Showing papers on "Solid-fuel rocket published in 1984"

Influence of particulates on infrared emission from tactical rocket exhausts

Abstract: The influence of radiation scattering on the infrared radiation signature of representative plumes from four types of tactical rocket motors is investigated using the recently developed JANNAF Standardized Infrared Radiation Model (SIRRM) numerical code. The plumes are modeled as isothermal cylinders with gas and particle compositions representative of 1) advanced liquid rocket exhausts (HC1, HF/carbon); 2) low-temperatu re metal fuel solid rocket exhausts (H2O, HC1, CO/aluminum oxide); 3) reduced smoke low visibility solid rocket exhausts (H2O, HC1, CO, CO2/aluminum oxide); and 4) advanced minimum smoke solid rocket exhausts (CO, CO2, H2O/zirconium oxide). The signatures of the plumes containing carbon particles are sensitive to the amount of carbon present, but insensitive to the carbon particle size. The signatures of the plumes containing aluminum oxide particles are sensitive to both the particle size and the amount of aluminum oxide present. The emission from the advanced minimum smoke plume becomes increasingly sensitive to particle size and concentration as the particle loading increases.

Pulse-Triggered Instability in Solid Rocket Motors

Abstract: This paper presents the results of a study to assess the ability of a recently developed comprehensive nonlinear combustion instability model to predict pulse-triggered instability in solid rocket motors. Performance models were developed to calculate the mass and energy flow rates produced by three laboratory pulsers (pyro, low brisance, and piston). The mass and energy flow rates are utilized as boundary conditions for the comprehensive nonlinear combustion instability model. The model predicts the temporal and spatial evolution of the resulting waveforms (amplitude and harmonic content) in the combustion chamber. Comparisons of theoretical predictions with experimental data for both laboratory and full-scale motors are presented. Very good agreement is demonstrated between the predicted and measured pulse amplitudes, wave shapes, limiting amplitudes, mean pressure shifts, and growth rates.

Modeling Slag Deposition in the Space Shuttle Solid Rocket Motor

Abstract: The burning of an aluminized propellant within a solid rocket motor produces liquid A1/A12O3 droplets on the propellant surface. Upon leaving the surface and experiencing collisions with other droplets large agglomerates may form. A significant percentage of these droplets and/or agglomerates may be deposited on the inner surfaces of the motor as a consequence of their inability to follow the gas streamlines through the nozzle. The study described herein develops an understanding of this deposition process through flow modeling and the subsequent accumulation and pooling of this slag material within the Space Shuttle solid rocket motor (SRM). From this , an analytical procedure was developed for predicting certain unknowns relating to the deposition and pooling processes. The results of an analysis using this procedure have been qualitatively verified by post-test observations of solidified slag within static horizontally fired motors. The results of the analysis were used to provide an explanation for the slag in the QM-2 and to predict the effects of flight acceleration on slag formation. The procedure is applicable to any rocket motor configuration, although IS was devloped for the SRM.

Solid rocket motor with nozzle containing aromatic amide fibers

Abstract: A solid rocket motor is provided which comprises a rocket case and a centrally ported propellant grain comprising a main portion and a nozzle portion, wherein the main portion is a shaped and cured first propellant and wherein the nozzle portion comprises a shaped and cured second propellant composition having a lower burn rate than the first composition and having a plurality of aromatic amide fibers dispersed therethrough.

Dynamics of variable mass systems with application to the star 48 solid rocket motor

Abstract: Existing methods for the derivation of equations of motion of variable mass systems are reviewed and compared, the end product being a system of general dynamical equations for variable mass systems. These equations are used to study the lateral stability problem associated with the Star 48 solid rocket engine. It is shown that the shape of the combustion chamber could have a significant effect on the lateral stability of the rocket; specifically, a short and wide combustion chamber is destabilizing, while a long and narrow chamber is stabilizing.

Self compensating ducted rocket motor fuel nozzle and method for regulating fuel generation

Abstract: Method and device for continuously and automatically maintaining optimal control settings for rich fuel gas generation within the pre-starting, starting, and operational modes of a ducted solid fuel rocket motor, through actual or potential control of pressure within a fuel rich gas generator by use of a temperature sensitive hydraulic/mechanical choke, the mechanical component of which comprises an expansion bellows endwise attached to a spring-biased slidable nozzle valve throat blockage element, and the hydraulic component comprises a temperature-sensitive constant volume fluid reservoir, or its equivalent.

