Showing papers on "Solid-fuel rocket published in 1980"

One- and Two-Phase Nozzle Flows

Abstract: A time-dependent technique, in conjunction with the boundary-fitted coordinates system, is applied to solve a gas-only one-phase flow and a fully-coupled, gas-particle two-phase flow inside nozzles with small throat radii of curvature, steep wall gradients, and submerged configurations. The emphasis of the study has been placed on one- and two-phase flow in the transonic region. Various particle sizes and particle mass fractions have been investigated in the two-phase flow. The salient features associated with the two-phase nozzle flow compared with those of the one-phase flow are illustrated through the calculations of the JPL nozzle, the Titan III solid rocket motor, and the submerged nozzle configuration found in the Inertial Upper Stage (IUS) solid rocket motor.

Thrust and ignition transients of the Space Shuttle solid rocket motor

Abstract: Techniques developed for predicting and analyzing ignition, pressurization, and thrust transients in large segmented motors were applied to the four developmental Space Shuttle solid rocket motors. Following the first test, attention was focused on understanding dynamic thrust and the rate of thrust increase during pressurization and predicting effects of reduced igniter mass flow rate and gas temperature. The close coupling between the igniter and the head-end segment and the high length-to-diameter ratio make understanding of the longitudinal pressure waves essential. The igniter performance along with igniter to head-end segment geometry govern the induction interval and head-end segment ignition. Following ignition of the head-end segment, subsequent flame spreading and pressurization tend to be motor properties which are not appreciably altered by changes in the head-end igniter mass flow rate.

Propulsion/control modular booster

Abstract: A modular, apogee-control package is disclosed which can be added to exisg missiles, and which will limit the apogee of the missile trajectory by implementation of thrust vector control (TVC). The package comprises a boost guidance unit, a solid rocket propellant motor, and jet vane TVC.

Laboratory investigations of cloud nuclei from combustion of Space Shuttle propellant

Abstract: Small quantities of solid rocket motor propellant, of the type to launch the Space Shuttle, were burned at ambient pressure in the laboratory to provide aerosol samples for characterization. A portion of each sample was injected into an isothermal cloud chamber and the remainder into a 770-liter holding tank. Portable ice nucleus (IN) counters, filter devices for IN determinations and a cloud condensation nucleus (CCN) counter sampled from the tank. The measurements show that particles resulting from the combustion of the propellant are active IN (3.3 times 10 to the 8th to 1.5 times 10 to the 11th/g active at 20 C). The portable counters and filters detected significantly fewer IN than the isothermal cloud chamber. The propellant aerosol is a prolific source of CCN that swamped the instrument.

Space Transportation System Solid Rocket Booster Thrust Vector Control System

Abstract: The Solid Rocket Booster, Thrust Vector Control (TVC) system was designed in accordance with the following requirements: self-contained power supply, failsafe operation, 20 flight uses after exposure to seawater landings, optimized cost, and component interchangeability. Trade studies were performed which led to the selection of a recirculating hydraulic system powered by Auxiliary Power Units (APU) which drive the hydraulic actuators and gimbal the solid rocket motor nozzle. Other approaches for the system design were studied in arriving at the recirculating hydraulic system powered by an APU. These systems must withstand the imposed environment and be usable for a minimum of 20 Space Transportation System flights with a minimum of refurbishment. The TVC system completed the required qualification and verification tests and is certified for the intended application. Substantiation data include analytical and test data.

Analysis of the measured effects of the principal exhaust effluents from solid rocket motors

Abstract: The feasibility of conducting environmental chamber tests using a small rocket motor to study the physical processes which occur when the exhaust products from solid motors mix with the ambient atmosphere was investigated. Of particular interest was the interaction between hydrogen chloride, aluminum oxide, and water vapor. Several types of instruments for measuring HCl concentrations were evaluated. Under some conditions it was noted that acid aerosols were formed in the ground cloud. These droplets condensed on Al2O3 nuclei and were associated with the rocket exhaust cooling during the period of plume rise to stabilization. Outdoor firings of the solid rocket motors of a 6.4 percent scaled model of the space shuttle were monitored to study the interaction of the exhaust effluents with vegetation downwind of the test site. Data concerning aluminum oxide particles produced by solid rocket motors were evaluated.

