Showing papers on "Solid-fuel rocket published in 1985"

Ignition/duct overpressure induced by Space Shuttle solid rocket motor ignition

Abstract: An analytical methodology, with a propagating solid rocket motor (SRM) exhaust front as a perturbation pressure wave generator, is developed to enhance understanding of the ignition/duct overpressure (IOP/DOP) induced during the Space Shuttle liftoff. Waveform, amplitude, and low-frequency responses of IOP/DOP are simulated. For a full-scale configuration, IOP and DOP responses are clearly isolated and distinguishable. However, in the subscale tests, an interaction region of IOP and DOP exists, which obscures interpretation of the test data. The methodology offers a first-order assessment of these phenomena with respect to the launch complex geometry and the SRM ignition transient, and provides guidance for the test data interpretation.

Solid rocket motor with dual interrupted thrust

Abstract: The invention disclosed is a solid propellant rocket motor, capable of providing two separate propulsive impulses to a missile. The rocket motor is connected at one end to the missile body, the other end including an exhaust nozzle. The rocket motor comprises two stages connected by an interstage bulkhead. The bulkhead includes a port opening which is closed by a frangible cover which prevents the second stage from igniting during burning of the first stage, but breaks up into harmless fragments during firing of the second stage.

Mechanical bond between a solid rocket propellant composition and a substrate and a method of effecting such a bond

Abstract: A mechanical bond between a solid rocket propellant composition and a substrate such as insulation material is provided by a quantity of molded member portions integral with the substrate and protruding from a substrate surface into the cured propellant composition whereby a linerless rocket motor may be provided. A method of effecting such a bond is also disclosed.

Active nutation control for spinning solid motor upper stages

Abstract: This paper discusses the application of a hydrazine-fueled, aftfacing thruster with pulses synchronized so as to neutralize coning on spin-stabilized rockets. Often used during long quiescent coast periods, the development of a more powerful system for overcoming dynamic-coupled disturbances during solid rocket burning is discussed. The contribution of vehicle modeling, disturbance characterization, and stabilizing models is shown, followed by evolution of the hardware from system requirements through flight results.

Rocket nozzles in layered construction

Abstract: An improved rocket nozzle liner for solid fuel rockets which is constructed from an inner liner of silicon carbide, a layer of highly anisotropic graphite and an outer shell of heat resistant material. The layer of anisotropic graphite permits thermal expansion of the silicon carbide, thus preventing the build up of thermal stresses and resulting fractures.

Electro-optical control of solid fuel rocket burn rate

Abstract: A solid fuel rocket burn rate control apparatus wherein grain preheating energy is supplied to the grain region just ahead of the regressing burn face by optical conductors such as fiber optic filaments that are buried in the grain and dispersed across the grain cross section. The optical conductors receive optical energy from one of several types of electrical-to-optical energy transducers such as a semiconductor laser or an incandescent source; the optical conductors promote coning burn action at the burn face of the rocket and allow control of the burn rate by electrically modified optical signals.

L*-Combustion Instability in Solid Propellant Rocket Combustion

Abstract: An analysis of the L*-combustion instability in solid propellant rockets is formulated to include (1) secondary or residual combustion in the rocket chamber and (2) the change of the mean chamber pressure. The aim was to explore if these factors might remedy the failure of the many transient heat-transfer theories with quasi-steady gaseous-phase reactions in predicting any L*-combustion instability. It became clear soon how and why the response functions derived from such quasi-steady gas-phase reaction theories must fail regardless of the aforementioned remedies.

Analysis of ballistic anomalies in solid rocket motors

Abstract: Ballistic anomalies, defined as pressure deviations from the normally expected trace, have occurred frequently during solid rocket motor firings. This paper describes techniques for determining the probable cause of such anomalies. Mass and energy balance relationships, which account for changes in chamber volume and throat area, are derived for the purpose of calculating a burning surface history that corresponds to pressure data from flight measurements or static tests. Results indicate that chamber temperature variations caused by ballistic anomalies are negligible. Characteristic burning surface signatures for propellant voids, cracks, unbonds, and high burn rate pockets are discussed. Conservation of momentum relationships are derived for the purpose of describing anomalies caused by mass ejection through the nozzle throat. Specific examples form inertial upper stage motors and the Titan T34-D solid rocket booster are presented to illustrate the application of the generic analysis techniques described in this paper.