Thixotropic flame inhibitors for rocket propellant grains

Abstract: Disclosed is a thixotropic flame inhibitor for a rocket propellant grain luding an organic thixotrope in amounts from about 5 parts to about 30 parts into HTPB liquid polymer in amounts from about 70 parts to about 95 parts by stirring at elevated temperatures. After cooling the mixture to room temperature, appropriate quantities of di- or polyfunctional isocyanates and cure catalysts are added to promote cure at room temperature. Fine particulate matter may also be added to increase viscosity. The resulting formulation is a room-temperature curable, thixotropic mixture which will not flow under its own weight when applied to wettable non-horizontal surfaces in appreciable thicknesses, and which, when cured, is suitable as an inhibitor for composite solid rocket propellants. Thixotropy may be controlled by the level of organic thixotrope in the formulation while viscosity may be adjusted by the level of particulate filler; these parameters determine the allowable thicknesses from one-coat applications.

Computation of the Space Shuttle solid rocket booster nozzle start-up transient flow

Abstract: The first NASA Space Shuttle flight (STS-1) produced an overpressure wave that exceeded preflight predictions by as much as 5 to 1. This second overpressure wave occurred just after the solid rocket booster (SRB) igniter wave. To understand this overpressure phenomenon, a numerical simulation effort was undertaken. Both the SRB static firing test and STS-1 geometries were studied for two-dimensional, inviscid and viscous flow. The inviscid calculations did not produce significant second overpressure waves. However, the viscous calculations did produce second overpressure waves that qualitatively agree with experiment. These overpressure waves were present in both the static firing test and STS-1 geometries. This second overpressure wave is generated by the motion of the boundary layer separation point and the subsequent radial motion of the exhaust jet during the start-up of the SRB nozzle flow. The presence of the mobile launch platform exhaust hole wale amplifies this wave, but does not appear to be the source of any additional overpressure waves. The lack of good quantitative agreement between theory and experiment indicates that other overpressure sources, not accounted for by this simulation, may be present.

136-foot main parachute for recovery of Space Shuttle Solid Rocket Boosters

Abstract: A 136-ft-D(0) main parachute system has been developed to replace the present 115-ft-D(0) main parachutes currently flown on the Space Shuttle Solid Rocket Boosters (SRB's) for the purpose of reducing the velocity at water impact. This cluster of three larger main parachutes will decelerate the 170,000-pound SRB to a terminal nominal impact velocity of 75 feet per second (ft/s) as compared to the present velocity of 88 ft/s. The paper discusses the design, development, and manufacturing of this 136-ft parachute.

Alternate nozzle ablative materials program

Abstract: Four subscale solid rocket motor tests were conducted successfully to evaluate alternate nozzle liner, insulation, and exit cone structural overwrap components for possible application to the Space Shuttle Solid Rocket Motor (SRM) nozzle asasembly. The 10,000 lb propellant motor tests were simulated, as close as practical, the configuration and operational environment of the full scale SRM. Fifteen PAN based and three pitch based materials had no filler in the phenolic resin, four PAN based materials had carbon microballoons in the resin, and the rest of the materials had carbon powder in the resin. Three nozzle insulation materials were evaluated; an aluminum oxide silicon oxide ceramic fiber mat phenolic material with no resin filler and two E-glass fiber mat phenolic materials with no resin filler. It was concluded by MTI/WD (the fabricator and evaluator of the test nozzles) and NASA-MSFC that it was possible to design an alternate material full scale SRM nozzle assembly, which could provide an estimated 360 lb increased payload capability for Space Shuttle launches over that obtainable with the current qualified SRM design.

Effect of Variable Solid Phase Properties on Propellant Combustion

Abstract: : Mathematical models derived for describing the combustion of solid rocket propellants have in the past assumed that the thermal properties, specific heat, and thermal conductivity are constant throughout the solid phase portion of the combustion zone. The values for specific heat and thermal conductivity can vary as much as 50 to 100 percent for the temperatures in the solid phase combustion wave, values from about 250 to 1100 K. Thus, the variation in solid phase properties due to the temperature variation in the solid phase have significant effects on the burning rate, the temperature sensitivity and the pressure coupled response. The paper is concerned with predicting the effect of the variable solid phase properties on the combustion properties. Parametric calculations were made to illustrate the effect of the magnitude of the variation of solid phase properties with temperature. The results indicate that significant differences exist between the values for the combustion parameters calculated using variable thermal properties as compared to those calculated using constant thermal properties.

