Solid Propellant Rocket Motors Research Articles

Influence of Strain on the Microstructure and Transverse Relaxation Characteristics of HTPB Coating Using Low-Field <sup>1</sup>H NMR Spectroscopy

During the storage process, the HTPB (hydroxyl-terminated polybutadiene) coating is continuously affected by the strain, and the microstructure and mechanical properties will be degraded, which will seriously affect the performance of composite solid propellant and solid rocket motor and cause great harm. In order to analyze the microstructure and transverse relaxation characteristics of HTPB coating under different strains, low-field 1H NMR tests was carried out under 0%, 5%, 10% and 15% strain conditions, and the crosslinking density and transverse relaxation parameters of HTPB coating were analyzed. The results show that, the transverse relaxation decay can be divided into two segmental mobilities corresponding to two distinct transverse relaxation times. With the increase of strain, the crosslinking density shows a decline tendency, the transverse relaxation decay amplitude slows down, and the inversion curve has a tendency to move to the right. The ratio of the fast transverse relaxation time and the peak area are much larger than the slow transverse relaxation time, and the proportion of the fast relaxation time and the peak area enlarge with the increase of the strain, while the proportion of slow transverse relaxation time is reduced. With the increase of strain, there is a transition from slow transverse relaxation to fast transverse relaxation, and there is an inverse linear relationship between crosslinking density and transverse relaxation time.

Parametric Study on Taper-ended Tubular Solid Propellant Grains

The design of the solid propellant grain is a decisive aspect of the solid propellant rocket motor performance. Tubular grain design is a favorable design since it produces a high neutral (time-invariable) thrust-time profile. However, neutrality of tubular grains deteriorates as the aspect ratio of the grain deviates from an optimum value that is dependent on the web thickness. In some cases, the undesirable phenomenon of erosive grain burning may take place. One simple solution to restore neutrality is to add taper to the ends of the grain. Loss of motor filling comes as penalty for adding these tapered ends. The grain should thus be tailored to simultaneously satisfy both desired design objectives namely, neutrality and filling. The present paper aims to address the dependence of these two design objectives on the design of a taper-ended tubular grain. The designs that are likely to yield erosive burning are also addressed. A parametric study is conducted involving the aspect ratio of the grain, its web thickness, and the taper angles on both ends. These parameters were varied, respectively, in the ranges 0.5: 5, 0.1:0.9, and 10o:90o. It was concluded that neutrality and filling of these grains are two competing objectives. The maximum neutrality is achieved using a grain with aspect ratios, web thickness, front and rear taper of 2:2.5, 0.1, 50o, and 30o, respectively. In contrast, the maximum filling can be achieved using a grain with aspect ratios and web thickness of 1:3.5 and 0.8, respectively, with no taper at both ends. A compromised grain can have aspect ratio and web thickness of 2 and 0.8, respectively, with no taper at both ends.

Study for Analysis of Effect of Machining Parameter & To Predict the Behaviour of Propellant Grain during Machining Operation

Solid Propellant Grain machining is one of the critical and hazardous operations in Solid Propellant Rocket Motor (SRM) processing. HTPB based Propellant grain surfaces are casted as per the feasible mandrel shape, but mostly the final grain surfaces are achieved by propellant machining operation, that generates various shapes and contours over casted Solid propellant grain. Some of these machined surfaces are directly exposed to the ignition and other acts as interface for further processing. These machined surfaces over propellant grain are generated by using CNC controlled Vertical Turn Mill Centre. Machining operation depends on three major machining parameters viz. speed, feed and depth of cut. Being a visco-elastic material machining of solid propellant surface is carried out with Special purpose Hollow Contouring Cutter with HSS Conical Insert. Solid Propellant in SRM is the prime mover to propel the payload to a particular range as per the mission requirement. Machining of extra propellant may lead to deviation from the target to be achieved. Due to visco-elastic behaviour of the propellant, during machining surface are compressed and after machining the machined surfaces are partially return back resultant the machined depths are varying with consecutive depth of cuts. This paper focuses on study of effect of depth of cuts on propellant grain surface keeping cutting speed and rotary table feed within a range. Study also analyzed to arrive at a conclusion to predict the machined depth over propellant grain, resulting in elimination of check cuts during propellant machining operation.

Нестационарные процессы в подкрепленном тонкостенной оболочкой вязкоупругом цилиндре при импульсном нагружении

Solid propellant rocket motor is considered as hollow viscoelastic cylinder inserted in multilayered elastic shell-like case. The material of propellant is considered to be compressible. An estimation of maximum unsteady stresses on cylinder-shell boundary and shell under growing pressure on interior or external cylindrical surface were calculated by FEM. Four corner isoparametric finite element is utilized. Numark method to integrate by time the dynamic equations is used. The problem of linear viscoelasticity have been employing of the Schapery method. `In the case of internal pressure, the possibility of tensile radial stresses on the contact surface of the propellant-shell during the transition process has been established. The dependence of the maximum contact stresses as well as circumferential stresses in the shell on the shell thickness is established. In the case of external pressure pulse, the presence of significant tensile radial stresses on the propellant-shell interface is shown. Insignificant tensile circumferential stresses in the transient wave process are possible in the shell.

