Showing papers on "Solid-fuel rocket published in 1998"

Improved time-dependent flowfield solution for solid rocket motors

Abstract: A model of the time-dependent velocity field in solid rocket motors is derived analytically for an oscillatory field that is subject to steady sidewall injection. The oscillatory pressure amplitude is assumed to be small by comparison to the mean pressure. The mathematical approach includes solving the momentum equation governing the rotational flow using separation of variables and multiple scales. This requires identifying scales at which unsteady inertia, convection, and diffusion are significant. A composite scale is obtained that combines three disparate scales. The time-dependent axisymmetric solution obtained incorporates the effects of unsteady inertia, viscous diffusion, and the radial and axial convection of unsteady vorticity by Culick's mean flow components (Culick, F. E. C., Rotational Axisymmetric Mean Flow and Damping of Acoustic Waves in a Solid Propellant Rocket, AIAA Journal, Vol. 4, No. 8, 1966, pp. 1462-1464). The resulting agreement with tbe numerical solution to the momentum equation is remarkable. The uncertainty in a short analytical expression is found to be smaller than the injection Mach number, which represents the error associated with the mathematical model itself. The multiple-scales solution agrees extremely well with Flandro's recent flowfield solution (Flandro, G. A., On Flow Turning, AIAA Paper 95-2730, July 1995). The present solution has the advantage of being shorter, more manageable in extracting quantities of interest, and capable of showing the significance of physical parameters on the solution.

A global modeling study of solid rocket aluminum oxide emission effects on stratospheric ozone

Abstract: Recent laboratory measurements [Molina et al. 1997] have indicated that the heterogeneous chlorine activation reaction ClONO2+HCl ⟶ HNO3+Cl2 has a reaction probability of about 0.02 on aluminum oxide particles (alumina). Since alumina is among those substances emitted by solid rocket motors (SRMs), we have assessed the heterogeneous chemical impact of SRM-emitted alumina on stratospheric ozone using the Goddard Space Flight Center two-dimensional photochemistry and transport model. Historical launch rates of the Space Shuttle, Titan III, and Titan IV rockets were used in time-dependent and steady-state model calculations. Variations in the temporal ozone decreases reflected the fluctuation in launch rate frequency. The annually averaged global total ozone (AAGTO) is computed to decrease by 0.025% by the year 1997. About one-third of this AAGTO change results from the SRM-emitted alumina while about two-thirds is due to SRM-emitted hydrogen chloride.

