Showing papers on "Solid-fuel rocket published in 2000"

Solid propellant chemistry, combustion, and motor interior ballistics

Abstract: This volume brings together international scientists in the field of solid rocket propulsion. Thirty-nine papers present in-depth coverage on a wide range of topics including: advanced materials and non-traditional formulations; the chemical aspects of organic and inorganic components in relation to decomposition mechanisms, kinetics, combustion and modelling; safety issues, hazards and explosive characteristics; and experimental and computational interior ballistics research, including chemical information and the physics of the complex flow field.

Structural Oscillation Considerations for Solid Rocket Internal Ballistics Modeling

Abstract: *† ‡ § The coupling of structural vibrations with the nonlinear internal ballistic flow and burning rate of a sleeved cylindrical-grain motor is investigated through numerically simulated motor firings configured for the evaluation of pulse-triggered combustion instability behavior. The predicted results illustrate the significant impact of the threedimensional structural acceleration field on the burning rate and axial wave development during unsteady motor operation. Parameters such as the structural damping have been found to significantly influence the axial shock wave development and base chamber pressure rise during simulated firings. Instability-related symptoms are demonstrated through this study to be dependent at least in part on the motor structural vibrations.

Virtual prototyping of solid propellant rockets

Abstract: Researchers seek a detailed, whole-system simulation of solid propellant rockets under normal and abnormal operating conditions. A virtual prototyping tool for solid propellant rocket motors based on first principles models of rocket components and their dynamic interactions meets this goal. Given a design specification (geometry, materials, and so on), we hope to predict the entire system's resulting collective behavior with sufficient fidelity to determine both nominal performance characteristics and potential weaknesses or failures. Such a response tool could explore the space of design parameters much more quickly, cheaply, and safely than traditional build-and-test methods. The article is a progress report on a project to design such a virtual prototyping tool for SRMs (solid rocket motors).

Experimental and numerical study of aeroacoustic phenomena in large solid propellant boosters

Abstract: The present research is an experimental and numerical study of aeroacoustic phenomena occurring in large solid rocket motors (SRM) as the Ariane 5 boosters The emphasis is given to aeroacoustic instabilities that may lead to pressure and thrust oscillations which reduce the rocket motor performance and could damage the payload The study is carried out within the framework of a CNES (Centre National d'Etudes Spatiales) research program Large SRM are composed of a submerged nozzle and segmented propellant grains separated by inhibitors During propellant combustion, a cavity appears around the nozzle Vortical flow structures may be formed from the inhibitor (Obstacle Vortex Shedding OVS) or from natural instability of the radial flow resulting from the propellant combustion (Surface Vortex Shedding SVS) Such hydrodynamic manifestations drive pressure oscillations in the confined flow established in the motor When the vortex shedding frequency synchronizes acoustic modes of the motor chamber, resonance may occur and sound pressure can be amplified by vortex nozzle interactionOriginal analytical models, in particular based on vortex sound theory, point out the parameters controlling the flow-acoustic coupling and the effect of the nozzle design on sound production They allow the appropriate definition of experimental testsThe experiments are conducted on axisymmetric cold flow models respecting the Mach number similarity with the Ariane 5 SRM The test section includes only one inhibitor and a submerged nozzle The flow is either created by an axial air injection at the forward end or by a radial injection uniformly distributed along chamber porous walls The internal Mach number can be varied continuously by means of a movable needle placed in the nozzle throat Acoustic pressure measurements are taken by means of PCB piezoelectric transducers A particle image velocimetry technique (PIV) is used to analyse the effect of the acoustic resonance on the mean flow field and vortex properties An active control loop is exploited to obtain resonant and non resonant conditions for the same operating pointFinally, numerical simulations are performed using a time dependent Navier Stokes solver The analysis of the unsteady simulations provides pressure spectra, sequence of vorticity fields and average flow field Comparison to experimental data is conductedThe OVS and SVS instabilities are identified The inhibitor parameters, the chamber Mach number and length, and the nozzle geometry are varied to analyse their effect on the flow acoustic couplingThe conclusions state that flow acoustic coupling is mainly observed for nozzles including cavity The nozzle geometry has an effect on the pressure oscillations through a coupling between the acoustic fluctuations induced by the cavity volume and the vortices travelling in front of the cavity entrance When resonance occurs, the sound pressure level increases linearly with the chamber Mach number, the frequency and the cavity volume In absence of cavity, the pressure fluctuations are damped

