Showing papers on "Solid-fuel rocket published in 1978"

Process for the leaching of AP from propellant

Abstract: Enhanced recovery of ammonium perchlorate from waste solid rocket propellant is effected by leaching shredded particles of the propellant with an aqueous leach solution containing a low concentration of surface active agent while stirring the suspension.

A Comprehensive Theory of Erosive Burning in Solid Rocket Propeilants

Abstract: The theory of erosive burning has been constructed front first principles using turbulent boundary layer concepts. It is shown that the problem constitutes one of solution of flame propagation equation for turbulent flow. The final approximate solution for the case of single step overall kinetics reveals the combined effects of fluid mechanics and chemical kinetics. The results obtained from this theory are compared with earlier experimental results. The dependence of erosive burning characteristics on various parameters has been elucidated.

Extensions to analysis of ignition transients of segmented rocket motors

Abstract: The analytical procedures described in NASA CR-150162 were extended for the purpose of analyzing the data from the first static test of the Solid Rocket Booster for the Space Shuttle. The component of thrust associated with the rapid changes in the internal flow field was calculated. This dynamic thrust component was shown to be prominent during flame spreading. An approach was implemented to account for the close coupling between the igniter and head end segment of the booster. The tips of the star points were ignited first, followed by radial and longitudinal flame spreading.

Optimization of solid rocket motor igniter performance requirements

Abstract: A numerical technique is presented for optimizing a set of solid rocket motor (SRM) ignition control variables to achieve a specific requirement or set of requirements to be used in preliminary design of an SRM igniter. The mathematical model of the igniter transients uses a simplified ignition simulation routine to calculate igniter and SRM performance. The object function to be optimized (minimized) is typically the summation of The absolute values of the differences between the desired and computed thrust values of the SRM at specified times during ignition. Constraints are imposed in the form of limiting values of igniter characteristics, flame-spreading speeds and/or maximum rate of pressure rise in the SRM. Optimization is obtained using a direct pattern search technique to determine the required values of the controlling variables. Examples are presented which illustrate the ability of the technique to meet practical design requirements. Computational results are shown to be consistent with the sta...

Silicone containing solid propellant

Abstract: Addition of a small amount, for example 1% by weight, of a liquid silicone oil to a metal containing solid rocket propellant provides a significant reduction in heat transfer to the inert nozzle walls eliminating metal oxide slag collection and blockage of the nozzle and increases burning rate by about 5 to 10% improving ballistic performance.

Gas-Deployed Skirt for Rocket Motor Nozzles

Abstract: The gas-deployed skirt (CDS) is a metal nozzle extension that is folded radially inward and packaged entirely within the cavity of a rocket motor nozzle. The skirt is deployed by low-pressure motor ignition gases and then stabilized in the open position by the internal pressure of fully developed nozzle exhaust gas flow. The conception, design, and analysis are discussed in the context of seven successful CDS tests on 100-lb liquid rocket rocket motors and 5000-lb solid rocket motors. Typical applications and the resulting performance improvements are also discussed.

Asymmetric Unsteady Flow in Forward-Facing Cavities

Abstract: An analysis has been performed of the flow oscillation observed in the "empty" nozzles of the space shuttle solid rocket motors (SRM's) after they have been jettisoned at burnout. The unsteady flow phenomenon is similar in many respects to the flow oscillations observed on bodies with flow separation spikes. Hence, a previously developed analytic method for prediction of the spiked-body oscillation frequency has been extended to apply to forward-facing cavities. The predicted frequencies are in good agreement with experimental results. Furthermore, a simple time history analysis produces time-average loads on the forward-facing SRM nozzles that agree well with the observed discontinuous load changes caused by the flow oscillation. The analytic methods have direct or indirect application to analysis of inlet dynamics, supersonic control buzz, and nosetip aerodynamics of ablating re-entry vehicles.

Effects of solid-propellant temperature gradients on the internal ballistics of the Space Shuttle

Abstract: The internal ballistic effects of combined radial and circumferential grain temperature gradients are evaluated theoretically for the Space Shuttle solid rocket motors (SRMs). A simplified approach is devised for representing with closed-form mathematical expressions the temperature distribution resulting from the anticipated thermal history prior to launch. The internal ballistic effects of the gradients are established by use of a mathematical model which permits the propellant burning rate to vary circumferentially. Comparative results are presented for uniform and axisymmetric temperature distributions and the anticipated gradients based on an earlier two-dimensional analysis of the center SRM segment. The thrust imbalance potential of the booster stage is also assessed based on the difference in the thermal loading of the individual SRMs of the motor pair which may be encountered in both summer and winter environments at the launch site. Results indicate that grain temperature gradients could cause the thrust imbalance to be approximately 10% higher in the Space Shuttle than the imbalance caused by SRM manufacturing and propellant physical property variability alone.