UV, Visible, and Infrared Spectral Emissions in Hybrid Rocket Plumes

Abstract: A survey was made of the spectral emissions from a 2 x 10 inch labscale hybrid rocket motor system The emissions in the Ultraviolet-Visible (300-750 nm), Near Infrared (750-1100 nm), and Mid Infrared (2-16 μm) regions were studied Baseline emissions were found to consist of the sodium and potassium atomic lines, present due to the use of silica phenolic insulators, and the C2, OH, and CH combustion bands Doped fuel studies were performed, using hydroxyl-terminated polybutadiene (HTPB) fuel mixed with metal salts to introduce emitters into the plume Metals studied included manganese, nickel, cobalt, copper, and iron Iron was studied in both the II and III oxidation states Manganese was also used to study the effect of concentration, and indicated that plume emission is quantitative, giving linear output for the range 5 to 40 ppm Overall, the labscale hybrid was found to offer a stable system for plume spectroscopy, whether for direct studies of the hybrid type rocket, or for use in plume simulations of other propulsion systems Introduction The hybrid rocket motor is of interest to the aerospace community for several reasons Two of these, the possible use of hybrids as boosters and their potential usefulness as a plume simulator for other rocket systems, particularly solids, indicate the need for a thorough survey of their spectroscopic emissions For potential use as boosters, which has been discussed as possible alternatives to the current Space Shuttle Solid Rocket Motors (SRM), hybrids must be evaluated as to base heating effects, which are related to total IR emissions The hybrid also appears to be an excellent motor system to use in ground based testing, especially for the general development of optical monitoring and other measurement techniques A properly designed system, built from the ground up for these types of applications, offers an attractive and safe alternative to the use of solid or liquid propellant systems This paper reports the results of a spectral study of such a hybrid system Atomic Emission Spectroscopy An atomic emission system basically consists of a source into which a sample can be introduced, a wavelength selector, and an optical detector The system used in this study is typical of those used for most emission experiments, but the source of emissions, a labscale hybrid motor, is different in several ways from those normally utilized 1 As a base for comparison, the ideal atomic emission source has the following characteristics: 1 Complete atomization of all elements 2 Controllable excitation energy 3 Sufficient excitation energy to excite all elements 4 Inert chemical environment 5 No background 6 Will accept solutions, gases, or solids 7 Tolerant to various solution conditions and solvents 8 Simultaneous multielement analysis 9 Reproducible atomization and excitation conditions 10 Accurate and precise analytical results The hybrid rocket plume does provide a high energy source for the atomization and excitation of the elements involved The plume temperature has been measured to be approximately 2500-2700 degrees C, which is about the same as a hydrogen-air flame The temperatures in the combustion chamber are even higher, on the order of 3000 degrees C, providing more energy for the atomization process The combustion stoichiometry is set to be fuel lean, which normally should provide lower background emissions Rocket plumes in general then do supply an environment which is capable of good atomization and excitation, and can provide these to solids liquids or gases presented to the combustion chamber Also, the motor control system allows precision metering of the oxidant, rigidly setting the operating point of the motor when fired This renders good reproducibility in firing conditions, which, along with fuel dopant seeding homogeneity, results in reproducibility between firings However, it is obvious that a rocket plume falls short on several of these characteristics Since the only method currently available for sample introduction is doping of the fuel grain, this limits us to metals, metal salts, or other solids, which then must not react with the polymer fuel while curing Hybrid motors display a significant amount of particulate matter in the plume, therefore, there is a component of blackbody type background radiation The plume is exposed to the atmosphere, which is not an inert chemical background, but allows additional oxidant to react with exhaust gases Characteristically, hybrid motors have an inherent tendency to pressure oscillations, on the order of 20-60 Hertz This causes fluctuations in plume intensity However, these fluctuations can normally be averaged or integrated out in the detection process, when using the "standard" atomic spectroscopy PMT or array detectors

Mechanical and burning properties of highly loaded composite propellants

Abstract: An improvement in the performance of solid rocket motors was achieved by increasing the oxidizer content of HTPB-based solid propellants. To minimize the adverse changes in the mechanical and rheological properties due to the increased amount of hard solid particles in the soft polymeric binder matrix, the optimum combi- nation of the particle sizes and volume fractions of the bimodal ammonium perchlorate and the aluminum powder in the solid load was obtained from the results of testing a series of propellant samples prepared by using ammonium perchlorate in four different average particle sizes, 9.22, 31.4, 171, and 323 mm. The maximum packing density of solids in the binder matrix was determined by changing the sizes and the volume fractions of fine and coarse ammonium perchlorate at constant solid loading. The aver- age size (10.4 mm) and concentration of aluminum powder used as metallic fuel were maintained constant for ballistic requirements. Optimum sizes and fine-to-coarse ratio of ammonium perchlorate particles were determined to be at mean diameters of 31.4 and 323 mm and fine-to-coarse ratio of 35/65. Solid content of the propellant was then increased from 75 to 85.6% by volume by using the predetermined optimum sizes and fine to coarse ratio of ammonium perchlorate. Mechanical properties of the propellant samples were measured by using an Instron tester with a crosshead speed of 50 mm/ min at 257C. The effect of oxidizer content and fine-to-coarse ratio of oxidizer on the burning rate of the propellant was also investigated by using a strand burner at various pressures. From experiments in which the size and the fine-to-coarse ratio of ammo- nium perchlorate were changed at constant solid loading, a minimum value of initial modulus was obtained for each fine-to-coarse ratio, indicating that the solids packing fraction is maximum at this ratio. The tensile strength and the burning rate increase, while the elongation at maximum stress decreases with increasing fine-to-coarse ratio of ammonium perchlorate. Experiments in which the total solid loading was increased at constant fine-to-coarse ratio of ammonium perchlorate show that the modulus, the tensile strength and the burning rate increase, while the elongation at maximum stress decreases with increasing solid loading. Propellants having solid loading of up to 82% exhibit acceptable mechanical properties and improved burning properties suitable for rocket applications. q 1998 John Wiley & Sons, Inc. J Appl Polym Sci 67: 1457-1464, 1998

