Showing papers on "Solid-fuel rocket published in 1995"

Space Propulsion Analysis and Design

Abstract: List of Authors and Editors Preface Chapter 1 Introduction to Space Propulsion 1.1 Rocket Fundamentals 1.2 The Design Process Chapter 2 Mission Analysis 2.1 Keplerian Orbits 2.2 Orbit Perturbations 2.3 Orbit Maneuvering 2.4 Launch Windows 2.5 Orbit Maintenance 2.6 Earth to Orbit Chapter 3 Thermodynamics of Fluid Flow 3.1 Mass Transfer 3.2 Thermodynamic Relations (Energy and Entropy) 3.3 Thrust Equations 3.4 Heat Addition 3.5 HEat Transfer 3.6 Design Example-Cold-Gas Thruster Chapter 4 Thermochemistry 4.1 The Chemical Heat Source: Bond Energy 4.2 Thermochemistry Basics 4.3 Products of Combustion 4.4 Flame Temperature: The Available-Heat Method 4.5 Chemical Kinetics: The Speed of the Chemical Reactions 4.6 Combustion of Liquids vs.Solids 4.7 Propellant Characteristics and Their Implications 4.8 Key Thermochemical Parameters: The Bottom Line Chapter 5 Liquid Rocket Propulsion Systems 5.1 History 5.2 Design Process 5.3 Preliminary Design Decisions 5.4 System Sizing, Design, and Trade-off 5.5 Case Study Chapter 6 Solid Rocket Motors 6.1 Background 6.2 Design Process 6.3 Preliminary Sizing 6.4 Solid Rocket Propellants 6.5 Performance Prediction 6.6 Case Study Chapter 7 Hybrid Rocket Propulsion Systems 7.1 History 7.2 Hybrid-Motor Ballistics 7.3 Design Process 7.4 Preliminary Design Decisions 7.5 Performance Estimate 7.6 Preliminary Component Design 7.7 Case Study Chapter 8 Nuclear Rocket Propulsion Systems 8.1 Introduction 8.2 Design Process 8.3 Preliminary Design Decisions 8.4 Size the Reactor 8.5 Size the Radiation Shield 8.6 Evaluate Vehicle Operation 8.7 Case Study Chapter 9 Electric Rocket Propulsion Systems 9.1 History and Status 9.2 Design Process 9.3 Specify the Mission 9.4 Select an Electric Thruster 9.5 Select Space Power 9.6 Assess System Performance 9.7 Evaluate the System 9.8 Case Study Chapter 10 Mission Design Case Study 10.1 Define Mission Requirements 10.2 Develop Criteria to Evaluate and Select a System 10.3 Develop Alternative Mission Concepts 10.4 Define the Vehicle System and Select Potential Technologies 10.5 Develop Preliminary Designs for the Propulsion System 10.6 Assess Designs and Configurations 10.7 Compare Designs and Choose the Best Option Chapter 11 Advanced Propulsion Systems 11.1 Air-Augmented Rockets 11.2 Rocket Advancements 11.3 Nonrocket Advancements 11.4 Interstellar Flight Appendix A Units and Conversions Factors Appendix B Thermochemical Data for Selected Propellants Appendix C Launch Vehicles and Staging Index

Overview of Combustion Instabilities in Liquid-Propellant Rocket Engines

Abstract: C OMBUSTION instabilities were discovered in solidand liquid-propellant rocket engines at about the same time in the late 1930s. Since then, unstable oscillations have occurred in most, if not practically all, new development programs. Indeed, because of the high density of energy release in a volume having relatively low losses, conditions normally favor excitation and sustenance of oscillations in any combustion chamber intended for a propulsion system. Figure 1 is an abbreviated chronology of some major events and features of the subject during the past 50 years. In one form or another, combustion instabilities have been under continuous study during all of that period. In time, however, the emphasis naturally has shifted, depending on what sort of full-scale systems experienced difficulties. During World War II in the United States, it seems that virtually all work in this subject was concerned with elimination of high-frequency resonant burning (the term used at the time) in small tactical solid rocket motors. The common treatment was usually a form of passive control, involving installation of baffles, resonance rods, or some other modification of geometry. Since then, the need to solve problems of instabilities in solid rockets has continued for motors of all sizes. Much of the basic understanding that has been gained is applicable to liquid rockets, despite the obvious differences in the systems. Although work on combustion instabilities in liquid rockets began in the early 1940s, significant progress was neither achieved nor required until after World War II with the development of large intercontinental ballistic missiles (ICBMs). During the 1960s, the needs of the Apollo program motivated a large amount of work on instabilities, rendered particularly important because of the astronauts

