Simulated Solid Rocket Motor Research Articles

Turbulent Flow Characteristics in a Model of a Solid Rocket Motor Chamber with Sidewall Mass Injection and End-Wall Disturbance

The present analytical, numerical, and experimental investigations are performed to study the flow field in acoustically simulated solid rocket motor (SRM) chamber geometry. The computational solution is carried out for a high Reynolds number and low Mach number internal flows driven by sidewall mass addition in a long chamber with end-wall disturbances. This kind of flow (transient, weakly viscous, and contains vorticity) have several features in common with a turbulent flow field. The numerical study is performed by solving the unsteady Reynolds-averaged Navier–Stokes equations along with the energy equation using the control volume approach based on a staggered grid system. The v2-f turbulence model has been implemented in the current study. A comparison of the SIMPLE and PISO algorithms showed that both algorithms provide identical results, and the computational time using the PISO algorithm is higher by about 6% than the corresponding value of the SIMPLE algorithm. A fair agreement has been obtained between the numerical, analytical, and experimental results. Moreover, the results showed that the complex turbulent internal flow patterns are induced inside the chamber due to the strong interaction of the sidewall injection with the traveling acoustic waves. Such a complex internal structure is shown to be dependent on the piston frequency and sidewall mass flux. The current study, for the first time, emphasizes the acoustic-fluid dynamics interaction mechanism and the accompanying unsteady rotational fields along with the effect of the generated turbulence on the unsteady vorticity and its impact on the real burning rate.

Laminar and Turbulent Characteristics of the Acoustic/Fluid Dynamics Interactions in a Slender Simulated Solid Rocket Motor Chamber

In this paper, analytical, computational, and experimental studies are integrated to examine unsteady acoustic/vorticity transport phenomena in a solid rocket motor chamber with end-wall disturbance and side-wall injection. Acoustic-fluid dynamic interactions across the chamber may generate intense unsteady vorticity with associated shear stresses. These stresses may cause scouring and, in turn, enhance the heat rate and erosional burning of solid propellant in a real rocket chamber. In this modelling, the unsteady propellant gasification is mimicked by steady-state flow disturbed by end-wall oscillations. The analytical approach is formulated using an asymptotic technique to reduce the full governing equations. The equations that arise from the analysis possess wave properties are solved in an initial-boundary value sense. The numerical study is performed by solving the parabolized Navier–Stokes equations for the DNS simulation and unsteady Reynolds-averaged Navier–Stokes equations along with the energy equation using the control volume approach based on a staggered grid system with the turbulence modelling. The v2-f turbulence model has been implemented. The results show that an unexpectedly large amplitude of unsteady vorticity is generated at the injection side-wall of the chamber and is then penetrated downstream by the bulk motion of the internal flow. These stresses may cause a scouring effect and large transient heat transfer on the combustion surface. A comparison between the analytical, computational, and experimental results is performed.

Convergence of Two Internal Mean Flow Solutions for Spinning Rocket Motors

In this work, two mathematical procedures are used to derive the incompressible mean flow profile in a simulated solid rocket motor that is modeled as a spinning right-cylindrical porous tube. The first approach starts with the Navier–Stokes equations and leads to a large wall-injection Reynolds number approximation. The second begins with the inviscid Bragg–Hawthorne equation, where the variation of the stagnation head is taken to reproduce the classical Taylor–Culick flow approximation for nonspinning rockets. To permit swirl to develop, the injected fluid is introduced with a finite angular momentum that is prescribed by the spinning rate of the motor. The core singularity that naturally evolves in the inviscid formulation is overcome through the use of viscous matched-asymptotic expansions. The closed-form expressions emerging from both techniques are then compared and verified using finite volume simulations of the unabridged Navier–Stokes equations. Results from all three approaches are found to be in substantial agreement over a wide range of Reynolds numbers, thus helping to validate the asymptotic treatment of the core boundary-layer analysis pursued in this and similar studies. Additional flow features, such as the vorticity and pressure, are also described and compared to the nonspinning motor case.