Modified Fracture Mechanics Approach in Structural Analysis of Solid-Rocket Motors

Abstract: A deŽ ciency in current failure initiation criteria used in the structural design and service-life prediction of solid-rocket motors, where the critical failure mode is structural failure, is the uncertainties associated with the criteria used in predicting crack initiation and propagation. Also, the constitutive models that describe the propellants’ nonlinear behavior require an extensive experimental testing program to characterize the softening functions. A modiŽ ed fracture mechanics method that accounts for bulk inelastic behavior in the calculation of a critical strain energy release rate and a relatively simple method for nonlinear viscoelastic analysis, where a three-dimensional interpolation scheme is used to solve the Prony series equations that represent Young’s and shear relaxation moduli as functions of time, temperature, and strain level, have been developed and implemented into a Ž nite element code. Predictions for a thermal shock loading on an end-burning research and development motor were made using the fracture mechanics and nonlinear viscoelastic analysis capabilities developed in this study. Reasonable agreement between the measured stresses from instrumented rocket motors and the predicted von Mises stresses was obtained. The regions of high propensity for crack propagation corresponded to the critical regions of maximum tensile principal stress in the motor.

Pulsed motor firings - Pulse amplitude, formulation, and enhanced instrumentation

Abstract: Over the last ten years the Naval Air Warfare Center has participated in an extensive effort to understand non-linear pulsed instability in tactical sized solid rocket motors. The purpose of this work was to broaden the knowledge of design factors that influence non-linear axial mode combustion instability. This effort concentrated on reduced smoke propellant systems at pressures around 1000 psi. A new effort has been undertaken over the last several years to examine pulsed instability in reduced smoke systems operating at higher pressure. This paper is a progress report on results of this multi-year effort. In this paper, combustion instability data will be presented on five tactical size motor firings, two at 1000 and three at 1500 psi. Each motor was pulsed two times during bum. The motors were instrumented with high frequency piezoelectric quartz pressure transducers. In addition to pressure, pulse amplitude, propellant and geometry were varied. The motors were instrumented with two or three high frequency Kistler type piezoelectric pressure transducers. The motors with three transducers had them mounted at the forward, middle and aft end of the motor. Several significant achievements were made during this years program. First, the stability boundary was determined by bracketing the pulse amplitude required to trigger a motor into high levels of combustion instability. Second, placement of three transducers mounted along the length of the motor presents hard to obtain wave-form and phase data of a motor undergoing combustion instability. Third, results showed that it is possible to pulse a motor into nonlinear limiting amplitude oscillations whose propellant contains a stability additive. BACKGROUND AND INTRODUCTION The Naval Air Warfare Center (NAWCWPNS) at China Lake has participated in a program to develop an improved understanding of linear and non-linear combustion instability in solid propellant rocket motors. One primary goal of this program was to develop a systematic data base of motor and stability data for future analysis. Earlier papers have reported on previous NAWCWPNS work on this program." The motors fired in the past program were 5 inches in diameter and 67 inches in length. The majority were loaded with an 88% solids reduced smoke AP/HTPB propellant with a nominal burning rate of 0.240 in/sec at 1,000 psi. In addition, motors were fired which contained 1 percent 8 micron aluminum oxide, 90 micron aluminum oxide, and 3 micron zirconium carbide as stability additives in place of 1 percent ammonium perchlorate. Motor pressures ranged from 500 to 1500 psi. Pressure coupled combustion response measurements were made at the nominal motor operating pressures for all propellants. Several motor configurations and propellant variations were included hi the program. A total of 23 motors were fired and each motor was typically pulsed three times during burn. The baseline grain geometry was a six-point star in the aft two-thirds of the motor and a cylindrical section hi the forward end. Most of these motors were fired using the baseline reduced smoke composite propellant and three were fired with propellants containing stability additives. Three motors with star-forward grains, one motor with a full star grains, two motors with This effort was sponsored by the Air Launched Weaponry Technology Block Program Office under the authority of Tom Loftus and Scott Fuller, Naval Air Warfare Center, China Lake, CA 93555-6100. * Research Scientist, Research and Technology Division, Senior Member A1AA Approved for Public Release. Distribution is Unlimited. cylindrical cross sections, and four half length higher frequency motors were also fired. The pulsing produced 10 unstable pulses (pulses that grew to a limiting oscillatory amplitude) and 32 stable pulses (pulses that decayed). A complete description of the motors can be found in references 8, 9, 11 and 12. In addition to the work done at China Lake, non-linear stability analysis was performed by Phillips Laboratory and motor firings and analysis were performed by Canada and the United Kingdom. Australia also participated by performing propellant response testing." Some of the data generated and analysis performed during the course of this program indicated a possible increase in instability tendencies at high motor operating pressures. In addition, there have been concerted efforts to develop tactical solid rocket systems which have greatly reduced plume signatures and use light weight composite motor cases. The pressure. These motors were also pulsed. The results of this study, besides providing needed motor combustion instability data, showed clearly that increasing a motor's operating pressure increases a motors susceptibility to pulsed instability. During the course of this years work five motors were fired and pulsed. The paragraphs below will discuss the motor hardware, pulsing methods, test matrix, propellants and, most importantly, the acoustic analysis of the motor firings data. MOTOR FIRINGS DETAILS Propellant: Table 1 shows the two different formulations used during these motor firings. Unfortunately, these propellants differ in more that one way, making conclusions about formulation effects difficult to make. There is one important difference, however, that conclusions can be drawn from. Propellant A contains one percent of the Table 1. Baseline Propellant Properties Ingredient AP RDX HTPB Carbon Black ZrC Burning Rate @ 1000 psi Exponent @ 1000 psi Propellant Density Flame Temperature @ 1000 psi Speed of Sound (a), 1000 osi Motors 1-4 Approximate % 82.0 4.0 12.5 0.5 1.0 0.267 in/sec 0.360 0.0648 lbs/in 4915 °F 3554 ft/sec Motor 5 Approximate % 86.5