Rocket motor joint construction including thermal barrier

Abstract: A thermal barrier for extremely high temperature applications consists of a carbon fiber core and one or more layers of braided carbon fibers surrounding the core. The thermal barrier is preferably a large diameter ring, having a relatively small cross-section. The thermal barrier is particularly suited for use as part of a joint structure in solid rocket motor casings to protect low temperature elements such as the primary and secondary elastomeric O-ring seals therein from high temperature gases of the rocket motor. The thermal barrier exhibits adequate porosity to allow pressure to reach the radially outward disposed O-ring seals allowing them to seat and perform the primary sealing function. The thermal barrier is disposed in a cavity or groove in the casing joint, between the hot propulsion gases interior of the rocket motor and primary and secondary O-ring seals. The characteristics of the thermal barrier may be enhanced in different applications by the inclusion of certain compounds in the casing joint, by the inclusion of RTV sealant or similar materials at the site of the thermal barrier, and/or by the incorporation of a metal core or plurality of metal braids within the carbon braid in the thermal barrier structure.

Solid rocket propellant

Abstract: A solid rocket propellant includes a hydroxy-terminated caprolactone ether binder and an oxidizer. The propellant may be disposed of by contacting it with an aqueous solution of 12 N NaOH or 6 N HCl at a temperature of about 140° F. for about 24 hours to decompose the binder. Solids remaining in the solution after the binder decomposes are removed.

Determination of Triggering Condition of Vortex-Driven Acoustic Combustion Instability in Rocket Motors

Abstract: A method to determine the sufe cient condition for the occurrence of acoustic combustion instability in solid rocket motors with slotted-tube grain is proposed. The condition is obtained by comparing the frequency of the shedding vortices at the entrance of the slots and the acoustic oscillation frequency in the upstream cylindrical port. To obtain the vortex shedding frequency, a transient e ow analysis is conducted with the consideration of grain-surfaceregression. The method is assessed by analyzing e ve practical solid rocket motors employing slottedtube grain in which acousticcombustion instability occurs in threecases, whereastheothertwo showno signie cant pressureoscillations.Theresultsprovetobesuccessfulforalle vemotors.Furthermore,goodagreementisobtained between the measurements and predictions, not only for the oscillatory frequencies but also for the occurrence time.

Experimental and numerical investigations of nozzle geometry effect on the instabilities in solid propellant boosters

