Solid Propellant Rocket Motors Research Articles

Effect of Using Fractals as Grain Configuration on the Performance of Solid Rocket Motors

Grain design is a very crucial consideration for determining the performance of solid propellant rocket motors. A rocket’s performance and combustion characteristics (thrust profile, chamber pressure variation, propellant mass flow rate, and sliver among other important performance parameters) depend very strongly on the grain configuration. We propose fractal shapes (Koch star, Koch snowflake, Gosper, and Cesaro) as new port geometries and theoretically predict their performance using simulations based on the fast-marching method implemented in the open-source software Open-Motor. The initial thrust for fractal geometry grains was found to be significantly higher than the commonly used industry standard BATES and Finosyl ports. The fractal grains also showed a significant decrease in sliver as compared to the Finocyl port that is generally used to increase performance as compared to the circular BATES ports. Especially, The Koch snowflake port had a drastically lower sliver. These new grain geometries proposed in this work can be used for applications in which low initiation and ignition times along with high initial thrust are crucial. Moreover, these geometries can improve the performance of the solid rocket motor (SRM) owing to their better regression characteristics which promise complete utilization of the entire grain web.

Novel Solid Propellants Enabled Through In Situ Martian Perchlorates

With evidence for the native perchlorates existing in the Martian regolith, this paper examines the feasibility and performance of propellants formed from perchlorate salts reported to be present on Mars. Thermochemistry calculations indicate that the Martian perchlorate-based propellants provide less theoretical specific impulse than AP composite propellants but could still be viable propellants. Three propellants made from Martian perchlorates were manufactured and compared to a control propellant with AP as the oxidizer. Deflagration experiments were performed to obtain the burning rates as a function of pressure, with results comparable to AP baseline propellant. The propellant energy density was evaluated through bomb calorimetry. The propellant formulation with a similar perchlorate mixture to the distribution found in Martian soil was then subjected to thermal analysis, elemental analysis, and sensitivity testing to examine its combustion behavior and suitability for handling. Further characterization and development work would be needed to field these propellants, but initial conclusions indicate an in situ blend of calcium perchlorate and magnesium perchlorate could serve as a novel oxidizer for future safe, high-performing, and economical solid propellant rocket motors, offering an alternative to most current proposals for Martian ascent vehicle architectures.

Solid rocket motor propellant health monitoring based on oxide-doped curved long-period fiber grating.

A kind of curved long-period fiber grating(CLPFG) engraved by CO2 laser based on oxide-doped fiber was designed to monitor the structural integrity of propellant. The mechanical damage characteristics of the propellant were analyzed. The sensor model is constructed and the refractive index modulation characteristics of the CLPFG are analyzed. The strain coupling characteristics and the strain transfer efficiency of the interface between the CLPFG and the propellant are clarified. Propellant modules with implanted CLPFG were fabricated. The novel grating sensor has been effectively coated and structurally packaged. Conducted experiments on strain and temperature of propellant modules. The large strain measurement of propellant from 0 με to 24000 με is realized. Solved the thorny problem of large strain measurement for propellants. In addition, the temperature discrimination measurement in the temperature range of 30 ℃ to 250 ℃ can be realized. Sensor exhibit extremely high stability characteristics and has good compatibility with propellants. The sensor implantation and extraction structure has been designed to improve the survival rate of the sensor inside the solid rocket motors (SRM). Sensors can accurately measure the mechanical and thermal state parameters of propellants, providing effective data support for the health management of SRM.

Risk Assessment of Solid Propellant Rocket Motor using a Combination of HAZOP and FMEA Methods