Vortex-Shedding Phenomena in Solid Rocket Motors

Abstract: In the past 20 years, periodic vortex shedding as a source of acoustic energy inside solid propellant rocket motors has been continuously studied, in connection with several motors that exhibited oscillatory behaviors although they were predicted stable by means of conventional linear stability methods. On that subject, correlations through Strouhal numbers are commonly used. In this article, the meanings of various Strouhal number definitions are discussed. The basic mechanisms for sound production in ducts or chambers are then discussed through simple tools, such as the linear hydrodynamic stability analysis, Flandro's method and the acoustic balance technique, which provide good insights into the physics of the phenomenon. Necessary conditions for vortex-shedding-driven motors are then arrived at and illustrated by actual firing results. Finally, the full numerical approaches are shown to provide unprecedented insight into the mechanisms behind the vortexshedding phenomenon and are believed to open the way to quantitative predictions of frequencies and oscillatory levels.

Triggering of longitudinal combustion instabilities in solid rocket motors: Nonlinear combustion response

Abstract: Pulsed oscillations in solid rocket motors are investigated with emphasis on nonlinear combustion response. The study employs a wave equation governing the unsteady motions in a two-phase flow, and a solution technique based on spatial- and time-averaging. A wide class of combustion response functions is studied to second-order in fluctuation amplitude to determine if, when, and how triggered instabilities arise. Conditions for triggering are derived from analysis of limit cycles, and regions of triggering are found in parametric space. Based on the behavior of model dynamical systems, introduction of linear cross-coupling and quadratic self-coupling among the acoustic modes appears to be the manner in which the nonlinear combustion response produces triggering to a stable limit cycle. Regions of initial conditions corresponding to stable pulses were found, suggesting that stability depends on initial phase angle and harmonic content, as well as the composite amplitude, of the pulse.

Flutter of Cantilevered Column Under Rocket Thrust

Abstract: This paper describes an experimental investigation on the flutter of viscoelastic cantilevers subjected to a tangential follower force. The force was produced by the direct installation of a real solid rocket motor to the tip end of the cantilevered columns. The columns lost their stability by flutter. The results were compared with theoretical flutter predictions made by accounting for an internal damping of the test columns, as well as the mass and size of the installed rocket motor. The introduction of the concept of instability in a finite time interval is of vital importance in predicting the experimental flutter force.

Solid rocket motor internal flow during ignition

Abstract: A numerical procedure is presented for the analysis of the internal flow in a solid rocket motor (SRM) during the ignition transient period of operation, along with the results obtained when this computer code was applied to several motors. The purpose of this code development effort was to achieve a detailed picture of the unsteady flowfield for a SRM of arbitrary design during this period of ignition delay, propellant ignition, flame spreading, and chamber filling/pressurization. The approach was to combine an unsteady, axisymmetric solution of the equations of inviscid fluid motion (Euler equations) with simple models for the convective and radiative heat transfer to the propellant surface during the run up to ignition. An unsteady, one-dimensional heat conduction solution for the propellant grain is coupled to this unsteady flow solution in order to calculate the propellant surface temperature. This solution, together with a surface temperature ignition criterion, determines the ignition delay and flame spreading. First, data were used from a Titan 5-1/2-segment solid rocket motor static firing to fix an unknown constant in the heat transfer model. Then, the computer code was applied to two solid rocket motors, Titan 7-segment and Space Shuttle, for which time-dependent chamber pressure measurements weremore » available from static firings. Good agreement with the data was obtained. 17 refs.« less