Prediction of Detonation Transition in Porous Explosives from Rapid Compression Loadings

Abstract: : Increasing the nitramine content of solid rocket propellants increases the overall performance of the system as well as the sensitivity to Shock to Detonation Transition (SDT) and Deflagration to Detonation Transition (DDT). This report deals primarily with the analysis and numerical modeling of a combined SDT/DDT event. The results show that in some instances a zone of burning granulated propellant, confined and adjacent to a zone of cast propellant, can provide a rapid enough pressure-rise rate to shock initiate the cast material. This type of detonation hazard scenario is a real possibility in any high-energy rocket motor environment. The modeling study also indicates areas where important assumptions need to be further researched. These include: (a) relations for dynamic (transient) collapse of the voids or pores; (b) relations for setting the volume percent of hot spots based on initial porosity; (c) the evaluation and expression for the chemical rate of decomposition of the reactive, shocked material; and (d) the assessment of two-phase mixture equilibrium. The predicted run-to detonation distance as a function of porosity for HMX explosive compares favorably with limited shock initiation experiments. There is no data available to check whether the predictions of ramp-wave compressions (where rise times exceed several microseconds) presented here are valid.

Particle sizing in a solid rocket motor using the measurement of scattered light.

Abstract: : A previous study discussed the formulation of the theory for the scattering properties in the general case of particles of arbitrary size and refractive index occurring in a polydispersion of finite optical depth. Roberts and Webb concluded that three volume-surface mean diameter (Sauter) D sub 32 of the polydispersion may be determined from the intensity of diffraction scattered light from spherical particles. An experimental investigation was conducted to determine the feasibility of measuring the change in particle size across the exhaust nozzle of a small solid propellant rocket motor. Light scattering measurements were made at small forward angles at the entrance and exit of the exhaust nozzle. The experimental technique was found to be practical, especially if used in conjunction with measurements of transmitted light of multiple wave lengths. However, the determination of D sub 32 is difficult in the motor environment and is biased toward the larger particles in the size distribution. Particle size measurements were in reasonable agreement with sizes determined from collected exhaust products. Recommendations for further improvement of the apparatus are made. Keywords: Diffractively scattered light; Particle sizing.

Explorer 46 Meteoroid Bumper Experiment: Earth Orbital Debris Interpretation

Abstract: The Meteoroid Bumper Experiment on Explorer 46 (launched 1972) was placed in Earth orbit to evaluate the effectiveness of using double-wall structures against meteoroids. This paper re-examines the data from this experiment. Certain sets of sensors were found to be penetrated much more frequently than other sets. The most plausible explanation is that nearly all of the penetrations were from an Earth orbiting population of particulates. In addition, because a large percentage of the penetrations occured soon after solid rocket motors were fired in space, the particulates are most likely 75 µm diameter aluminum oxide. Aluminum oxide particulates are a major exhaust product from solid rocket motors. The size of particulates from most current solid rocket motors is found to range between 0.1 pm to 20 µm. Modeling the orbits of particulates from these rockets predicts that measurements in Earth orbit of interplanetary dust in this size range are also likely to include Earth orbiting particulates from solid rocket motors.

Analytical investigation of solid rocket nozzle failure

Abstract: On April 5, 1983, an Inertial Upper Stage (IUS) spacecraft experienced loss of control during the burn of the second of two solid rocket motors. The anomaly investigation showed the cause to be a malfunction of the solid rocket motor. This paper presents a description of the IUS system, a failure analysis summary, an account of the thermal testing and computer modeling done at Marshall Space Flight Center, a comparison of analysis results with thermal data obtained from motor static tests, and describes some of the design enhancement incorporated to prevent recurrence of the anomaly.

Optical control method for solid fuel rocket burn rate

Abstract: A solid fuel rocket burn rate control method wherein grain preheating energy is supplied to the grain region just ahead of the regressing burn face by optical conductors such as fiber optic filaments that are buried in the grain and dispersed across the grain cross section. The optical conductors receive optical energy from one of several types of electrical-to-optical energy transducers such as a semiconductor laser or an incandescent source; the optical conductors promote coning burn action at the burn face of the rocket and allow control of the burn rate by electrically modified optical signals.

Water impact testing of a filament wound case

Abstract: A lightweight Filament Wound Case (FWC) Solid Rocket Booster (SRB) is being developed by NASA to increase the payload capability of the space shuttle. As with the steel boosters, the current plan is to recover the FWC SRB's after they impact the ocean at 65 to 85 ft/sec. As the boosters enter the ocean (nozzle first) the water moves away from the vehicle creating a cavity, which then collapses on the vehicle, and results in a significant loading event. To understand this loading event, tests were conducted on a quarter scale FWC model to measure cavity collapse pressure distributions, deflected shape and the effects of end conditions and pressure scaling.