Abstract: Introduction The nozzle design effect on sound production is investigated at VKI to improve the understanding of the aeroacoustic coupling that occurs in the Ariane-5 booster. Acoustic measurements performed by means of PCB piezoelectric transducers are taken in a 1/30scale axisymmetric cold flow model of the Ariane5 SRM with purely axial injected flow. Flowacoustic coupling is observed for the nozzles including cavity and an analytical model of the resonance occurrence is developed. Furthermore, the maximum resonance amplitude is highly dependent on the nozzle design. Without cavity, the fluctuations are damped by at least one order of magnitude. The experimental pressure spectra are compared to numerical simulations performed using the code CPS. The frequencies are well simulated by the numerical code even if the pressure levels are overestimated. The nozzle design effect on sound production is also observed numerically. Furthermore, the nozzle cavity modifies the flow field around the nozzle head. With cavity, the recirculation bubble is shorter and the flow close to the nozzle head presents high amplitudes of radial mean velocity and fluctuation. That explains why the vortices break up when interacting with the nozzle head generating acoustic pressure fluctuations. *Ph.D. Candidate, AIAA Student Member t Professor *Head of CFD Group, AIAA Member Copyright © 2000 by the von Karman Institute for Fluid Dynamics. Published by the American Institute of Aeronautics and Astronautics, with permission. The aeroacoustics of solid propellant booster is currently being investigated at the von Karman Institute (VKI) as a part of the ASSM program (Aerodynamics of Segmented Solid Motors), initiated by the CNES to support the development of the Ariane-5 solid propellant motor (EAP). This accelerator has a segmented combustion chamber consisting of three cylindrical segments and a submerged nozzle (figure 1). Two inhibitor rings ensuring thermal protection separate the three segments. The hot burnt gas flow originates radially from the burning surface of the combustion chamber and then develops longitudinally before reaching the exhaust nozzle. During the combustion, the regression rate of the burning surface is faster than those of the inhibitor rings. Then, the lasts become obstacles into the flow-field and generate a significant risk of hydrodynamic instabilities. Pressure oscillations have been already observed for solid rocket motors, such as the U.S. Space Shuttle, the Titan SRM and the Ariane-5 EAP [1, 2, 3]. Similar results were obtained on sub-scale models of the Ariane-5 booster [4].

2D model for unsteady burning heterogeneous AP/binder solid propellants

Abstract: The burning of a periodic sandwich of the solid oxidizer, ammonium perchlorate (A) and solid fuel, HTPB is modeled by considering two-dimensional energy balances in both the solid and gas phases, two-dimensional gas species concentration, considering a reduced chemistry model for three global reactions and eight chemical species. Full heat coupling between the solid and gas phase allows the prediction of the AP, binder, and average regression rates. Flame structure including the AP decomposition flame and the diffusion flames with the binder are predicted to occur within regions ranging from 10 jjm to 200 um. Solutions are presented for various AP/binder ratios, at solid rocket pressures, ranging from 40-100 atm. Parametric studies identify the sensitivity of the burning rates to the chemical kinetics constants and the pyrolysis relations, as well as the solid-phase heat exchange coefficient, as.

Eulerian Approach for Unsteady Two-Phase Solid Rocket Flows with Aluminum Particles

Abstract: To increase the specie c impulse in a solid rocket motor, aluminum particles are embedded in the propellant formulation; combustion of these particles causes two-phase and reactive e ow features. A model of unsteady compressible e ow is presented, which includes particle-phase effects. Thecomposition of gaseous combustion products of such a solid propellant involves many species and, to reduce the computational time, a reduced description is used for the gas phase. The particles are assumed to form a continuum; then, the dispersed phase is treated by an Eulerianapproach.Abasicdescriptionofaluminum combustionand thecurrent generalknowledgearepresented. Aluminum combustion is computed by using Law’ s model, which assumes a steady vapor-phase diffusion e ame; a D n law giving the vaporized aluminum mass e ow rate is also used. The computations are performed on a simple cylindrical port motor. After some computational verie cations, a parametric study of several important quantities of the two-phase e ow is carried out. The ine uence of particle injection velocity, initial diameter of aluminum particles, and effects of oxidizing species are analyzed.

Pulsed Motor Firings

Abstract: : Combustion stability additives like zirconium carbide (ZrC), aluminum oxide (Al2O3), and zirconium ortho-silicate (ZrSiO4) have long been known to suppress combustion instability in reduced smoke, composite propellant solid rocket systems. Often, as little as 0.5% additive can stabilize an otherwise unstable rocket motor. The additives appear to have effects on both linear and nonlinear pulsed instabilities. Although several theories have been proposed, the actual mechanism on how stability additives work remains unknown. The common belief that additive particle damping alone stabilizes rocket motors is not true. Somehow, the additives change the response behavior of the propellant. Past studies have shown that the additive effect is a combination of particle damping and a reduction in the combustion response of the propellant. In the past study, four propellants were studied containing 0, 1, 3, and 5% ZrC. In this study, the 3% propellant used before will be used again, except 3% HMX will be used in one formulation and 3% ultra fine aluminum or ALEX will be used in another. The emphasis here is to examine the combustion response changes. This paper will present the results of T-burner response testing and compare the results to past propellants containing additives. The reason for the work is that recent evidence suggests that traditional additives may not work as well when solid motors are operating at higher pressures. In addition, additives like the ones proposed, add energy to the propellant which would be a performance advantage over classical additives like ZrC.