Effect of concentrated mass on stability of cantilevers under rocket thrust

Abstract: The paper describes the effect of an intermediate concentrated mass on the dynamic stability of cantilevered columns subjected to a rocket thrust. It is assumed that the rocket thrust is produced by the installation of a solid rocket motor at the tip end of the cantilevered columns having the intermediate mass. The rocket motor is assumed to be a rigid body having finite sizes but not a mass point. The importance of the magnitude and size of the intermediate mass is demonstrated by theory and experiment. The experimental results are compared with theoretical predictions made by taking into account the mass and size of the rocket motor as well as intermediate mass effect. The internal damping was neglected in the theoretical predictions. It is shown that theoretical stability predictions and experimental flutter limits agreed well.

Radiation-augmented combustion of homogeneous solids

Abstract: The thermal effects of incident radiation on the burning characteristics of homogeneous solids are considered and the practical implications of these effects in solid rocket motors are assessed. The applicability of the equivalence principle (equivalence between radiation and initial temperature) is investigated experimentally and analytically. Experimental steady burning rates as a function of initial temperature and mean radiant flux are presented for a fine ammonium perchlorate (A P) composite propellant which indicate that the equivalence principle is accurate to within experimental error. The equivalence principle is also assessed analytically by considering the worst case conditions of condensed phase controlled burning. For deflagration and pyrolysis of solids controlled by condensed phase reactions it is shown that a modification of Ibiricu and Williams' high activation energy asymptotic burning rate expression allows consideration of the effect of incident radiation on steady burning rat...

Modeling Solid Rocket Booster Exhaust Plumes in the Stratosphere with SURFACE CHEMKIN.

Abstract: : The results of a detailed chemical model of the transient stratospheric chemistry following passage of a large solid rocket booster motor is described. The model is based on SURFACE CHEMKIN, which is a newly developed multiphase chemical kinetic model. The model incorporates 34 chemical species and over 100 gas phase, heterogeneous, and photochemical reactions. The results show that passage of a Titan IV-sized rocket should produce an ozone 'hole' 10 km in diameter at 20 km altitude, and 28 km in diameter at 30 km altitude, lasting from a few hours to a day. The size and persistence of the hole are very sensitive to the rate of dissipation of the rocket plume, which is poorly understood at present.

An ultrasonic fuel regression rate measurement technique for mixture ratio control of a hybrid motor

Abstract: A method for real-time control of mixture ratio and chamber pressure in a hybrid motor using an ultrasonic pulse-echo technique is discussed. The technique allows rapid sequential measurement of fuel web thickness during motor operation at multiple combustion port axial locations, thereby enabling direct computation of instantaneous fuel regression rate, fuel flow rate, and, ultimately, motor operating mixture ratio. Using such data, oxidizer flows into motor combustion ports and the aft mixing chamber of a hybrid motor can be varied to achieve operation at a constant pressure and constant mixture ratio. This technique can be used to generate a constant enthalpy combustion gas environment useful for a variety of ablative material thermal response and characterization studies. In related work, the technique can provide accurate spatialltemporal regression rate histories to assist in anchoring computational fluid dynamics analytical models used for regression rate prediction. Methods of coupling ultrasonic transducers to the fuel grains of 11and 24-in. hybrid motors during motor operation, to image the regressing fuel surface with a portable ultrasonic regression rate analysis system, are discussed, and data resulting from system feasibility testing is presented. INTRODUCTION results in measuring regressing fuel or solid propellant surfaces in operating ramjets and solid rocket motors were reported using transducers operating at a frequency of 2.25 MHz to penetrate fuel or solid propellant web thicknesses of approximately 1.5 in. In the application discussed here, ultrasonic pulse-echo methods were developed to penetrate approximately 5 in. of a highdensity hybrid fuel and image the regressing fuel surface. The work was conducted as part of an initiative to develop a feedback control system for regulation of mixture ratio and chamber pressure in a hybrid motor for NASA's Large Scale Solid Rocket Combustion Simulator (LSSRCS) program (NAS~-39874).4 The objective of the LSSRCS program is to reproduce the nozzle throat aerothermodynarnic environment generated during solid rocket combustion by using a hybrid gas generator. To accomplish this task, both mixture ratio and combustion pressure must be controlled to specified levels during tests in the LSSRCS. These tasks may be best accomplished by providing real-time data on hybrid fuel consumption, which can be compared with instantaneous data for oxygen flow rate and motor pressure. A control system can then make adjustments to motor oxygen flow rate to ensure that both motor pressure and mixture ratio always remain within the tolerances desired for solid rocket combustion simulation. Combustion Simulation Use of ultrasonic techniques to measure the burning rate in solid rocket motors and fuel surface regression rate in solid fuel ramjets has been explored by ort tin^,' ~ r a i n e a u , ~ and ~ i j k s t r a . ~ Ultrasonic methods employed by these and other experimenters have been oriented toward furthering the understanding of motor internal combustion and grain burnback mechanisms. Good Figure 1 illustrates the conceptual operation of a constant-pressure, constant-mixture ratio hybrid gas generator. Given a specific fuel formulation, the general problem of constant enthalpy operation is one of varying the oxidizer flow rates to the fuel grain combustion ports (mOF) and aft mixing chamber (mOA) so as to maintain constant chamber pressure and mixture ratio as the throat