Solid Rocket Motor Propellant Optimization with Coupled Internal Ballistic–Structural Interaction Approach

This research aims to optimize the geometric design of slotted propellant grains for solid rocket motors with respect to coupled internal ballistic performance and structural strength criteria. In-house codes such as a zero-dimensional internal ballistic solver and an analytical burnback solver are implemented to compute the variation of chamber pressure and the rocket thrust transiently. Structural analysis of the solid propellant is achieved by using a parametric linear viscoelastic model and a parametric cooldown heat transfer model, both of which are based on the finite element method. The transient temperature distribution data derived from the cooldown process are required inputs for the material properties to be used in the viscoelastic structural analysis. To enable an efficient optimization process, a surrogate heat transfer model that predicts the cooldown time of the system by eliminating expensive iterations is also implemented and validated. Within a coupled analysis approach, the pressure data obtained from the internal ballistic performance analysis are used for the ignition step of the linear viscoelastic analysis. The structural analysis results are evaluated by using a deterministic approach based on the margin of safety with respect to the stress and strain criteria. Finally, the optimum geometrical parameters for a slotted grain subjected to both structural and internal ballistic performance constraints are investigated through multidisciplinary optimization techniques.

Rocket motor exhaust thermal environment characterization

This study experimentally examines the heat flux to objects outside of a firing solid propellant rocket motor plume by measuring the heat flux to gages located at various locations with respect to the rocket nozzle. Because the application of interest may involve multiple motors firing simultaneously, the heat flux from multiple motors is projected based on data collected for a single motor test, and compared to the data from two, three-motor configuration test. Data showing the enhancement from three motors firing can be substantially higher than a single motor firing when the three motors are arranged in a triangular bundle, but this was not found to be the case when the three motors were arranged in a linear bundle (linear to the instrumentation). Based on results of this study, it was concluded that a material of concern which is exposed to as many as 14 motors firing simultaneously, should survive.

Design and Implementation of a Thrust Vector Control (TVC) Test System

The rocket engines are tested statically to evaluate the performance of engine based upon thrust produced. One of the most important parameters of the rocket engine static testing evaluation is to measure the thrust produced by the engine. The thrust produced is measured using a Thrust Vector Control (TVC) test system which is a structural element equipped with load cells. In this study, a load sensor system was designed to measure the propulsion performance of a solid propellant rocket motor. The forces and moments of the rocket motor with respect to the six degrees of freedom of the test system were measured during firing. It is seen that the obtained experimental results and the analysis results are compatible with each other. The designed stand is capable of measuring axial thrust and lateral (misaligned) thrust components, and the rolling moment for rocket motors producing axial thrust up to 50 [kN].

Features of heat transfer in a pre-nozzle volume of a solid-propellant rocket motor with charges of complex shapes

Local features of thermophysical processes in the channels and pre-nozzle volumes of solid-propellant rocket engines with case-bonded charges of different cross-sectional shapes are considered. The influence of the charge shape on the heat exchange in the nozzle bottom is investigated. It is shown that the value of the Nusselt number at a critical point on the multi-nozzle bottom is determined both by the charge channel form and by the geometry of the pre-nozzle volume. By processing the numerical experimental results the criterial dependences for determining the Nusselt number in the areas of local increase of heat exchange intensity are obtained. The obtained dependences are compared with the known empirical formulas [1–4]. It is found that the use of empirical relationships to estimate the Nusselt number leads to incorrect determination of the parameters of heat transfer on the armored surfaces of the charge, the nozzle covers, and the input parts of the submerged rotating nozzle.

Failure Mode Avoidance of Solid Rocket Motor Pressure Monitoring Joint Seals

The sealing joints used for pressure monitoring of solid propellant rocket motors (SRMs) of launch vehicles are very critical, as they are large in number, and leak through any of them is a single point failure mode. Identification of failure modes and its prevention is the key for reliable performance of an SRM. Failure modes are identified and the failure mechanisms of different seals in the pressure monitoring system studied through investigative tests with deliberately induced variations in the design parameters and nonconformance. Systematic analysis is carried out for the proposed designs through a failure mode effects analysis (FMEA), failure modes ranked in accordance with Risk Priority Number (RPN) and reliability of the joints worked out from the data. Design concerns are analyzed, alternate designs explored and innovative design solutions evolved. The effectiveness of the final design is brought out quantitatively by reduced RPN ratings and quantum jump in the reliability. Critical design, process and quality control parameters were identified, and procedures to ensure them evolved for failure mode avoidance.