Propellant-Liner Bond Test Improvement Program,

Abstract: : The purpose of this program was to improve test methods for characterizing the propellant-linear-insulation bond system used in solid rocket motors. An industry-wide survey gathered information on current test specimen configurations and assessment methods. From this base an upgraded specimen was developed, then optimized, using two-dimensional finite element analyses and experimental evaluations. The optimized specimen is a rectangular flapped analog that has been improved to minimize bond stresses at the end plates. Also, it can be pulled at tens-shear angles to duplicate (analog) the tension and shear stress components in solid rocket motors. In experimental evaluation, the recommended specimen demonstrated excellent behavior when pulled in tension, shear, or combined loading.

The Suppression of Afterburning in Solid Rocket Plumes by Potassium Salts

Abstract: : The exhaust plume of a minimum-smoke solid rocket contains significant concentrations of hydrogen and carbon monoxide which when mixed with ambient air react to water and carbon dioxide producing visible flash and increased infrared radiation. Both reactions produce undesirable signatures and interference with optical guidance systems. Potassium salts have been added to propellant charges to inhibit afterburning in both guns and rockets. They have not always been effective, the inhibiting effect of the salt being related to gas composition and temperature in a complex manner which is not completely understood. Further, there is disagreement as to whether it is KOH, KO2, or K that is most important in the afterburning suppression. The results are presented here of the first year of the investigation on the efficacy of each of these on the combustion of diluted H2/CO-O2-N2 mixtures. Potassium added to the fuel-side of a H2-CO-N2-O2 flat diffusion flame at near stoichiometry is more effective in inhibiting the flame reactions than KOH added to a H2-N2-O2 flame at a stoichiometric ratio of 0.61. A description is given of burner, optical and flow metering system used in experiments. Originator supplied keywords include: Rocket plume afterburning, Combustion, and Flame spectroscopy.

An apparatus for sizing particulate matter in solid rocket motors.

Abstract: : A light scattering apparatus to measure particle size (D32) in a solid rocket motor was improved. Multiple consecutive scans of two photodiode arrays were accomplished with a pacing circuit and added memory. The device was calibrated using various suspended particle samples and found to make accurate measurements. Addtional keywords: Theses, Data acquisition, Data reduction, Diffraction, Polydispersion, Solid propellants, Computer programs. (Author)

Thermal Protection System for the Space Shuttle External Tank

Abstract: The External Tank (ET) has two major roles-to contain and deliver quality cryogenic propellants to the Space Shuttle main engines and to serve as the structural backbone for the attachment of the orbiter and solid rocket boosters. The Thermal Protection System (TPS), composed of cryoprotective foam insula tion and ablator (a sacrificial heatshield material), is applied to the outer sur faces of the ET to maintain the cryogenic propellant quality, to protect the structure from ascent heating, and to prevent ice from forming. The original foam insulation, BX-250, was replaced by CPR-421, then CPR-488 as re quirements changed. Ablator reduction accompanied the changes in foam with their attendant material-properties improvements. Ultimately, both CPR-488 and the ablator were replaced by NCFI 22-65 on the aft dome of the External Tank. The TPS design requirements, including thickness and configuration, are primarily defined by allowable structural design temperatures, propellant quality considerations, and we...

High altitude chemically reacting gas particle mixtures. Volume 2: Program manual for RAMP2

Abstract: All of the elements used in the Reacting and Multi-Phase (RAMP2) computer code are described in detail. The code can be used to model the dominant phenomena which affect the prediction of liquid and solid rocket nozzle and orbital plume flow fields.

Measurement and analysis of high-frequency pressure oscillations in solid rocket motors