Ablation behavior of C/C-SiC-ZrB2 composites in simulated solid rocket motor plumes

In order to improve the ablation resistance of carbon/carbon (C/C) composites, SiC-ZrB2 was infiltrated into them by reactive melt infiltration (RMI). Their ablation behavior with a leading edge shape was investigated in the combustion environment of a simulated solid rocket motor (SRM). The results show that after introducing SiC-ZrB2 ceramic phases into C/C composites, the linear and volume ablation rate of modified leading edge decreased by 26% and 45%, respectively. The high erosion rate of the composites leading edges was mainly attributed to the synergistic effect of thermochemical ablation and mechanical erosion in a high-velocity, high particle concentration two-phase flow of SRM plume environment.

Impact of the Surface Morphology on the Combustion of Simulated Solid Rocket Motor

An advanced and intensive computational solution development is integrated with an asymptotic technique, to examine the impact of the combustion surface morphology on the generated rotational flow field in a solid rocket chamber with wide ranges of forcing frequencies. The simulated rectangular chamber is closed at one end and is open at the aft end. The upper and lower walls are permeable to allow steady and unsteady injected air to generate internal flow mimicking the flow field of the combustion gases in real rocket chamber. The frequencies of the unsteady injected flow are chosen to be very close or away from the resonance frequencies of the adapted chamber. The current study accounts for a wide range of wave numbers that reflect the complexity of real burning processes. Detailed derivation for Navier-Stokes equations at the four boundaries of the chamber is introduced in the current study. Qualitative comparison is performed with recent experimental work carried out on a two-inch hybrid rocket motor using a mixture of polyethylene and aluminum powder. The higher the percentage of aluminum powder in the mixture, the more the corrugations of the combustion surface. This trend is almost similar to the computational and analytical results of a simulated solid rocket chamber.

Numerical modeling for the combustion of simulated solid rocket motor propellant

The numerical procedure for the burning of Ammonium Perchlorate (AP) with a Fuel-Binder (Hydroxyl Terminated Polybutadience HTPB) heterogeneous propellant is presented. This model accounts for the two-step reaction mechanism for the primary diffusion flame at the interface between Binder (B) and the oxidizer AP and the AP premixed flame. The complete coupling between the gas-phase, the condensed-phase, and the unsteady non-uniform regression of the propellant surface is considered. The parameters used in this model are chosen to fit experimental data for the burning of the composite AP/Binder. The propagation of the unsteady non-planer regression surface is described, using the Essentially-Non-Oscillatory (ENO) scheme with the aid of the level set strategy. The Alternating-Direction-Implicit (ADI) solver is employed to solve the full Navier–Stokes equations in the gas phase for the variable density model. The results show the effect of various parameters on the surface propagation speed, flame structure, and the burning surface geometry. A comparison between the computational and experimental results is presented.

Silicone polymer composites for thermal protection system: fiber reinforcements and microstructures

A new class of silicone polymer matrix composites was evaluated using a simulated solid rocket motor test apparatus. Conversion of this organic silicone polymer to a ceramic (i.e. silica) structure on exposure to flame impingement or high temperature, accounts for its outstanding thermal stability. A research program was aimed to develop and evaluate this new class of thermal protection materials for military applications. This article presents the effects of the type and form of reinforcements on the thermal performance of a novel class of silicone polymer matrix composites. Reinforcement types such as glass, silica, quartz, NextelTM, and NicalonTM were used. Reinforcement forms such as random continuous-fiber mat, chopped-fiber mat, 2-D fabric, 3-D fabric, chopped roving, and broadgood tapes with different ply angles were tested. Detailed microstructural, mass loss, and peak erosion analyses were conducted on the phenolic-based matrix composite (control) and silicone-based matrix composites to understand their protective mechanisms.