Numerical simulation of ignition transient in solid rocket motor: a revisit

PurposeThe capability to predict and evaluate the motor pressure during each phase by means of a numerical analysis can significantly increase the efficiency of the preliminary design process with a reduction of both the motor development and operational costs. This paper aims to perform numerical simulation to analyze the ignition transient in solid rocket motor by solving Euler equation coupled with some semi-empirical correlations. These relations take into account the main phenomena affecting the ignition transient. Coupling relationships include the heat transfer of the gas to the propellant and erosive burning rate relationship.Design/methodology/approachThe current research effort divides motor into series of control volumes along the port axis, and the variation of port area, burning surface and burning rate along the port are taken into account. A set of governing equations are then solved using explicit, time-dependent, predictor-corrector finite difference method. The numerical model helps to capture and embed shock wave associated with igniter flow within the solution. Second-order artificial viscosity dampens out the numerical oscillations due to sharp gradient within the flow field. The developed computer code predicts the start-up characteristics of motor. The study also provides comparison of simulation results with in-house experimental motor.FindingsSimulations are performed with and without erosive burning to demonstrate that the flow model is a good physical approximation of motor. Numerical results calculated by this model without erosive burning are not in good agreement with experimental results. This minor discrepancy has motivated the inclusion of erosive burning in numerical model. The simulated results are then compared with the experimental data for head-end and rear-end pressure. The agreement between simulation and experiment is remarkable. In summary, major finding of this study is that unsteady quasi-one-dimensional gas dynamic model can capture the flow field in the motor during ignition transient effectively.Research limitations/implicationsUnsteady quasi-one-dimensional gas dynamic model can capture the flow field in the motor during ignition transient effectively. However, in systems where two- and three-dimensional effects are pre-dominant, one would require to develop a more elaborate, multi-dimensional model which will allow for further understanding of the flow behavior and eventually lead to modeling of rocket motors with more complex geometries.Practical implicationsThe close agreement between experimental and simulation results can be considered as forced to some degree, because the general mathematical model of erosive burning contains a free variable erosive burning exponent. However, in future, this variable can be established a priori by erosive burning tests.Originality/valueThe solid propellant ignition process consists of series of rapid events and must be completed in a fraction of a second. An understanding of the dynamics of ignition has become increasingly vital with the development of larger and more sophisticated solid propellant rocket motors. This research effort provides the simulation framework to predict and evaluate the motor pressure during each phase by means of a numerical analysis, thus significantly increasing the efficiency of the preliminary design process with a reduction of both the motor development and operational costs.

Calculation of pressure in a solid-propellant rocket motor with the use of a real dependence of the solid propellant burning rate on pressure

Five variants of calculating the burning rate of a solid propellant as a function of the pressure in a solid-propellant rocket motor are considered. Two variants of analytical expressions are proposed for approximating real dependences. In all variants, the pressure in the rocket motor can be presented by simple analytical expressions as a function of solid propellant parameters, charging conditions, and structural factors of the charge and motor.

Flow of Combustion Products Containing Condensed-Phase Particles over a Recessed Vectorable Jet Nozzle

The flow of combustion products containing condensed-phase particles over the recessed vectorable nozzle of a solid-propellant rocket motor was investigated with the use of the Reynolds-averaged Navier–Stokes equations, equations of the k–e model of turbulence, and the Lagrange approach. The fields of flows of combustion products and the mechanical trajectories of condensed-phase particles in the charge channel, the prenozzle space, and the nozzle unit of this motor were calculated for different angles of swing of the nozzle. The formation of vortices in the gas flow in the neighborhood of the downstream cover of the nozzle and their influence on the movement of particles different in size were considered.

Reliability Assessment of Solid-Propellant Rocket Motors Under Storage and Transportation Loads

A methodology for the assessment of the service life of solid-propellant rocket motors under random storage and transportation loads is presented. In storage conditions, the propellant grain is subject to random thermal environments while during transportation to vibration loads, both of which induce strain and stress. Over long periods of time, storage and transportation cause damage accumulation in the propellant. Aging mechanisms also contribute to the accumulated damage. Thermal loading is modeled as a harmonic function of time and vibratory loads of different transportation scenarios, such as ground, air, and sea transportation are used with the corresponding acceleration spectral density functions. The loading and material parameters are taken into consideration with their uncertainties. A linear viscoelastic material model is used for the propellant. Degradation of the structural capability is represented using Laheru’s cumulative damage model and the aging effect is simulated with Layton’s model. To find the stress and strain accumulated in the grain, the finite element method is used. The response surface method is used to use the surrogate mathematical models for the induced stress and strain. The limit-state functions are formed to predict failure. The instantaneous reliability is calculated within a confidence interval using a first-order reliability method.