Advanced designs for high pressure, high performance solid propellant rocket motors

Abstract: A solid rocket propellant formulation with a burn rate slope of less than about 0.15 ips/psi (2.54 cms/69.102Pa) over a substantial portion of a pressure range of about 1,000 psi (69.102Pa) to about 7,000 psi (69.102Pa) and a temperature sensitivity of less than about 0.15 %/°F (0.15 %/0.56 °C) is provided. A high performance solid propellant rocket motor including the solid rocket propellant formulation is also provided. The rocket motor is encased in a high strength low weight motor casing which is further equipped with a nozzle throat constructed of material that has an erosion rate not more than about 2 to about 3 mils (2.54.10-3cm) per second during motor operation. The solid rocket propellant formulation can be cast in a grain pattern such that an all-boost thrust profile is achieved.

Numerical investigation of 3D two-phase turbulent flows in solid rocket motors

Abstract: A numerical investigation of the three-dimensional turbulent flow in a solid rocket motor with a submerged, vectored nozzle is presented in this paper. The simulations have been performed employing an Eulerian-Lagrangian approach: the full Navier-Stokes equations are solved numerically for the gas phase and a Lagrangian deterministic model is adopted for the discrete phase. Computations have been carried out on a realistic segmented solid rocket motor. Both an axisymmetric configuration and a 3-D geometry with a canted nozzle have been studied: a nozzle deflection of 6 degrees has been assumed. Three grid levels have been used in all computations to check grid independence. Very strong 3-D effects are predicted: an asymmetric pressure gradient is established at motor aft-end, and a vortical and circumferential flow pattern in the aft-dome is obtained. Alumina droplets trajectories have been computed for two diameters and different injection points; in most cases, three-dimensional, swirling paths have been predicted. Introduction Many solid rocket motors employ aluminized propellants to increase performance. During the combustion process liquid droplets of aluminum oxide are produced at the propellant grain surface, and are dragged into the combustion chamber by the internal flow. Some of these droplets, under certain circumstances, collect in the aft-end of the booster and give rise to alumina slag deposition which results in motor performance loss', damage of thermal protection, and possible sloshing and ejection of liquid agglomerates through the nozzle;' 3 this, in turn, is believed to cause pressure pulses and thrust imbalance. The possibilities that large lumps of alumina are exhausted through the nozzle is even greater during the thrust vector controlled phases of the vehicle ascent, in which the gimbaled nozzle operates at a certain deflection angle. This gimbaled position of the nozzle may induce strong threedimensional effects in the booster internal flow, especially in proximity of the motor aft-dome. These effects may lead to significant changes in the flow field patterns and in the alumina droplet trajectories with respect to an axisymmetric configuration, which may influence the deposition process significantly. Experimental evidence of 3-D motions in the booster aft-end region exists today. Slag movement and preferential accumulation in one sector of the motor were observed during the testing of a strategic-size solid rocket motor. Also, circumferential flows were observed by Waesche et al. in water tunnel tests when the submerged nozzle was gimbaled. Booster flow field characteristics in the motor aft-dome are very difficult to predict accurately due to the two-phase turbulent nature of the flow, mass injection from the grain surface, regions of strong recirculation and complex combustion chamber geometry, especially in three-dimensional configurations. Several numerical analyses of the internal flow field in a solid rocket motor directed to the investigation of the slag deposition phenomenon have been performed." A variety of models have been adopted: inviscid rotational or viscous flow models, including laminar and turbulent models, for the gas * Aerospace Engineer, Propulsion Section, Head; Member AIAA Aerospace Engineer, Computational Methods Section; currently with the Center for Turbulence Research, Stanford University, Stanford, CA Aerospace Engineer, Computational Methods Section, Head Copyright © 1998 by CIRA. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission American Institute of Aeronautics and Astronautics