Abstract: During the operation of a solid propellant rocket motor, a high-frequency , finite-amplitu de pressure oscillation may develop which is superimposed on the large, time-average chamber pressure. A number of novel techniques to isolate the high-frequency pressure signal from the near-D.C. chamber pressure are described. The hardware to do this consists of off-the-shelf instrumentation modified to give the required characteristics. The software package which is used to display and analyze these high frequency signals is also discussed here. This package will allow the editing and display of the signal in the time and frequency domains and the application of such analytical tools as the fast Fourier transform. N unstable rocket motor is characterized by a finite- amplitude pressure oscillation superimposed on the larger, near-D.C. chamber pressure. To better understand rocket motor instability and to improve on the prediction of its occurrence, high-resolution pressure data are required from the unstable firing of rocket motors. With these data and the appropriate analytical tools, the rocket motor designer can identify the source of the disturbance in the motor which causes the anomalous behavior and take the steps to eliminate it. Similarly equipped, the theoretician studying the instability can determine empirical constants (such as modal growth rates) or verify his theories. The high-frequency pressure oscillations are superimposed on the large, near-D.C. chamber pressure and so may be lost in the noise if conventional signal handling techniques are used. A novel technique has been developed1"4 to separate the high-frequency wave from the near-D.C. chamber pressure and to analyze the resulting signal. The hardware and soft- ware developed for this report were intended for use with piezoelectric type gages; however, with the appropriate modifications, any high-frequency gage system can be used. The paper is divided into four parts. The first part deals with the problems which may arise when attempting to isolate the small-amplitude wave for analysis. The second part is a comparison of the various techniques undertaken to resolve the problem. The third section describes briefly the computer system developed to analyze the resulting signals. Finally, the last part demonstrates the application of the analytical tools to the pressure signals of some problem motors. This measurement and analysis technique has been applied to centrally perforated rocket motors with relatively large length-to-diameter ratios. The propellant used was a hydroxyl-terminated polybutadiene binder with ammonium

Rocket motor nozzle

Abstract: PURPOSE:To reduce thickness of a nozzle skirt, by a method wherein, in a nozzle having a thermal insulation member between the nozzle skirt and a member for a structure installed on the outer periphery thereof, the juncture between heat insulation member and the nozzle skirt is formed by a graphite felt. CONSTITUTION:The submerge type nozzle of a solid rocket motor loaded on the uppermost stage of a rocket for space observation is formed by a reinforcing member 2, made of fiber-reinforced plastic, securely fixed to a nozzle retaining member 6 made of a metal and mounted to a motor case 10, and thermal insulation materials 3a and 45 made of fiber-reinforced plastic and securely fixed to the inner and the outer wall of the member 2. Said thermal insulation material 3a forms a staged structure in which the end part of the small size part is formed thicker than other part. The outer peripheral surface of the nozzle skirt 1, made of a ultra-heat-resistant material, is covered by sticking a graphite felt 3b thereto, and the work is formed integrally with the thermal insulation material 3a to form a thermal insulation structure.

An Acoustical Imaging System for the Inspection of Solid Rocket Motors.

Abstract: : This report describes the development and operation of an acoustical imaging, nondestructive test system for the inspection of small rocket motors. This report is intended to document the accomplishments of the program, the theoretical background, and the operational characteristics of the inspection system. The major objective of this program was to develop an acoustical imaging, nondestructive testing system for small rocket motors (i.e., 13-20 cm diameter). This inspection system was to be capable of detecting two types of defects within these rocket motors: (1) internal defects of inner-bore cracks; and (2) separations or debond between the various layers making up the wall of the rocket motor. The system was shown to be capable of detecting all of the flaw types it was originally designed to detect: (1) case/liner debonds; (2) liner/propellant debonds; and (3) inner-bore cracks. Case/liner debonds are reliably detected using the chirp frequency debond test, and the other two flaws are evident using the low frequency transmission test.

Solid fuel rocket motor

Abstract: PURPOSE:To stably fly at desired low speed by providing a partition wall between a propulsion nozzle and a solid propulsion powder, providing another nozzle at the wall and setting to form a stable combustion atmosphere for the powder in the nozzle coefficient. CONSTITUTION:When a launch signal is transmitted through a leg line 22 to an ignitor 20, the ignitor 20 is energized to fill flame in a sustainer combustion chamber 14 and to fill the flame also in a booster combustion chamber 15 through a nozzle 11 provided at a partition wall 10. Thus, the inner and end combustion surfaces of the combustion surface 2a of sustainer propulsion powder 2 and booster propulsion powder 7 are simultaneously ignited, a nozzle closure 9 is discharged by towing the remaining ignition ball 21, and combustion gas is injected from the nozzle 8. In this case, when the internal pressure of the chamber 15 decreased upon burning of the powder 7, combustion gas of the powder 2 is supplied from the nozzle 11 which is set to the prescribed nozzle coefficient to the chamber 15. Accordingly, the internal pressure in the chamber 15 can be maintained constantly.