The development of rockets in Indonesia has long been carried out by the National Aeronautics and Space Agency (LAPAN), which has now been integrated into the National Research and Innovation Agency (BRIN). The Research Centre for Rocket Technology, which is one of the centres within BRIN, has been developing solid propellant-based rockets with a various sizes and types. Solid-propellant rocket technology is commonly used because of their reliability, cost-effectiveness, and simple design. However, this technology is one of the high-risk technologies, whose failure can harm humans, damage the environment and cause huge losses to assets. As a high-risk technology, risk assessment activities must be carried out, starting from the design, manufacturing, testing and up to the launching stage. In this paper, we studied a risk assessment for general Solid-propellant Rocket Motor (SRM). SRM is basically a device that processes chemical energy in solid propellant into thrust (kinetic energy) in a container that functions as a pressure vessel. The risk assessment methods commonly used in this technology are the HAZOP or FMEA methods. The HAZOP is excellent in identifying failure modes systematically through identifying the deviation of physical process parameters but has difficulties in prioritizing the risk. The FMEA has effectiveness in understanding failure mechanisms and establishing necessary countermeasures, but for a product with a lot of components, the worksheet is also complex. By combining these two methods, integrating the superiority of each method, this research can identify modes, causes and effects of failure that may occur in SRM effective and accurately. In addition, this research also proposes corrective or preventive actions for each failure mode. As the objective of the risk assessment, results of the research can be used as input for the designers to improve their design and as inspection and surveillance objects for QC officers.

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.

A novel in-flight thrust modulation technique for SPRM: proof of concept

Thrust modulation (termination or reduction) is one of the most sophisticated topics in rocket motors. Thrust termination is employed when combustion must be terminated to ensure proper stage separation, to avoid motor explosion, to attain certain range, or for research purposes. While liquid and hybrid propellant rocket motors have the ability to extinguish thrust, thrust termination in solid propellant rocket motors (SPRMs) need the use of specialized technology. In-flight thrust termination of (SPRMs) can be achieved by sudden and sufficient increase in throat area. Increasing throat area causes high depressurization rate which consequently leads to extinguishing the rocket motor. In contrast, in-flight thrust reduction can be theoretically attained either by reducing chamber pressure to a new equilibrium level or by reducing the divergent section length. The objective of the present paper is to prove the concept of using Linear Shaped Charge (LSC) for thrust modulation; a technique novel to SPRM applications. Thrust termination concept is proved via a set of static firing tests on a standard test motor. In addition, a mathematical model is developed to predict the drop in thrust due to cutting the nozzle at any arbitrary location. The model is validated by comparing with published experiments for thrust termination and reduction. A parametric study is conducted based on model results. It is confirmed that by cutting the nozzle along its divergent section, thrust reduction less than 20% can be attained. Further reduction can be achieved by cutting the nozzle along its convergent section.

Rapid optimization method of solid rocket motor propellant with structural strength

AbstractTo design a solid rocket motor with a high mass ratio, it is necessary to find out the parameters of the grain shape to meet the structural strength under volume loading fraction and initial burning areas conditions. A rapid optimization method for complex grain of solid rocket motors based on parametric modelling and GA‐BP is proposed. Through the sensitivity analysis of the geometric parameters related to the grain, the sensitive parameters are determined. 30 groups of parameters are determined by the mixed horizontal orthogonal test and the learning samples of the neural network are obtained by FEA. The artificial neural network is established based on the FEA results, and the accuracy of the network is verified by 10 groups of test data sampled by the LHS method. The comparison results show that the prediction error of the maximum Von Mises strain is 5.09 %, the prediction error of the volume loading fraction is 0.16 %, and the prediction error of the initial burning surface is 0.865 %. Based on this, the neural network prediction results are optimized as a function of fitness in GA, and the optimization results show that the maximum strain is reduced by 20.41 % while the constraints are satisfied.

Practical Methods of Optimizing the Maximum Range for Rockets

Range extension for artillery projectiles, rockets, and missiles is a problem of trajectory optimization based on a set of constraints aimed at maximizing the range as a measure of performance. Solid Rocket Motors (SRM) are rocket motors that use solid propellants. For higher engine efficiency and cost minimization, optimization of the engine is necessary, which implies a modification of the schematics of the engine. SRM optimization is currently a key topic in aerospace engineering research. Because some input data remain constant, others can be considered as variables, and their influence on the performance parameters of the rockets was investigated. In some investigations, the time until impact is an important performance parameter; in others, the range can be considered as the performance parameter to be optimized. In this study, pulsed solid propellant rocket motor technology was considered to demonstrate a feasible method for range extension. For the targeted optimization problem, the influence of the thrust profile along the launching angles was studied, and the results from the performed calculations were analyzed and discussed.