Slag accumulation in the Titan solid rocket motor upgrade

Abstract: Numerical simulations of slag accumulation in the aft end of the Titan solid rocket motor upgrade (SRMU) are described. These quasisteady, two-phase flow solutions at 0-, 30-, 55-, 80-, 110-, and 125-s burnback geometries involve a gas phase and an A12O3 liquid phase of a single droplet size, where four droplet sizes (10, 35, 60, and 100 fim) have been considered. The two-phase flow calculations are inviscid and rotational (Euler equations), with full momentum and energy coupling between the phases. Both phases are treated with an Eulerian approach. The stability of the AI2O3 droplets with respect to breakup is shown not to affect our predictions. The capture rate as a function of time is determined from the solutions at the five burn times, which is then integrated over the total burn time to determine the total slag captured. It is found that slag capture starts at time zero and continues throughout the firing. Using a droplet size distribution from a recent experimental study produces a total slag accumulation of 2265 kg, which is in good agreement with the static test results. Also, predictions of the slag pool depth as a function of time show good agreement with real-time radiography measurements. Small changes in propellant grain design, such as the use of a short-lived inhibitor in the submerged nozzle region, only provide a small reduction in the total slag captured during the burn. Slag accumulation is unaffected by the g levels typical of an SRMU flight.

Solid rocket motor grain burnback analysis using adaptive grids

Abstract: This article describes a solid rocket motor ballistics tool capable of calculating burning surface area progressions in either axisymmetric or two-dimension al grains using adaptive gridding techniques. The main advantages of this tool are its ease of use and generalized capabilities in handling spatially dependent burning rate information. The solution algorithm determines optimal locations for the number of grain surface points desired by utilizing weighting factors. Wall proximity, surface curvature, and spatial burning rate variations are considered in assigning the value of this weighting factor. Nomenclature 0, b = constants in weighting function, Eq. (4) F{ = local value of weighting function, Eq. (4) k = constant in burning rate expression, Eq. (7) n = burning rate exponent P = chamber pressure R = radial coordinate Rc = radius of curvature of propellant surface ^case = radial distance to case wall r = propellant burning rate s = natural coordinate measured parallel to local grain surface Wj = local wall proximity distance Wiimit = minimum allowable W,- value, Eq. (4) w = web distance Z = axial coordinate