Solid rocket motor

Abstract: A solid rocket motor having a main propellant comprising glycidyl azide polymer, combustible solids and a plasticizer.

Onboard adaptive guidance design for air-to-surface missiles

Abstract: An explicit onboard guidance algorithm is developed for an air-to-surface missile The model considered is a generic solid rocket with a single thrusting stage flying over a spherical rotating earth Modern control theory methods of a state space formulation are employed to find values of the control variables that optimize the performance index while satisfying the system constraints.

Solid rocket motor

Abstract: PURPOSE:To shorten the total length of a solid rocked motor having a plug insertion system nozzle and reduce the weight of same by receiving an igniter for firing solid propusive chemicals in a plug supported by a rear retatiner. CONSTITUTION:A chamber 31 in a rocket motor is provided on the rear end with a nozzle 32. A front end plate 33 is fixedly screwed into an end of the chamber 31 to close the front end opening of the chamber 31. Between a retainer claw 34 and rear retainer 35 is sandwiched propulsive chemical grains 39 charged in the chamber 21 and having abore 38. Said grain 39 consists of inner and outer surface combustion type solid propulsive chemicals 40 and restricters 41, 42. A plug body 46 is loosely fitted in the nozzle 32 to form a nozzle throat 50 between said body 46 and the inner surface of the nozzle 32. In a cavity 51 of a plug 52 is received a igniter 55 consisting charge gunpowder 53 and a squib 54 embedded in the charge gunpowder 53.

The Solid Rocket Booster Auxiliary Power Unit: Meeting the Challenge

Abstract: The thrust vector control systems of the solid rocket boosters are turbine-powered, electrically controlled hydraulic systems which function through hydraulic actuators to gimbal the nozzles of the solid rocket boosters and provide vehicle steering for the Space Shuttle. Turbine power for the thrust vector control systems is provided through hydrazine fueled auxiliary power units which drive the hydraulic pumps. The solid rocket booster auxiliary power unit resulted from trade studies which indicated significant advantages would result if an existing engine could be found to meet the program goal of 20 missions reusability and adapted to meet the seawater environments associated with ocean landings. During its maturation, the auxiliary power unit underwent many design iterations and provided its flight worthiness through full qualification programs both as a component and as part of the thrust vector control system. More significant, the auxiliary power unit has successfully completed six Shuttle missions.

The aerodynamic challenges of SRB recovery

Abstract: Recovery and reuse of the Space Shuttle solid rocket boosters was baselined to support the primary goal to develop a low cost space transportation system. The recovery system required for the 170,000-lb boosters was for the largest and heaviest object yet to be retrieved from exoatmospheric conditions. State-of-the-art design procedures were ground-ruled and development testing minimized to produce both a reliable and cost effective system. The ability to utilize the inherent drag of the boosters during the initial phase of reentry was a key factor in minimizing the parachute loads, size and weight. A wind tunnel test program was devised to enable the accurate prediction of booster aerodynamic characteristics. Concurrently, wind tunnel, rocket sled and air drop tests were performed to develop and verify the performance of the parachute decelerator subsystem. Aerodynamic problems encountered during the overall recovery system development and the respective solutions are emphasized.

Coated oxidizers for combustion stability in solid-propellant rockets

Abstract: Experiments are conducted in a laboratory-scale (6.25-cm diameter) end-burning rocket motor with state-of-the-art, ammonium perchlorate hydroxy-terminated polybutadiene (HTPB), nonmetallized propellants. The concept of tailoring the stability characteristics with a small amount (less than 1 percent by weight) of COATING on the oxidizer is explored. The thermal degradation characteristics of the coat chemical are deduced through theoretical arguments on thermal diffusivity of the composite material (propellant). Several candidate coats are selected and propellants are cast. These propellants (with coated oxidizers) are fired in a laboratory-scale end-burning rocket motor, and real-time pressure histories are recorded. The control propellant (with no coating) is also tested for comparison. The uniformity of the coating, confirmed by SEM pictures and BET adsorption measurements, is thought to be an advance in technology. The frequency of bulk mode instability (BMI), the pressure fluctuation amplitudes, and stability boundaries are correlated with parameters related to the characteristic length (L-asterisk) of the rocket motor. The coated oxidizer propellants, in general, display greater combustion stability than the control (state-of-the-art). The correlations of the various parameters are thought to be new to a field filled with much uncertainty.