Ultrasonic Method Applied to Full-Scale Solid Rocket Motors

Abstract: The solid rocket design engineer requires a multitude of data during the static and e ight test phases of a developmentprogram. Forinstance,a betterknowledgeofthebehavioroftheinternal thermalinsulator, especially in the e ap zones, is useful for determining the margins of safety. This, in addition to the heavy work on the modelizationoftheinsulationdegradation,iswhyameasurementtoolfordeterminingthebehavioroftheinsulators hasbeendeveloped.Thistechniqueisdirect,nonintrusive,andbasedonthepropagationofultrasonicwavesthrough the materials: case, insulator, solid propellant, etc. By following the evolution of the echoes coming back from the regression surface or the degraded zone of the insulator in real time, many times per second during the test, one can deduce thefollowing data: thesequenceofeventsatthemeasurement location, the starting timeof degradation and ablation of the insulator, and, coupled with thermoablative computing code, the pyrolysis rate and the heat e ux evolutionreceivedon theinsulatorsurface.Forsolid propellant,thedisplacementofitscombustionfrontalong the insulator is followed, and of course its burning rate is deduced.

Method for manufacture of mandrels for solid fuel rocket motors and other applications

Abstract: Cores (12) for manufacture of molded products, such as a solid fuel rocket motor housing (16), are fabricated from ceramic microspheres in combination with 5-20 % polymer binder in water solution, the binder comprising poly(2-ethyl-2-oxazoline).

Numerical Simulation of Axisymmetric Solid Rocket Motor Ignition Transient with Radiation Effect

Abstract: A numerical simulation is performed to analyze the ignition transient in a solid rocket motor by employing time-dependent compressible SIMPLER algorithm fur gas phase governing equations and FVM for radiative transfer equation. Two dimensional conduction equation fur propellant grain is also coupled. The standard model with wall function is used far turbulence. A reasonably good agreement with the experiment is obtained in predicting a variation of head end pressure. Based on this comparison, a detailed development of unsteady flow and temperature fields obtained is helpful for understanding the physic phenomena involved. Especially, the radiation is found to play a significant role in igniting the solid propellant by promoting the heat feedback to propellant surface.

Modeling the combustion response function with surface and gas phase dynamics

Abstract: Considerable data exists suggesting that the response functions for many solid propellants tend to have higher values, in some ranges of frequencies, than predicted by the conventional QSHOD theory. It is a familiar idea that such behavior is associated with dynamical processes possessing characteristic times shorter than that of the thermal wave in the condensed phase. The QSHOD theory, and most of its variants, contains only the dynamics of that process, which normally has a characteristic frequency in the range of a few hundred hertz. Two previous works seeking to correct this deficiency (T’ien, 1972; Lazima and Clavin, 1992) have focused their attention on including the dynamics of the thermal wave in the gas phase. Both include effects of diffusion that complicate the analysis although the second achieves some simplification by applying the ideas of ‘activation energy asymptotics’. While their results differ in detail, both works show influences at frequencies higher than those near the broad peak of the response due to the thermal wave. The work reported in this paper has the primary purpose of constructing a simple model of the problem so that the possible dynamics of the thin region adjacent to the interface of the condensed phase may be incorporated and investigated in heuristic fashion, with and without approximations in the gas phase. It is well known from many observations, both with high-speed films and from pictures taken with scanning electron microscopes, that the surface of a burning solid propellant is certainly not smooth and in general contains both liquid and solid particles. For metallized propellants the agglomeration of aluminum drops is an important process affected, for example, by small amounts of impurities or additives. The dynamics of this region may be significant to the response of a burning propellant to external disturbances, but this phenomenon has not been previously been studied. In OUT analysis we include both phenomenological modeling of that surface layer as well as the thermal waves in both the gas and solid phase. Particular attention is given to the selection of the boundary conditions and their effect on the solution of the problem. Assumptions made in the analytical approach are tested against direct numerical integration of the relevant equations. Response functions are shown for realistic ranges of the chief parameters characterizing the dynamics of solid phase and surface layer. The results are also incorporated in the dynamical analysis of a small rocket motor to illustrate the consequences of the combustion dynamics for the stability and nonlinear behavior of unsteady motions in a motor. That is part of the primary objective of the Caltech MURl program, to understand the influences of propellant composition and chemistry on the global dynamical behavior of a solid rocket combustor by connecting the microscopic and macroscopic through the response function.