Synthesis and characterization of iron and ruthenium complexes with the hydrotris(3,5-dimethyl-pyrazolyl)borate ligand and application as burning rate catalysts for solid rocket motor propellant

In this work, we reported two complexes [Tp*FeCp*] (1) and [Tp*RuCp*] (2) obtained from the reaction of respective pentamethylcyclopentadienyl metal precursors and [hydrotris(3,5-dimethyl-pyrazolyl)borate] (Tp*). Both complexes were characterized by 1H and 13C NMR, elemental analysis, FT-IR, UV–vis, and cyclic voltammetry. The synthesized complexes [Tp*MCp*] and analogous compounds [(Cp*)2M] (with M = Fe and Ru) were tested as catalysts on the thermal decomposition of ammonium perchlorate by a differential scanning calorimetry (DSC) technique. Compound 2 caused a decrease ammonium perchlorate’s decomposition temperature displacing it to 326 °C compared to compound 1 and [(Cp*)2M] species. However, compound 2 presented the lowest energy released compared to the complexes here reported (408 J·g−1). In addition, the results obtained from the cyclic voltammetry characterization evidenced that the electrochemically active species of compounds 1 and 2 exhibit noteworthy oxidation peaks. These results suggest that compounds 1 and 2 here presented are suitable and competitive alternatives to be used as a modifier for composite solid propellants.

Numerical Simulation of Chemical Propulsion Systems: Survey and Fundamental Mathematical Modeling Approach

This study deals with the mathematical modeling and numerical simulation of chemical propulsion systems (CPSs). For this, we investigate and summarize a comprehensive collection of the simulation modeling developments of CPSs in academic works, applications, and industrial fields. Then, we organize and analyze the simulation modeling approaches in several ways. After that, we organize differential-algebraic Equations (DAEs) for fundamental mathematical modeling consisting of the governing Equations (ordinary differential equations, ODEs) for the components and other equations derived from several physical rules or characteristics (algebraic equations or phenomenological equations, AEs) and then synthesize and summarize the fundamental structures of analytic mathematical modeling by types (liquid-propellant rocket engines, solid-propellant rocket motors, and hybrid-propellant rocket motors) of CPSs.

Transient thrust behavior of low‐aspect‐ratio rocket propulsion systems**

AbstractIn this work, the transient thrust force of the solid propellant rocket motor is numerically computed to illustrate the shortcomings of the common thrust formula in transient periods of operation. The internal ballistics of the solid rocket motor is simulated using the quasi one‐dimensional governing equations in which the finite volume method and the Roe′s scheme are utilized in calculation procedure. This is demonstrated that in transient operation of the propulsion systems, the common thrust formula may ignore some parts of physical phenomena because it has been derived based on steady state assumption, but there are some moving waves in transient operation in the combustion chamber which induce the fluctuations on the thrust force. So, the thrust‐time history must be computed by pressure force integration over the whole solid boundaries of the motor at any instant. The results show that such transient thrust behavior is more intense in low aspect ratio propulsion systems like small space thrusters.

Improved Mass Flow Rate Regulation Methods Based on Variable Frequency Control: A Case Study of Oxidizer Agent Weighing for Solid Propellants

The feeding and weighing of oxidizer agents is the key process of solid rocket motor propellant preparation, and its accuracy directly affects the burning performance of solid rocket motors. At present, the existing multi-batch feeding methods have the problem of low accuracy and high time consumption of the oxidizer agent. In this paper, an improved mass flow rate regulation method based on variable frequency control is proposed to improve accuracy and reduce time consumption. The nonlinear variation process of the mass flow rate during the opening and closing process of the air-operated pinch valve is analyzed. The periodic opening and closing frequency of the air-operated pinch valve is introduced to establish the mathematical model of the mass flow rate and frequency, and then, the model parameters are obtained through the discrete element method. The plan of the method of variable frequency regulation and the frequency parameters were determined using the multi-objective optimization method. The experiments are carried out, and the results show that compared to the existing multi-batch feeding method, optimized with the improved mass flow rate regulation methods based on the variable frequency control method, improved the feeding and weighing accuracy by 0.37% and reduced time consumption by 25.6%.