Analysis of Ignition and Flame Spreading in Solid Rocket Motor Star Slots

Abstract: Introduction T HE ignition transient of a solid rocket motor (SRM) employing a pyrogen igniter can be defined as the time interval from the onset of the igniter flow to the time a quasisteady flow develops. A little-understood portion of the starting transient for (star) slotted head-end grain configurations is the time interval encompassing the initiation of the igniter flow, the first appearance of a flame on the star grain, and the subsequent flame spread over the star slot region. Previous analyses' for motors such as those used on the Space Shuttle agree quantitatively well with test data, except for the time period that directly involves burning of the headend star grain segment. Discrepancies during this time period are believed to arise from three factors: 1) the flowfield is assumed to be one-dimensional, 2) the star geometry in the head-end segment is approximated by variations in port area and grain burning perimeter, and 3) the igniter flowfield is not accounted for. The authors have previously presented a numerical calculation method utilizing the time-dependent, two-dimensional Navier-Stokes equations, which calculates the subsonic flow induced in the slot by the supersonic igniter plume. The focus of the present study is the calculation of the initial portion of the ignition transient, beginning with the start of igniter flow and ending when the head-end star slot segment of the motor is fully burning. The expanding igniter plume, the complex flow patterns within the star slot, heat transfer to the propellant grain, and subsequent propellant burning are considered.

Analysis of multidimensional and two-phase flows in solid rocket motors

Abstract: A physicomathema tical model describing the processes in the combustion chamber of solid rocket motors (SRMs) is suggested. The model can be used in practical calculations of the parameters of real motors with solid-propellant charges of complex geometry. The model allows the description of two-phase flows (from averaged zero dimensional to three dimensional) for the complete operation cycle of SRMs, including transition to steady-state operating conditions and pressure drop accompanying the propellant burnout. Equations obtained can readily be employed in designing computer codes. Methods for the solution of these equations are discussed. Results of studying the gasdynamics parameters of internal flows are reported. Calculated values are compared with experimental data obtained for a modeled gas flow turning from the head-end to the charge channel. A three-dimensional calculation of flow parameters in a slotted-tube charge of a large-size motor is illustrated by distribution patterns for pressure and velocity. The influence of the redistribution of aluminum in a solidpropellant charge on the combustion efficiency coefficient of the metal, specific impulse, and in-chamber losses is analyzed.

Solid fuel launch vehicle destruction system and method

Abstract: A system and method of shattering a launch vehicle (10) into relatively small pieces are described. The launch vehicle (10) includes at least one solid fuel rocket motor (12) having a propellant (14) disposed about a combustion chamber (16) within a rocket motor case (18). Each rocket motor (12) also includes at least one motor igniter (20) to ignite the propellant (14) and at least one explosive charge (30) adjacent the rocket motor case (18). A firing unit (44) is capable of generating a motor ignition signal and a charge explosion signal. A first propagator (58) carries the motor ignition signal to the motor igniter so the signal arrives after a propagation time T?ignition? and causes ignition of the previously unignited rocket motor. A second propagator (52) carries the charge explosion signal to the explosive charge so the signal arrives after a propagation time T?explosion? and causes an explosion against the rocket motor case. The time T?explosion? is greater than the time T?ignition? by a pressurization time T?pressurization? that is sufficient to allow the pressurization of the combustion chamber before actuation of the explosive charge. Pressurization of the combustion chamber (16) exerts forces on the rocket motor case (18) which act in combination with the subsequent force from the explosion to shatter both the case (18) and the propellant (14) much more effectively than conventional destruction systems.

History of the Development of Solid Rocket Propellant in France

Abstract: The real development of solid rocket propellents in France started in 1946 for tactical missiles applications mostly based on cast or extruded double-base propellants. The decision by General de Gaulle at the beginning of the 1960s to develop an independent Strategic Force based on ballistic missiles had a tremendous effect on the research, development, and production level of activity in the field with a strong effort on composite and high-energy propellants. The results of these efforts were applied to three generations of strategic systems, many tactical missiles, and to space boosters like DIAMANT, the first French satellite launcher (1965), or European launchers like Ariane V, which will fly in 1995. This article describes the evolution of the main propellant families and their applications during this period of 50 years.