Coupled simulations of solid rocket motors

Abstract: We describe simulations of solid rocket motors that involve coupling between the core fluid flow, the structural response of the propellant and case, and the combustion of the propellant. A partitioned predictorcorrector algorithm is employed to treat the fluidstructure interaction. The combustion rate of the propellant is coupled to the fluid flow via an empirical power law relationship. Our algorithm couples the physical processes involved using a partitioned approach, enabling us to use existing codes to perform the bulk of our simulations. We give special consideration to the jump conditions that hold at the fluid-structure-combustion interface, and specialize them for the early burn phase. The interface between the eroding solid and the fluid is treated using an ALE formulation, which provides a consistent technique for handling the eroding solid. Data are presented that demonstrate the parallel performance of our code on a variety of architectures. Results from simulations of the space shuttle solid rocket motor demonstrate the applicability of our approach. Future extensions of the simulation capability to include thermal effects, turbulence and material failure will be discussed.

Combustion of aluminum in solid rocket motor flows

Abstract: The combustion of aluminum droplets was studied in a solid-rocket chamber flow-field with variable chamber pressure. Rocket motor chamber conditions were generated directly from an aluminized solid propellant combustion products flow field. The experimental test conditions were 1-20 atm pressure and ~2300 K temperature. Images of the burning aluminum droplets were made with a high-speed CCD video camera, and were de-convoluted with an Abel transformation to give radial intensity profiles of the flame and A12O3 smoke cloud surrounding each droplet. High-magnification experiments have shown major differences in the combustion of aluminum between well-studied 1 atm cold air environments and solid rocket motor conditions. The aluminum droplet flame burns with a much greater intensity at chamber pressures, indicating an increase in flame temperature, and the flame/Al2O3 cloud region thickens and extends close to the droplet surface, possibly indicating a reaction-limited mechanism in the H2O/CO2 environment. The mean value for the location of the peak intensity in the flame/Al2O3 cloud region is (r^/r,^) 2.5, and there was no apparent correlation between flame/Al2O3 smoke size and chamber pressure. However, as droplet diameter decreases, the overall A12O3 smoke cloud size increases, indicating a reaction-limited system becoming more fuel-lean as the droplet surface area decreases. Lastly, studies were conducted to observe combustion events at the surface of the aluminized propellant. Aluminum particles were first almost fully 'cleaned' before ignition began, but there was still enough binder to hold the aluminum particles onto the surface during ignition, which can lead to agglomeration.

Method for satellite injection using a solid fuel rocket motor

Abstract: A method of injecting a missile, having a solid fuel motor, into a satellite orbit. The method includes determining an along-track displacement and cross-track displacement with respect to the satellite orbit. The method also includes determining a velocity-to-be-gained for the satellite orbit, with respect to the current missile velocity. Along-track and cross-track wasting of the solid fuel motor are used to reach the satellite orbit, at the proper velocity for the satellite orbit, when the solid fuel motor burns out.