Method of Measurement of Admittance of Composite Solid Propellants Using Impedance Tube Technique

ABSTRACT Mitigation of combustion instability of solid propellant rocket motors is ever being attempted by several researchers as the problem is very complicated in nature. The acoustic admittance of composite solid propellants (AP/HTPB/RDX/Al) is experimentally investigated using Impedance tube technique. It is one of the indirect experimental techniques to determine the combustion response of the solid propellants besides T-burner. The governing equations based on the conservation of mass, momentum and energy (reactive flow is simplified to have only the axial temperature gradient) are linearized and solved to determine the acoustic admittance of the burning propellant. Different methods of analysis were explored to validate/compare the results obtained and are matching well. Some of the unreported details of the analysis are brought out. A novel way of utilizing the outer tube besides the inner tube as against the literature (wherein the inner tube is kept inside the outer tube) to maintain the chamber pressure is explored. A test conducted at fundamental acoustic mode (110 Hz) at 2 MPa at room temperature (303 K) is considered for the analysis. A rotary valve capable of operating up to a mean chamber pressure of 12 MPa is developed in this work to act as an acoustic driver for the experiments. The advantage of the impedance tube technique over T-burner is that the former is capable of measuring both the real and imaginary part of combustion response whereas the latter can only measure the real part. The imaginary part of the combustion response besides the real part is useful to predict the stability characteristics of the solid propellant rocket motor.

Rapid Optimization Method of Solid Rocket Motor Propellant Under Multiple Load Condition

In order to design a solid rocket motor with a high mass ratio, it is necessary to find out the parameters of the grain shape to meet the structural strength under volume loading fraction condition. A rapid optimization method for a complex grain of solid rocket motors based on parametric modeling and GA-BP is proposed. Based on this, the neural network prediction results are optimized as a function of fitness in GA, and the optimization results show that the maximum Von Mises strain is reduced by 9.8% while the constraints are satisfied.

A Review on the Influence of Test Bed Dynamic Characteristics on Thrust measured during Static Fire Testing of Solid Rocket Motor case

Abstract: Solid propelled rocket motors (SRMs) static firing tests are crucial tests in the aerospace industry while developing new motors as well as to ensure the quality of a motor batch. The operator can use this sort of test to determine whether the motor performance meets the project criteria by obtaining the measured "thrust versus time of burning" graphic the motor produces while burning. In this paper over all thrust measurement uncertaintyof a solid propellant rocket motor test bed is studied. During static test of solid motors requires the dynamic characteristics of the Static test configuration. To find out the magnitude of thrust oscillation, characterization of test setup and force transfer characteristics from motor case to load cell are required.

Gas-dynamic losses in the flow part of the channel charge of a solid-propellant rocket engine

Modern methods of computational fluid dynamics have been used to study flows with mass supply in axisymmetric channels of solid-propellant rocket motor charges. The studies were carried out with the aim of increasing the accuracy of predicting the intraballistic parameters for performing engineering calculations. The analysis of changes in intraballistic characteristics in the flow part of the channel charge in the classical and nozzleless solid propellant rocket motors is carried out for different rates of mass supply from the combustion surface. For an isobaric combustion chamber, a characteristic velocity profile is shown along a tubular charge path and a charge with sudden expansion. It is shown that for a steady cosine profile of axial velocity after sudden expansion of the channel, a length of more than three gauges is required. For a high-speed combustion chamber, a comparison of the axial velocity profile is carried out depending on the flow velocity under different conditions of mass supply. The tendency of the influence of flow velocity on the character of the profile is noted. It is shown that at a Mach number over 0.5, an increase in the mass flow rate from the combustion surface provides a less filled velocity profile tending to cosine. The differences between the pressure losses in the flow passage of the channel charge, calculated in the axisymmetric approximation, and the losses determined using gas-dynamic functions, are shown. An increase of the mass flow rate in the charge channel of a nozzleless solid-propellant rocket motor leads to a decrease in pressure losses at Mach number over 0.8.