Advanced Cryogenic Solid Hybrid Rocket Engine Developments: Concept and Test Results

Abstract: : ORBITEC has conducted considerable RD (2) solid hydrogen/gaseous oxygen; (3) solid methane/gaseous oxygen; and (4) solid methane-aluminum/gaseous oxygen. The primary focus of this paper is on solid oxygen/gaseous hydrogen. Work achieved to date includes: (1) a total of over 50 solid oxygen test firings; (2) establishment of regression rate data for the different propellant combinations, where the rates can be a factor of 20 to 40 times higher than conventional HTPB-based hybrids; (3) achievement of burn durations from 1 to 30 seconds; and (4) engine chamber pressures as high as 250 psi The potential applications include. research devices to test high-energy density matter (HEDM); hybrid rocket launch vehicle upper stages; or orbit transfer vehicles. During a current sponsored USAF Research Laboratory (RL, Edwards Air Force Base, CA) project, ORBITEC is to design, develop and test a larger, SOX/LH2 flight-type engine that will have throttling and O/F ratio control.

Implications of unsteady analytical flowfields on rocket combustion stability

Abstract: In the combustion stability assessment of internal burning solid rocket motors, a new destabilizing term must be incorporated. This term arises from the nonzero radial velocity oscillations at the propellant surface. The origin of this velocity component stems from a careful resolution of intrinsic fluid dynamics, including acoustico-vortical interactions that must satisfy mass and momentum conservation principles while accommodating the no-slip condition at the propellant surface. The source of this destabilizing term appears explicitly in two separate, independently derived, analytical formulations of the internal flowfield. Predictions generated by these analytical models are shown to coincide with reliable computational data produced recently by a numerical code that solves the nonlinearized Navier-Stokes equations. Unequivocal verifications of the analytical formulations via theoretical considerations, numerical comparisons, and global error assessments are also undertaken before examining the character and impact of the time-dependent radial velocity on rocket stability. Irrevocably, radial velocity fluctuations induce an effect whose significance is reminiscent of pressure coupling at the propellant surface. It is suggested that this effect be accounted for in current and future investigations.

Measurement system for determining solid propellant burning rate using transmission microwave interferometry

Abstract: A new and unique measurement system is presented, based on microwave transmission interferometry. The system has been developed after successful experiments with re ection microwave interferometry and operates in the Ka-band (35 GHz). It is used for the direct and continuous measurement of the instantaneous burning rate of solid rocket propellants at different pressures and different gas  ows over the burning surface in a solid rocket motor. The system consists of an experimental motor, microwave installation, hardware, and special software for data reduction. The software yields burning-rate data immediately after test runs. The principle of the measurement method and data reduction is given in this paper, together with a description of the test motor and microwave installation. Base and erosive burning rates are given for three types of ammonium perchlorate – polyvinyl chloride-based composite rocket propellant formulations. The results demonstrate that the transmission microwave interferometry system is a reliable tool for nonintrusive determination of instantaneous burning surface location of a centerperforated cylindrical propellant grain.