Gas Dynamics of a Recessed Nozzle in Its Displacement in the Radial Direction

Numerical simulation of gasdynamic processes accompanying the operation of the recessed nozzle of a solid-propellant rocket motor in its linear displacement is carried out. Reynolds-averaged Navier–Stokes equations closed using the equations of a k–e turbulence model are used for calculations. The calculations are done for different rates of flow of the gas in the main channel and in the over-nozzle gap, and also for different displacements of the nozzle from an axisymmetric position. The asymmetry of geometry gives rise to a complicated spatial flow pattern characterized by the presence of singular points of spreading and by substantially inhomogeneous velocity and pressure distributions. The vortex flow pattern resulting from the linear displacement of the nozzle from an axisymmetric position is compared with the data of experimental visualization. The change in the vortex pattern of the flow and in the position of the singular points as a function of the flow coefficient and the displacement of the nozzle from the symmetry axis is discussed.

3D grain burnback analysis using the partial interface tracking method

The evolution of multi-dimensional grains of solid rocket motor propellant was carried out using the partial interface tracking method. This method applies the Lagrangian approach to the axisymmetric area of the X–Y plane and the fin area of the Y–Z plane to analyze the multi-dimensional grain burnback. The Lagrangian approach is an effective way to simulate the axisymmetric or 2D shape of a grain, but it is difficult to determine the 3D shape using this method. Therefore, the partial interface tracking method was designed to solve this problem and to get solutions quickly. The grid data resulting from this method is generated in a standard file format, which makes it easy to check the grid shape through commercial modeling software. This method showed good agreement with previous literature and was effective for numerous different grains. Also, it was confirmed that the code can be applicable to the 0D flow field model.

Experimental Investigation of Dual-Thrust Rocket Motor with Intermediate Nozzle

Dual-thrust solid propellant rocket motors are used in applications where the vehicle is boosted with a high thrust to reach very quickly a high speed that is then sustained by a low thrust. Dual-thrust rocket motors have a variety of designs; however, the one with intermediate nozzle yields high thrust ratio as well as a stable operation. Motivated by the shortage in related studies, the present paper is intended to shed more light on the topic of dual-thrust rocket motors with intermediate nozzle. The paper discusses the results of a set of static firing tests on a developed test motor covering all modes of motor operation. It is found that the thrust ratio can be maximized if the pressure difference across the intermediate nozzle is minimized. If the flow through the intermediate nozzle is initially reversed, a risk of motor damage should be taken into account.

Calculation of Pressure in a Solid-Propellant Rocket Motor with the Use of a Real Dependence of the Solid Propellant Burning Rate on Pressure

Calculation of Pressure in a Solid-Propellant Rocket Motor with the Use of a Real Dependence of the Solid Propellant Burning Rate on Pressure

Modelling and Verification of Solid Propellant Rocket Motor Operation

Modelling and Verification of Solid Propellant Rocket Motor Operation

Effects of liner properties on the stress and strain along liner/propellant interface in solid rocket motor

The aim of this study is to perform viscoelastic structural analyses to determine the effects of liner properties on the stress and strain along the liner/propellant interface in a solid propellant rocket motor. A simplified axisymmetric model composed of propellant, liner, insulation, and case was established, and nonlinear viscoelastic analyses were implemented in the finite element package ANSYS. The responses of the model under the cooling load and the ignition pressurization load were calculated. The results show that under the cooling load, the increasing of the thermal expansion coefficient (CTE) and equilibrium modulus of liner leads to the increasing of the stress along the liner/propellant interface. Under the ignition pressurization load, the stresses along the liner/propellant interface increase with the increasing of the initial modulus, and decrease with the increasing of Poisson's ratio of liner. In order to reduce the interface stress and improve the serve life of solid rocket motor, a liner with low modulus and CTE, and high Poisson's ratio should be applied between the insulation and propellant from the viewpoint of interface stress.

Correlation of parameters in the burning rate law and its influence on intraballistic characteristics of a rocket motor

Experimental data demonstrating the correlation of parameters in the power-law dependence of the burning rate of composite solid propellants on pressure are reported. The reasons for changes in the burning rate due to changes in propellant mixing conditions are discussed. The deviation of the pressure in the combustor of a solid-propellant rocket motor is analyzed with due allowance for the correlation of parameters in the burning rate law. It is shown that the relative deviation of the burning rate depends on pressure at which propellant combustion occurs. Moreover, for each propellant, there exists a pressure level at which the burning rate deviation is theoretically equal to zero, regardless of the differences in propellant compositions and properties.