Properties of steady‐state solutions for a simulated solid rocket motor

PurposeThis paper's aim is to propose a quasi‐steady numerical model of a solid rocket motor that includes the coupling of motor chamber gas dynamics with the composite solid propellant combustion.Design/methodology/approachThe paper considers a model problem of steady‐state burning of a pure monopropellant coupled with a quasi‐steady gas dynamic model of the combustion chamber. In order to simulate the time evolution as the propellant burns back with time, the flow‐field in the chamber, the burning rate and the linear response function parameters are calculated for three port diameters of a simple cylindrical geometry.FindingsIt is shown that the pressure‐coupled linear response function remains approximately constant along the propellant surface but can change very strongly as the chamber pressure rises due to increase in the burn surface.Research limitations/implicationsOnly simplified motor geometry is considered but more realistic geometries can also be analyzed using a similar approach.Originality/valueThis study is the first step in building a comprehensive fully coupled model for numerical simulation of the internal flow‐fields of solid rocket motors. In addition, it demonstrates how to use the steady‐state results to calculate linearized pressure‐coupled response of the propellant.

Vorticity generation and acoustic resonance of simulated solid rocket motor chamber with high wave number wall injection

An asymptotic technique is integrated with computational solution development to describe the generation and evolution of intense unsteady vorticity and the accompanying temperature response in a model of solid rocket motor (SRM) chamber with low Mach number, weakly viscous internal flow. The chamber considered here has a rectangular cross-section, has to equally permeable walls, closed at the head end, and is opened at the downstream end. An initially steady internal flow from steady sidewall injection is generated. Then an additional transient sidewall mass injection with different wave numbers at non-resonance and resonance frequencies is imposed on the generated steady one. This mechanism is used to simulate the transient propellant combustion occurring on the sidewall in a real SRM chamber. The analytical approach is based on the reduced form of the Navier–Stokes using asymptotic technique. The results show that unexpectedly large transient shear stresses and temperature gradients are created at the sidewall of the chamber at resonance frequencies and low wave numbers. A comparison between the analytical, computational and experimental results is performed.

Transition Characteristics of Flowfield in a Simulated Solid-Rocket Motor

The e ow characteristics in a two-dimensional porous-walled duct simulating a solid-propellant rocket motor are numerically computed to investigate the effects of viscosity, compressibility, and ine ow turbulence ( w) on the e ow transitions. The e nite volume technique is used to solve the time-dependent compressible Navier‐ Stokes equations with a subgrid-scale turbulence model, and the numerical e uxes are computed using a modie ed Godunov scheme. In addition to computed axial mean velocity and turbulence intensity proe les, the axial variations of skin friction coefe cient and the transverse location of peak turbulence intensity are used to identify the mean-e ow transition and turbulence-intensity transition, respectively. In particular, a new way of identifying turbulence-intensity transition by the use of the power spectrum of velocity e uctuations is presented for the e rst time in the present study. The minimum centerline Mach number for the onset of mean-e ow transition is obtained as the compressibility is considered alone. The critical values of w for the onset of turbulence-intensity transition and mean-velocity transition advance as well as for the concurrence and delay between the two transitions are also determined to illustrate why some researchers could observe only a single transition whereas others observed two transitions.

Study of the Flow Turning Loss in a Simulated Solid Rocket Motor

The flow turning loss, one of the processes that damp instabilities in solid rockets, is caused by the interaction of the axial acoustic field in an unstable motor with the flow of combustion products from the propellant surface. While state-of-the-art stability models generally account for the flow turning loss, its characteristics have never been fully investigated. In order to study the contribution of the flow turning loss to acoustic stability, a onedimensional acoustic stability equation that includes the flow turning loss term was derived, providing the theoretical background and expressions needed to guide an experimental study. Experiments were performed to determine the dependence of the flow turning loss upon the injection and mean-flow Mach numbers and the location of the flow turning region relative to the standing acoustic wave. These studies showed that the flow turning loss strongly depends upon the magnitude of the gas velocity at the propellant surface and the location of the flow turning region relative to the standing pressure wave. A simple numerical simulation of the experimental setup was developed and its predictions were found to agree well with the measured data. Nomenclature a = speed of sound E~ = quantity representative of the total acoustic energy in the cavity H = height of two-dimension al duct used in analysis hT = total specific enthalpy i = unit vector in the axial x direction j = unit vector in the transverse y direction L = length of duct m = mass flux rate u = axial, x direction, component of velocity V = velocity vector v = transverse, y direction, component of velocity Subscripts b = value at side wall t = average over time y = average over the y direction 0,1,2 = acoustic order