Feasibility Assessment of Thermal Barrier Seals for Extreme Transient Temperatures

Abstract: The assembly joints of modem solid rocket motor cases are generally sealed using conventional O-ring type seals. The 5500+ F combustion gases produced by rocket motors are kept a safe distance away from the seals by thick layers of phenolic insulation. Special compounds are used to fill insulation gaps leading up to the seals to prevent a direct flowpath to them. Design criteria require that the seals should not experience torching or charring during operation, or their sealing ability would be compromised. On limited occasions, NASA has observed charring of the primary O-rings of the Space Shuttle solid rocket nozzle assembly joints due to parasitic leakage paths opening up in the gap-fill compounds during rocket operation. NASA is investigating different approaches for preventing torching or charring of the primary O-rings. One approach is to implement a braided rope seal upstream of the primary O-ring to serve as a thermal barrier that prevents the hot gases from impinging on the O-ring seals. This paper presents flow, resiliency, and thermal resistance for several types of NASA rope seals braided out of carbon fibers. Burn tests were performed to determine the time to burn through each of the seals when exposed to the flame of an oxyacetylene torch (5500 F), representative of the 5500 F solid rocket motor combustion temperatures. Rope seals braided out of carbon fibers endured the flame for over six minutes, three times longer than solid rocket motor burn time. Room and high temperature flow tests are presented for the carbon seals for different amounts of linear compression. Room temperature compression tests were performed to assess seal resiliency and unit preloads as a function of compression. The thermal barrier seal was tested in a subscale "char" motor test in which the seal sealed an intentional defect in the gap insulation. Temperature measurements indicated that the seal blocked 2500 F combustion gases on the upstream side with very little temperature rise on the downstream side.

Experimental database describing pulse-triggered nonlinear instability in solid rocket motors

Abstract: Pulse-triggered combustion instability constitutes a considerable problem in the design and operation of solid propellant rocket motors. Results from an experimental study involving 45 full-scale motor e rings are presented to provide guidance in motor design and a database for comparison with existing and future nonlinear theories. The repeatability of pulse-triggered instability along with the effect of variations in propellant formulation, operating pressure, grain cone guration, motor scale, and motor length were evaluated according to the parameters fraction of dc shift and wave strength. The stability rating of a cone guration based on these two different parameters was not always consistent; the fraction of dc shift was judged superior. It was shown that oxidizer particle size distribution and the presence of a stability additive affected pulse-triggered instability. In addition, it was shown that instability tended to increase with pressure, that cylindrical grains tended to be less stable than star grains, that shorter motors tended to be less stable than longer ones, and that smaller motors tended to be less stable than larger ones.

Tactical Missile Guidance with Passive Seekers Under High Off-Boresight Launch Conditions

Abstract: For future short-range air-to-air missile concepts, it has been recently demonstrated that throttleable hybrid rocket motors, coupled with linear optimal guidance laws, provide signie cant performance improvements over traditional solid rocket motors utilizing similar guidance laws in high off-boresight launch conditions. A problem associatedwiththeseoptimalguidancelawsisthattheyrequiremissile-to-targetposition,velocity,andacceleration. For practical missile applications, only the initial values of relative position and velocity and an initial guess of the target’ s acceleration would be available from the launching aircraft. To further complicate the situation, most short-range missiles use a passive seeker, providing angle-only measurements, and accelerometers to measure the missile’ s accelerations. An extended Kalman e lter (EKF) is developed to estimate the required guidance information using passive seeker measurements. The performance of a missile with a passive seeker in high offboresight launch conditions is evaluated using the combination of an EKF and the guidance law. In particular, a comparison is made between three guidance schemes using estimated states from an EKF. The three guidance schemes are 1 ) an optimal guidance law coupled with a solid rocket motor, 2 ) an optimal guidance law coupled with a hybrid rocket motor, and 3 ) proportional navigation coupled with a solid rocket motor. The results show that the optimal guidance law coupled with a hybrid rocket motor provides a signie cant improvement in high off-boresight, air-to-air engagements over proportional navigation and the optimal guidance law coupled with a solid rocket motor.