Rockets for spin recovery

Abstract: The potential effectiveness of rockets as an auxiliary means for an aircraft to effect recovery from spins was investigated. The advances in rocket technology produced by the space effort suggested that currently available systems might obviate many of the problems encountered in earlier rocket systems. A modern fighter configuration known to exhibit a flat spin mode was selected. An analytical study was made of the thrust requirements for a rocket spin recovery system for the subject configuration. These results were then applied to a preliminary systems study of rocket components appropriate to the problem. Subsequent spin tunnel tests were run to evaluate the analytical results.

Determination of solid propellant admittances by the impedance tube method

Abstract: The adaptation of the impedance tube technique for the measurement of solid propellant admittances and response functions is described. These quantities are needed for combustion stability analysis of solid rocket motors. The experimental set up consists of a tube with a disk of solid propellant sample placed at one end and a combination of an exhaust valve and an acoustic driver placed at the other end. Performing a test consists of turning on the acoustic driver to excite a standing wave of a predetermined frequency in the tube, ignition, and burnout of the solid propellant sample. During a test, acoustic pressure data is measured by fifteen pressure transducers distributed along a distance that includes, at least, two standing wave pressure minima. The data are transferred via a fast analog-to-digital converter to a minicomputer system for immediate storage and later analysis. The measured acoustic pressure amplitude and phase data are input into a newly developed data reduction procedure that is based upon the solution of the impedance tube wave equations, for the determination of the admittance and response function of the tested solid propellant sample and the acoustic energy losses in the gas phase. The paper presents results obtained in a series of tests conducted with an A-14 propellant. The paper demonstrates the capability of the developed experimental technique to determine simultaneously the acoustic characteristics of the flow inside the impedance tube and the admittance of the burning solid propellant for the duration of the experiment. Nomenclature A /y = defined in the Appendix C = constant defined in Eq. (8) c = velocity of sound, m/s cp - specific heat at constant pressure, kcal/kg K cv = specific heat at constant volume, kcal/kg K E = error function defined in Eq. (15),, (N/m 2)2 or Pa 2 F = losses due to viscosity and gas phase damping, kg/m2s2 G = gas phase bulk loss coefficient, kg/m3 s / = identity matrix M = Mach number m = mass flow rate per unit cross-sectional area, kg/m2 s p = pressure, N/m2 or Pa Q = volumetric heat source, kcal/m3 s mole R = gas constant, kcal/kg mole K r = burning rate, m/s s = entropy, kcal/kg K T = temperature, K; also transmission matrix defined in Eq. (10) TWQ = wall temperature at x = 0, K u = axial velocity, m/s Y = specific admittance, dimensionless Z = variable used to represent oscillatory quantities, defined by Eq. (6) ps = density of solid propellant, kg/m3 p = density, kg/m3 Superscripts (~) = variable describing a steady-state quantity ()' = variable describing a perturbation quantity

U. S. Air Force approach to plume contamination

Abstract: Exhaust products from rocket engine firings can produce undesirable effects on sensitivesatellite surfaces, such as optical systems, solar cells, and thermal control surfaces. TheAir Force has an objective of minimizing the effect of rocket plume contamination on space-craft mission effectiveness. Plume contamination can result from solid rocket motors, liq-uid propellant engines, and electric thrusters. To solve the plume contamination problem,the Air Force Rocket Propulsion Laboratory (AFRPL) has developed a plume contamination com-puter model which predicts the production, transport, and deposition of rocket exhaust pro- ducts. In addition, an experimental data base is being obtained through ground -based vacu-um chamber experiments and in- flight measurements with which to compare the analytical re- sults. Finally, the experimental data is being used to verify and improve the analyticalmodel. The plume contamination model, known as CONTAM, has been used to make contaminationpredictions for various engines. The experimental programs have yielded quantitative data,such as species concentrations and temperatures, in all regions of the plume. The resultof the modelling and experimental programs will ultimately be computer models which can beused by the satellite designer to analyze and to minimize the effect plume contaminationwill have on a particular spacecraft system.IntroductionSpacecraft contamination can not be attributed to one or even a few sources, but rathermust be attributed to all sources of migratory material which are foreign to the specificsystem of interest. Examples of potential contamination sources include outgassing of sol-id materials, the exhaust effluents of onboard propulsion systems, lubricants used formechanical devices, etc. The spacecraft designer must therefore be careful in choosing anymaterial for use on the spacecraft, basing his decision on the material properties and be-havior under vacuum conditions. In addition, proper care prior to launch must be consider-

The Dynamics of Solid Propellant Combustion.

