacting with the equipment cabin, (3) due to the insulation design of the equipment cabin, only a small part of the heat is transferred to the equipment cabin. During the whole working process of solid rocket motor, the temperature field of equipment cabin increases sharply at both ends and gently in the middle.

Thermal Structure Strength Analysis of Nozzle of Solid Rocket Motor with the Coupled Algorithm

In order to achieve the thermal structural integrity analysis of the solid rocket motor nozzle accurately and efficiently, the multifield (flow-thermal-mechanical) coupled numerical investigation was carried out based on the mesh-based parallel code coupled interface. The numerical simulation process and finite element model of the coupled algorithm and engineering algorithm were obtained, while the physical model was simplified appropriately. The coupled interface parameters, internal flow field, temperature field, and stress field of the coupled algorithm were compared with the engineering algorithm results, and the effectiveness and accuracy of the numerical simulation were validated. The numerical investigations shown that both the temperature field and stress field obtained by the coupled algorithm were slightly lower than which obtained by the engineering algorithm. These were considered to be impacted by the Bartz empirical formula and the one-dimensional isentropic flow assumption. Further experimental investigations shown that the exterior surface temperature and strain of the nozzle throat obtained by the coupled algorithm were much closer to the experimental results, which further verified the accuracy of the coupled algorithm.

Investigation of Liquid Droplet Flow Behavior in a Vertical Nozzle Chamber

Abstract This study aims to better understand the aluminum oxide agglomerates breakup mechanism, consequently determining the best solution for the solid rocket motor (SRM) nozzle erosion problem. Two-phase air-water flow experimental investigation was conducted as a substitute for liquid aluminum agglomerates and exhaust combustion gases. The results show that increasing the exhaust air velocity enhances the droplet's breakup tendency to reduce the average diameter and increase droplet numbers per the testing channel volume. Numerical models were constructed and validated using the experimental results. The percentage error in the droplets’ average diameter and the number is between 6 and 15% and 8 and 18%, respectively. Furthermore, the effect of reducing the liquid surface tension was studied. The results showed that it facilitates water bodies’ separation from the interface surface, because of the reduced bounding forces between surface’s molecules, which enhances the breakup process (0.5–17% increase in the droplets’ average diameter and 4–100% increase in its number) and reduces the droplets impact on the nozzle walls, hence reducing the SRM nozzle erosion problem.

Research on thrust vector control of nonlinear solid rocket motor nozzle based on active disturbance rejection technology

In this paper, based on the auto disturbance rejection control technology, the electromechanical servo system is used as the control actuator, and the thrust vector control of the solid rocket motor nozzle with typical nonlinear friction characteristics is studied and analyzed. In this paper, the realization of the classical PID algorithm and the lack of dynamic performance are analyzed, and then the compensation algorithm based on the auto disturbance rejection control technology is added. The algorithm compensates for the phase lag of the system due to the nonlinear friction characteristics. As a result, the frequency characteristics of the system have been significantly improved.

Radiative Heat Transfer in Solid Rocket Nozzles

The radiative heat transfer within solid propellant rocket motor nozzles due to the emission and isotropic scattering by the liquid-phase alumina particles present in the combustion products is investigated. A method is presented for computing the effective absorption and scattering coefficients. The discrete ordinates method participating media radiation solver in the STAR-CCM+ code is used to compute the radiative heat flux to the nozzle wall for the HIPPO and TM-3 motor test cases. Multiple studies are performed to investigate the impact that various factors have on these calculations, including complex index of refraction, particle size distribution, exit plane radiation boundary condition, and the dimensionality of the flow and radiation solutions. The effects of boundary layers and coupling between the flow and radiation solutions are also investigated. Comparisons are made between these high-fidelity radiation solutions and simpler correlations historically used to estimate radiative heat flux in rocket nozzles; it is found that the most common emissivity correlation underpredicts radiative heat flux. Recommended methods are presented that should yield radiative heat flux predictions for rocket nozzles with increased accuracy.