Internal Flow Analysis of Large L/D Solid Rocket Motors

Abstract: Traditionally, Solid Rocket Motor (SRM) internal ballistic performance has been analyzed and predicted with either zero-dimensional (volume filling) codes or one-dimensional ballistics codes. One dimensional simulation of SRM performance is only necessary for ignition modeling, or for motors that have large length to port diameter ratios which exhibit an axial "pressure drop" during the early burn times. This type of prediction works quite well for many types of motors, however, when motor aspect ratios get large, and port to throat ratios get closer to one, two dimensional effects can become significant. The initial propellant grain configuration for the Space Shuttle Reusable Solid Rocket Motor (RSRM) was analyzed with 2-D, steady, axi-symmetric computational fluid dynamics (CFD). The results of the CFD analysis show that the steady-state performance prediction at the initial burn geometry, in general, agrees well with 1-D transient prediction results at an early time, however, significant features of the 2-D flow are captured with the CFD results that would otherwise go unnoticed. Capturing these subtle differences gives a greater confidence to modeling accuracy, and additional insight with which to model secondary internal flow effects like erosive burning. Detailed analysis of the 2-D flowfield has led to the discovery of its hidden 1-D isentropic behavior, and provided the means for a thorough and simplified understanding of internal solid rocket motor flow. Performance parameters such as nozzle stagnation pressure, static pressure drop, characteristic velocity, thrust and specific impulse are discussed in detail and compared for different modeling and prediction methods. The predicted performance using both the 1-D codes and the CFD results are compared with measured data obtained from static tests of the RSRM. The differences and limitations of predictions using ID and 2-D flow fields are discussed and some suggestions for the design of large L/D motors and more critically, motors with port to throat ratios near one, are covered.

Development of Thermal Barriers For Solid Rocket Motor Nozzle Joints

Abstract: Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

Consistent definitions for burning-rate measurement in solid-rocket motors

Abstract: Inherent problems in burning-rate measurement in small solid-rocket motors are discussed in terms of subfactors comprising an overall absolute scale factor. A set of data-analysis definitions for burning-rate measurement that is consistent with bulk-mode theory governing pressure in solid-rocket motors is presented. The procedure using these definitions specifically treats the effects of variable exponent, noninstantaneous burnout, mean propellant shrinkage, hardware variations, and nonneutral pressure to reduce bias and scatter in measured burning rates. The present procedure does not address the effects of heat losses, formulation gradients, or nonequilibrium burning. Application of the procedure on motor data is discussed briefly.

Economical erosion-resistant rhenium coating on carbon substrates

Abstract: Erosion of material surfaces in a hot-gas valve designed to change thrust direction on a rocket motor is a limiting design consideration. Current materials used to produce solid rocket motor nozzles do not have adequate erosion resistance for the mechanical demands of hot-gas valves. This work was intended to develop rhenium as a coating on carbon-carbon substrates in lieu of solid rhenium, which has been used successfully in other hot-gas solid propellant applications. CVD processing was developed successfully depositing rhenium on pitch type carboncarbon substrates. This approach was successful in rocket erosion tests along with SiC infiltrated carboncarbon. These materials had more than 15 times greater resistance to erosion than other typical materials used on current production rocket motors.

Pressure control in a semi-closed volume upon combustion of solid propellants with an exponent in the combustion law greater than unity

Abstract: The problem of pressure control in a semi-closed volume by changing the critical cross-sectional area of a gas-release channel is considered upon solid-propellant combustion with the pressure, the combustion rate, and the free volume varied over a wide range (not smaller than one order of magnitude). For a system of automatic pressure control, a control algorithm is chosen and the conditions of partial parametric invariance with respect to the variable dynamic properties of the object to be controlled are formulated. The experimental results obtained upon improvement of the control system for solid rocket propellants whose exponent in the combustion law is greater than unity are given. The reasons for substantially nonstationary modes of operation of this system are considered, and a simplified model that approximates the phenomena of nonstationary combustion of a solid rocket propellant is proposed. The model is identified and the results of mathematical modeling are given. Recommendations on pressure control in the nonstationary modes of operation are given.