Two-Phase Flows in the Nozzles of Solid Rocket Motors

Abstract: Physicomathematical models of multivelocity and multitemperature two-phase flows are suggested. The analyses take into account the particle rotation due to off-center collisions in solid rocket motor (SRM) nozzles and the effects of crossing the trajectories of other particles in the combustion chamber. The rotation of particles is treated by means of quasi-one-dim ensional approximation. The effect of the rotation on particle-size distribution and two-phase losses is shown. Three-dimensional two-phase flows in SRM nozzles with rectangular cross sections are considered. The influence of nozzle geometry on the gasdynamic structure of two-phase flow is determined. Axisymmetric calculations are performed for the flows in the combustion chambers and nozzles of SRM. The calculations show the importance of taking into account the intersection of particle trajectories in investigating flows in motors with charges of complex geometry. Nomenclature A = number of particle collisions a = sound velocity c = specific heat cp = specific heat of gas at constant pressure d = particle diameter d4:, = mass mean particle diameter E = total energy of unit volume

Space Shuttle solid rocket motor slag expulsion mechanisms

Abstract: A 13 psi pressure perturbation occurred at approximately 68 seconds on the right Redesigned Solid Rocket Motor (RSRM) during the STS-54 space shuttle mission. While pressure perturbations are a normal characteristic of RSRM operation, the magnitude of the STS-54 perturbation and the resulting thrust imbalance between the left and right motors was outside of flight experience. A joint Marshall Space Flight Center (MSFC) and Thiokol Corporation (RSRM manufacturer) team soon narrowed the probable cause to a temporary nozzle restriction due to slag expulsion. In support of the team, Rockwell Aerospace performed fluid finite element simulations and vehicle flight dynamic correlations to investigate possible slag expulsion mechanisms responsible for pressure perturbations. Results of the simulations and analyses provided evidence that the combination of flight induced accelerations acting on accumulated slag and nozzle vectoring were the most probable cause of RSRM slag expulsion.

Flow-structural interaction inside a solid rocket motor during ignition transient

Abstract: In this article, a numerical procedure for the analysis of the flow-structural interaction inside a solid rocket motor (SRM) is developed from existing codes. This computer code simultaneously models the developing flowfield and the associated propellant grain deformation during the ignition transient period of SRM operation. It was created by coupling together an ignition-transient flow code, which provides a detailed picture of the time-dependent flowfield and flame spreading inside the motor, with a set of structural influence coefficients, which enable us to calculate the grain deformation that results from a given surface pressure distribution. Also included in the numerical package is a grid-generation code, which generates a grid mesh for the internal-flow passages, based on the deformed shape of the propellant grain. This grid mesh then is used in the ignition transient flow code. As these three component calculations march forward in time, they interact continuously; in this way the flowfield and grain shape are updated continuously. The computer analysis was validated by and used to study the static test failure of the Titan solid rocket motor upgrade (SRMU) that occurred on April 1, 1991 at Edwards Air Force Base. It has since been used to predict the stability of other motors.

Effects of Slag Ejection on Solid Rocket Motor Performance

Abstract: In past firings of the Reusable Solid Rocket Motor (RSRM) both static test and flight motors have shown small pressure perturbations occurring primarily between 65 and 80 seconds. A joint NASA/Thiokol team investigation concluded that the cause of the pressure perturbations was the periodic ingestion and ejection of molten aluminum oxide slag from the cavity around the submerged nozzle nose which tends to trap and collect individual aluminum oxide droplets from the approach flow. The conclusions of the team were supported by numerous data and observations from special tests including high speed photographic films, real time radiography, plume calorimeters, accelerometers, strain gauges, nozzle TVC system force gauges, and motor pressure and thrust data. A simplistic slag ballistics model was formulated to relate a given pressure perturbation to a required slag quantity. Also, a cold flow model using air and water was developed to provide data on the relationship between the slag flow rate and the chamber pressure increase. Both the motor and the cold flow model exhibited low frequency oscillations in conjunction with periods of slag ejection. Motor and model frequencies were related to scaling parameters. The data indicate that there is a periodicity to the slag entrainment and ejection phenomena which is possibly related to organized oscillations from instabilities in the dividing streamline shear layer which impinges on the underneath surface of the nozzle.