The role of turbulence interactions in solid propellant combustion instability

Abstract: The nonlinear stability theories now being developed for solid propellant rockets require and depend strongly on what is traditionally termed a velocity response function. Prior work by the author has shown, through computational solutions of a turbulence and combustion model, that a significant source of response to solid rocket aeroacoustics can result from the interaction between acousticallyinduced turbulence and combustion processes. This mechanism is reconsidered in view of recent experimental results which are in qualitative agreement with the earlier predictions. These include a pronounced and abrupt increase in propellant response function at elevated pressures, a substantial increase in mean burning rate and consequently mean chamber pressure, and a threshold or triggering behavior which is dependent on mean chamber pressure. Also discussed are recent experimental data and theoretical results which indicate that the acoustic velocity response is not associated with steady-state erosive burning. Consideration is given to means of avoiding the potential substantial response from turbularization by reduction of average flame heights in composite propellant formulations.

3D turbulent Navier Stokes simulations of ignition transients in solid rocket motors

Abstract: A numerical model has been used to compute the ignition transients in solid rocket motors. The model consists of a multi-dimensionnal unsteady turbulent Navier Stokes code coupled with a ID unsteady heat conduction model in the propellant. The unsteady behavior of the propellant combustion during the ignition transient is taken into account. The numerical model has been validated on subscale test motors. Parametric studies were made to select the procedures fitting at best the experimental results. These procedures have then been retained for evaluation of the numerical model in full scale simulations. Two full scale motors ignitions related to actual complex 3D motor configurations were simulated. Calculated pressure transients are in good agreement with test results. The ignition process shown in the computed flow field is very complex; the 3D simulation proves to be a powerful mean of understanding this process. Useful information concerning the design of igniters can be deduced. Further work is needed on other motor configurations and physical models in order to achieve a complete validation. Introduction Solid rocket motor ignition is a complex process involving competing or successive physical mechanisms within the tridimensional flow field and the propellant grain surface. Design of igniters and analysis of test firings require numerical tools modelling these mechanisms. 3D turbulent Navier Stokes simulation of ignition is considered as the best way of predicting and understanding this process. The numerical code to be used in the 3D simulations must include advanced physical models to represent the energy transfers from the igniter jet to the motor propellant and the unsteady combustion processes involved at the surface of the propellant. Copyright © 1998 by SEP, Division de Snecma. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. This paper describes the 3D model used to examine the ignition transient of solid rocket motors and some of its applications. Validation computations on subscale motors are detailed. Two examples of computations of full scale motors ignition transients are given.

Solid rocket propellant behavior during static firing test using real time X ray radioscopy

Abstract: Unlike ordinary motors, solid propellants rockets operate only once and for a very short period of time. So, a method able to investigate the behavior of the grain during static firing test is of high interest for the people in charge of the development program especially regarding the following phases : - ignition - burning surface Propagation - defects behavior and motors failure modes. The connection of high speed video camera with real time X Ray equipment allows valuable data acquisition during the firing and fulfil the need of direct investigation of the related phenomenon. In this paper some examples are presented, analyzed and compared with computed simulation.

Monitoring Direct Effects of Delta, Atlas, and Titan Launches from Cape Canaveral Air Station

Abstract: Launches of Delta, Atlas, and Titan rockets from Cape Canaveral Air Station (CCAS) have potential environmental effects that could arise from direct impacts of the launch exhaust (e.g., blast, heat), deposition of exhaust products of the solid rocket motors (hydrogen chloride, aluminum oxide), or other effects such as noise. Here we: 1) review previous reports, environmental assessments, and environmental impact statements for Delta, Atlas, and Titan vehicles and pad areas to clarity the magnitude of potential impacts; 2) summarize observed effects of 15 Delta, 22 Atlas, and 8 Titan launches; and 3) develop a spatial database of the distribution of effects from individual launches and cumulative effects of launches. The review of previous studies indicated that impacts from these launches can occur from the launch exhaust heat, deposition of exhaust products from the solid rocket motors, and noise. The principal effluents from solid rocket motors are hydrogen chloride (HCl), aluminum oxide (Al2O3), water (H2O), hydrogen (H2), carbon monoxide (CO), and carbon dioxide (CO2). The exhaust plume interacts with the launch complex structure and water deluge system to generate a launch cloud. Fall out or rain out of material from this cloud can produce localized effects from acid or particulate deposition. Delta, Atlas, and Titan launch vehicles differ in the number and size of solid rocket boosters and in the amount of deluge water used. All are smaller and use less water than the Space Shuttle. Acid deposition can cause damage to plants and animals exposed to it, acidify surface water and soil, and cause long-term changes to community composition and structure from repeated exposure. The magnitude of these effects depends on the intensity and frequency of acid deposition.