Abstract: : A nonlinear stability analysis of solid rocket propellant burning was carried out, within the framework of a thermal theory for thin (quasi-steady gas phase) heterogeneous flames. This required an integral method in reducing the partial differential equation for the condensed phase heat conduction to an approximate ordinary differential equation. A nonlinear algebraic function, called static restoring function, was found that contains all basic properties of equilibrium and stability of burning solid propellants. This function depends on the nature of the solid propellant (including its flame) and the operating conditions. Analysis of the static restoring function reveals the existence of lower and upper burning stabilities. Nonlinear static and dynamic burning stability boundaries were determined and stability properties of several unsteady flame models were compared. Domains of stationary reacting solutions, damped oscillatory burning, self-sustained oscillatory burning, non sustained burning were singled out. Pressure deflagration limit was evaluated. Validity and plausible extensions of this theory are discussed. Most of the analytical predictions were verified by computer simulated burning tests. Experimental results obtained in shock tube apparatus, depressurization strand burner, laser doppler velocimetry rig are presented. (Author)

Low-thrust chemical rocket engine study

Abstract: Parametric data and preliminary designs on liquid rocket engines for low thrust cargo orbit-transfer-vehicles are described and those items where technology is required to enhance the designs are identified. The results of film cooling studies to establish the upper chamber pressure limit are given. The study showed that regen cooling with RP-1 was not feasible over the entire thrust and chamber pressure ranges. The thermal data showed that the RP-1 bulk temperature exceeded the study coking temperature limit of 1010 R. Based upon the results presented, O2/H2 and O2/CH4 regen engine systems and O2/H2 film cooled engines were selected for further study in the system analysis. Six engine design concepts are examined.

One- and two-phase nozzle flows

Abstract: A time-dependent technique, in conjunction with the boundary-fitted coordinates system, is applied to solve a gas-only one-phase flow and a fully-coupled, gas-particle two-phase flow inside nozzles with small throat radii of curvature, steep wall gradients, and submerged configurations. The emphasis of the study has been placed on one- and two-phase flow in the transonic region. Various particle sizes and particle mass fractions have been investigated in the two-phase flow. The salient features associated with the two-phase nozzle flow compared with those of the one-phase flow are illustrated through the calculations of the JPL nozzle, the Titan III solid rocket motor, and the submerged nozzle configuration found in the Inertial Upper Stage (IUS) solid rocket motor.

Prediction of finite pressure oscillations in stable rocket motors

Abstract: A methodology is outlined for predicting the amplitude of forced vibrations in the acoustic cavity of a solid rocket motor. The equation for forced vibration of the motor cavity acoustic system is written by parallel with the acoustic mechanical analogy. Acoustic and aeroacoustic theory are used to predict the frequency and intensity of vortex systems or turbulence created by passage of the mean flow over geometric discontinuities in the motor port. Approximate methods are presented for coupling the acoustic field due to the flow noise with the chamber acoustics and for summing the effect upon the multiple acoustic modes.

MHD measurement of acoustic velocities in rocket motor chambers

Abstract: A linear analysis of longitudinal acoustic waves (either forced or self-induced) at or near the fundamental frequency in solid rocket chambers revealed that the acoustic velocity phase angle relative to the head-end pressure at the chamber midpoint is strongly dependent on the real part of the overall pressure-coupled response function and the imaginary part of the overall velocity-coupled response function. A flowmeter based on Faraday's Law was devised to demonstrate the feasibility of such velocity measurements under rocket motor conditions. The interaction of an externally-excited magnetic field with the unsteady velocity of hightemperature combustion gases results in a corresponding unsteady electrical potential whose magnitude is proportional to the flow velocity. The system is readily calibrated since electrical output does not depend on knowledge of ionization level and electrical conductivity (within certain limits) and measurements are independent of propellant type. Motor firing data has been used to deduce the real part of the pressure-coupled response from the phase angles between head-end pressures and midpoint velocities and the imaginary part of the velocity-coupled response from the phase angles between headand nozzle-end pressures. The overall response functions measured in the device are compatible to the form used in stability prediction models.