High-frequency instability of combustion in solid rocket motors

Abstract: Physical and chemical grounds for the unstable combustion of solid propellants in rocket motors are considered. Micro-oscillations of electric conductivity in combustion zones are measured. Nonuniform burning of propellant components and auto-oscillating gas-phase reactions are analyzed as possible reasons for the microoscillations. The correlation between unstable combustion of solid propellants and micro-oscillations in combustion zones is established. Possible applications of the results obtained are suggested for future motor design.

Space Shuttle Solid Rocket Booster Lightweight Recovery System

Abstract: The cancellation of the Advanced Solid Rocket Booster Project and the earth-to-orbit payload requirements for the Space Station dictated that the National Aeronautics and Space Administration (NASA) look at performance enhancements from all Space Transportation System (STS) elements (Orbiter Project, Space Shuttle Main Engine Project, External Tank Project, Solid Rocket Motor Project, & Solid Rocket Booster Project). The manifest for launching of Space Station components indicated that an additional 12-13000 pound lift capability was required on 10 missions and 15-20,000 pound additional lift capability is required on two missions. Trade studies conducted by all STS elements indicate that by deleting the parachute Recovery System (and associated hardware) from the Solid Rocket Boosters (SRBS) and going to a lightweight External Tank (ET) the 20,000 pound additional lift capability can be realized for the two missions. The deletion of the parachute Recovery System means the loss of four SRBs and this option is two expensive (loss of reusable hardware) to be used on the other 10 Space Station missions. Accordingly, each STS element looked at potential methods of weight savings, increased performance, etc. As the SRB and ET projects are non-propulsive (i.e. does not have launch thrust elements) their only contribution to overall payload enhancement can be achieved by the saving of weight while maintaining adequate safety factors and margins. The enhancement factor for the SRB project is 1:10. That is for each 10 pounds saved on the two SRBS; approximately 1 additional pound of payload in the orbiter bay can be placed into orbit. The SRB project decided early that the SRB recovery system was a prime candidate for weight reduction as it was designed in the early 1970s and weight optimization had never been a primary criteria.

Pocket model application to the combustion of AP/RDX/Al/HTPB propellants

Abstract: Aluminum powders tend to agglomerate on the burning surface of the high energy low burn rate APRDWAUHTPB propellants, thus resulting in a lower combustion and n o d e expansion efficiencies, and consequently degrading the propulsion performance of the solid rocket motor. During the past two decades, tremendous experimental and analytical works had been done to deal with the problems. The current study is a continuation of our previous w r k on the similar subject, but the focus is skifted to analyze whether Cohen’s Pocket Model can be applied to the aluminized composite propellants containing RDX as the secondary oxidizer. The results show that the cinephotomicrography measured mean aluminum agglomerates size and the quenched particle collection bomb experimental data deduced fractional agglomeration can be correlated with the Pocket Model parameters after making some minor modifications on the original model. Also the specifc impulse under the conditions of different aluminum combustion efficiencies is assessed. The significance of this study is to provide an approach to improve the aluminum combustion efficiency of an existing APRDWA!JHTPB propellant, or to design a new propellant with higher aluminum combustion efficiency through carefully modulating the effective pocket fine structure. ‘d

Plume particle collection and sizing from static firing of solid rocket motors

Abstract: A unique dart system has been designed and built at the NASA Marshall Space Flight Center to collect aluminum oxide plume particles from the plumes of large scale solid rocket motors, such as the space shuttle RSRM. The capability of this system to collect clean samples from both the vertically fired MNASA (18.3% scaled version of the RSRM) motors and the horizontally fired RSRM motor has been demonstrated. The particle mass averaged diameters, d43, measured from the samples for the different motors, ranged from 8 to 11 mu m and were independent of the dart collection surface and the motor burn time. The measured results agreed well with those calculated using the industry standard Hermsen's correlation within the standard deviation of the correlation . For each of the samples analyzed from both MNASA and RSRM motors, the distribution of the cumulative mass fraction of the plume oxide particles as a function of the particle diameter was best described by a monomodal log-normal distribution with a standard deviation of 0.13 - 0.15. This distribution agreed well with the theoretical prediction by Salita using the OD3P code for the RSRM motor at the nozzle exit plane.