On the use of high order central difference schemes for differential equation based wall distance computations

A computationally efficient high-order solver is developed to compute the wall distances by solving the relevant partial differential equations, namely: Eikonal, Hamilton–Jacobi (HJ) and Poisson equations. In contrast to the upwind schemes widely used in the literature, we explore the suitability of high-order central difference schemes (explicit/compact) for the wall-distance computation. While solving the Hamilton–Jacobi equation, the high-order central difference schemes performed approximately 1.4–2.8 times faster than the upwind schemes with a marginal improvement in the solution accuracy. A new pseudo HJ formulation based on the localized artificial diffusivity (LAD) approach has been proposed. It is demonstrated to predict results with an accuracy comparable to that of the Eikonal equation and the simulations are ≈ 1.5 times faster than the baseline HJ solver using upwind schemes. A curvature correction is also incorporated in the HJ equation to correct for the near-wall errors due to concave/convex wall curvatures. We demonstrate the efficacy of the proposed methods on both the steady and unsteady test cases and exploit the unsteady wall-distance solver to estimate the instantaneous shape and burning surface area of a dendrite propellant grain in a solid propellant rocket motor.

Drag and heat transfer of metal and oxide agglomerates in flow of combustion products of solid propellant

Modeling and simulation of the motion of agglomerates with a complex composition and shape in the flow of combustion products is of interest for problems related to design and optimization of solid propellant rocket motors. An approach is developed for modeling of unsteady processes that occur when a particle consisting of an aluminum droplet and a condensed oxide particle attached to it (aluminum droplet with oxide cap) is transported by a viscous incompressible fluid flow. Methods are proposed to take into account the sliding of an attached oxide particle over the surface of the aluminum droplet. The results of numerical computations are used to find the drag and heat transfer coefficients of a non-spherical particle at flow velocities corresponding to the formation of separation zones. The constructed model is one of the components of a more general mathematical model designed to calculate two-phase flows of combustion products based on a multi-scale approach to the simulation of two-phase flows in a complex environment.

Damage-ignition mechanism studies on modified propellant with different crosslinking density under dynamic loading

The study of high-energy and low-vulnerability propellants is important for the power performance and safety of solid propellant rocket motors. The modified split Hopkinson pressure bar (SHPB) tests are performed on two kinds of propellant with different crosslinking density to study the dynamic mechanical responses and damage-ignition mechanism. SHPB apparatus is equipped with a high-performance infrared camera and high-speed camera to capture the deformation, damage-ignition feature and temperature evolution images in the impact process. The results suggested that the mechanical responses and damage-ignition mechanism of the propellants were affected by the strain rates and crosslinking density. The damage-ignition degree is more intense and the reaction occurs earlier with the increase of strain rates. For propellant 1 with higher crosslinking density, the critical ignition strain rate is 4500 s−1. Two kinds of propellants show different ignition mechanism, i.e. crack generation, propagation and final fracture for propellant 1 while viscous shear flow for propellant 2. Meanwhile, the SEM images also reveal the difference of damage-ignition mechanism of the two kinds of propellants. Finally, the ignition mechanism under different strain rates and critical ignition strain rate of propellants are further explained by the theoretical calculation of temperature variations.

Health monitoring of high-damping viscoelastic materials using sub-resonator enriched electro-mechanical impedance signatures

This article presents a recent progress of the electromechanical impedance spectroscopy methodology for monitoring high-damping viscoelastic materials (solid rocket motor propellant for instance). It employs piezoelectric wafer active sensors (PWAS) with sub-resonators which can create additional resonance peaks for enriching the diagnostic information. In order to develop an in-depth understanding of the mechanism behind the sub-resonator sensor system, an analytical investigation is conducted first. The theoretical formulation treats the sub-resonator PWAS as a mass-in-mass model, which can be seamlessly merged into the traditional 1D constrained PWAS case. The analytical solution for the impedance signature of the new sub-resonator PWAS transducer is derived. Meanwhile, parametric case studies are carried out to analyze the influence of material damping and additional mass of the sub-resonators on the impedance signature. Moreover, numerical simulations are performed to demonstrate the effectiveness of the new sensor creating additional resonance peaks. Comparative investigations are carried out with both the conventional PWAS and the sub-resonator PWAS. EMIS damage indices based on the spectral amplitude and frequency variation features are used to quantify the material degradation. Additionally, the performance of the proposed sub-resonator PWAS is demonstrated via a material creep experiment. The article finishes with summary, concluding remarks, and suggestions for future work.