Design and development of the quad redundant servoactuator for the space shuttle solid rocket booster thrust vector control

Abstract: The design and theory of operation of the servoactuator used for thrust vector control of the space shuttle solid rocket booster is described accompanied by highlights from the development and qualification test programs. Specific details are presented concerning major anomalies that occurred during the test programs and the corrective courses of action pursued.

Convective heating of solid propellant during ignition

Abstract: A pattern search optimization procedure is used to deduce from test data on the solid rocket boosters of the Space Shuttle the convective heat transfer coefficients between the igniter and main motor combustion gases and the surface of the solid propellant by utilizing a spatial and temporal ignition transient analysis. The 'best' convective coefficients are found which will minimize the overall deviation of the analytical prediction from the ignition transient test data. Computationally, this is achieved by treating the motor in such a way that only four optimization variables must be determined. The resulting coefficients greatly improve the prediction capabilities of the ignition transient analysis for the Space Shuttle SRM.

Ice nuclei measurements from solid rocket motor effluents

Abstract: The ice crystal forming nuclei (IN) measured in solid rocket motor (SRM) exhaust products is discussed in relation to space shuttle exhaust. Preliminary results from laboratory investigations and flight preparations for March 1978 Titan launch are discussed. The work necessary to provide adequate measurements of IN and cloud condensation nuclei (CCN) in the stabilized ground clouds from SRM's is studied.

Device for piloting the propulsion system of a solid propellant rocket

Abstract: This invention concerns the piloting of propulsion systems for solid rocket propellant missiles. It consists of a piloting assembly in which the front section of a thrust nozzle integrated in the body of a propulsion system is pierced with at least three orifices which can be covered by means of valves and which are designed to provide direct injection into the thrust nozzle of hot gases from the powder block of the propulsion system to pilot the missile. An auxiliary block of "cold" powder is fitted around the nozzle on the input side of the orifices. The application is for piloting of ballistic missiles.

Effect of silicone oil on solid propellant combustion in small motors. [for rockets

Abstract: The feasibility of reducing troublesome nozzle blockage (by condensation deposits) in laboratory-scale solid rockets by addition of a silicone oil as a propellant ingredient was explored experimentally. An aluminized composite propellant and its counterpart with 1% silicone oil replacing part of the binder were fired in a 63.5 mm diameter, end-burning, all-metal burner. Pressure-time histories were recorded for all of the tests by a Taber gauge mounted at the downstream end of the chamber; temperature-time data at the nozzle throat were obtained in some of the runs by thermocouples having junctions positioned at the wall but insulated from the metal. Deposition of condensables on the nozzle walls causing a progressive increase in the chamber pressure with time was noted. The fraction of firings exhibiting practically no condensation was 59% with silicone and 32% without. On the average, temperature readings at the nozzle throat were higher with the silicone propellants. Although various phenomena may contribute to these findings, the results are not understood completely.

The 17th JANNAF Combustion Meeting, Volume 1

Abstract: The combustion of solid rocket propellants and combustion in ramjets is addressed. Subjects discussed include metal burning, steady-state combustion of composite propellants, velocity coupling and nonlinear instability, vortex shedding and flow effects on combustion instability, combustion instability in solid rocket motors, combustion diagnostics, subsonic and supersonic ramjet combustion, characterization of ramburner flowfields, and injection and combustion of ramjet fuels.

Proceedings of Shuttle Environmental Effects Program Review

Abstract: Measurements of Titan exhaust cloud effluents are documented and compared, mesoscale and microphysical acid rain models are described, and a submesoscale model is proposed. Various instruments and facilities for measuring ice nuclei and other constituents of solid rocket motor exhaust effluents are discussed. Regional air quality monitoring and rain collection systems are described, and the ecological impact of solid rocket motor exhaust effluents is examined. The potential effect of space shuttle launches is estimated where data are adequate.