Characterization of the laminar boundary layer in solid rocket motors

Abstract: This paper investigates the structure of the boundary layer in cylindrical rocket motors in light of two recent analytical solutions to the timedependent axisymmetric flowfield that have been shown to agree with numerical and experimental predictions in the forward portions of the motor where the flow remains laminar. To that end, closed form expressions that define the character of the oscillatory boundary layer are obtained in order to bring physical details into focus. The short flowfield solution published recently (Majdalani, J., and Van Moorhem, W.K.,

Energetic oxetane propellants

Abstract: An energetic solid rocket motor propellant having one or more plateau regions of low operating pressure exponent is disclosed. The propellant is formulated from ingredients including an energetic polyoxetane, an effective amount of a plasticizer, an inorganic oxidizer in at least two discrete particle size ranges, and a refractory oxide burn rate modifier.

Two-Phase Turbulent Flow Interactions in a Simulated Rocket Motor with Acoustic Waves*

Abstract: A numerical study of two-phase turbulent flows in a simulated solid rocket motor with acoustic oscillations has been carried out. The main purpose is to study the interactions between particle dynamics and gas-phase flowfield in an oscillatory environment and on this basis to further explore the particle effects on combustion instability in a solid rocket motor. The physical model consists of an axisymmetric chamber with a closed head end. As a first approach, premixed gas and particle mixtures are injected into the chamber through the side walls to simulate the evolution of combustion products from solid propellants. Aluminum particle is selected in this work due to its extensive use as combustion stabilizer in practical solid rocket motors. Periodic pressure oscillations are imposed at the exit to simulate acoustic standing waves in the chamber. The analysis is based on a two-phase flow model in which the Eulerian approach is used for the gas phase. The coupling between the two phases are treated by adding appropriate source terms in the gas-phase conservation equations. The Lagrangian approach is employed for the particle phase in order to resolve detailed particle dynamics. The numerical technique for the gas phase is based on a preconditioning technique previously developed to overcome computational difficulties for chemically reacting flows at low Mach numbers. Results in the present study show that turbulence kinetic energy follows the sinusoidal motion but with different phase compared with the mean velocity. An oscillatory flow may provide additional mechanisms to transfer kinetic energy from periodic motions to turbulence flow, and consequently may increase turbulence intensity in various parts of the chamber. On the other hand, the shear waves arising from acoustic oscillations are substantially damped because of the turbulence-enhanced viscous effects. As a result, the acoustic waves in the turbulent flow exhibits a 1-D pattern except in the near-wall region. Observations of particle trajectories reveal a strong turbulence effect on particle dispersion.

Solid rocket motor

Abstract: PROBLEM TO BE SOLVED: To provide a solid rocket motor for the propulsion power generation of flight vehicles including a rocket, a flight body or the like capable of protracting a span of combustible time without lengthening a combustion chamber. SOLUTION: This rocket motor consists of a cylindrical film 51, installing a space with the tip of a grain 2, the tip is set up in the rear, being installed concentrically in the rear in the grain 2, and dividing this grain 2 into a cylindrical inner grain 21 and an annular outer grain 22, and an end film 52 being extendedly installed from the rear end of the cylindrical film 51 to a peripheral surface of the outer grain 22, and partitioning off the rear end of the outer grain 22, and in this constitution, a buried type restrictor 5 checking the combustion progress of the grain is embedded in the grain 2. With this, a size of thrust on the whole is not varied at all, but a span of combustion time is protractible without lengthening a combustion chamber.