Methods used for Space Shuttle SRB thrust shape design

Abstract: Optimization of the Space Shuttle trajectory is discussed with reference to the low acceleration profile required for Shuttle missions. Static tests of the nominal flight curve are described in terms of impulse requirements, vacuum thrust, and burn time. Attention is given to BARF (Burning Anomaly Rate Factor), and it is noted that mandrel fabrication is intended to include the flexibility to counter BARF, should it occur. Test results are presented in which both BARF and specific impulse are considered as independent variables. It was found that no erosive burning occurred, BARF did not occur, specific impulse was on the order of 265 sec, and flow anomalies in the star region produced head-to-aft stagnation pressure drops in excess of theoretical predictions. In other areas, good agreement is noted between theoretical prediction and empirical data.

Space Shuttle SRM metal case component fabrication

Abstract: The 146 in. diam metal case components of the Space Shuttle Solid Rocket Motor (SRM) being developed have been successfully static-tested. The limitations placed on the program included current practice and facilities at the steel mills, forging suppliers, heat treaters, and machining operations. In addition, Thiokol had not previously fabricated metal components of this size with a minimum fracture toughness of 90 ksi-in. to 1/2 power. To insure that the SRM was producible within the established guidelines, it was necessary to coordinate all data heat by heat, forging by forging, and heat treat run by heat treat run. The basic fabrication sequences are outlined, and the data from the heat treat programs are presented.

Nitric oxide production in the stratosphere within the Space Shuttle's solid rocket motor plumes

Abstract: This study focuses on establishing the sensitivity of predictions of NO production to uncertainties in altitude, reaction rate coefficients, turbulent mixing rates, and Mach disk size and location. The results show that relatively large variations in parameters related to these phenomena had surprisingly little effect on predicted NO production.

Characterization of the non axial thrust generated by large solid propellant rocket motors in three axis stabilized ascent

Abstract: Nonaxial thrusts produced by solid rocket motors during three-axis stabilized attitude control have been determined from ascent experience on twenty three Burner II, Burner IIA and Block 5D-1 upper stage vehicles. A data base representing four different rocket motor designs (three spherical and one extended spherical) totaling twenty five three-axis stabilized firings is generated. Solid rocket motor time-varying resultant and lateral side force vector magnitudes, directions and total impulses, and roll torque couple magnitudes, directions, and total impulses are tabulated in the appendix. Population means and three sigma deviations are plotted. Existing applicable ground test side force and roll torque magnitudes and total impulses are evaluated and compared to the above experience data base. Within the spherical motor population, the selected AEDC ground test data consistently underestimated experienced motor side forces, roll torques and total impulses. Within the extended spherical motor population, the selected AEDC test data predicted experienced motor side forces, roll torques, and total impulses, with surprising accuracy considering the very small size of the test and experience populations.

Effects of entrained water and strong turbulence on afterburning within solid rocket motor plumes

Abstract: During the first few seconds of the space shuttle trajectory, the solid rocket boosters will be in the proximity of the launch pad. Because of the launch pad structures and the surface of the earth, the turbulent mixing experienced by the exhaust gases will be greatly increased over that for the free flight situation. In addition, a system will be present, designed to protect the lifting vehicle from launch structure vibrations, which will inject quantities of liquid water into the hot plume. The effects of these two phenomena on the temperatures, chemical composition, and flow field present in the afterburning solid rocket motor exhaust plumes of the space shuttle were studied. Results are included from both a computational model of the afterburning and supporting measurements from Titan 3 exhaust plumes taken at Kennedy Space Center with infrared scanned radiometers.

Solid rocket engine in two-stage thrust mode

Abstract: PURPOSE: To enable the thrust force of an engine to be varied according to the necessity by allowing a rocket engine which stores solid propellant to be equipped with igniters which operate at a same time or at arbitrary separate times. CONSTITUTION: A rocket engine stores plateau type double base propellant in a rocket chamber 1, and in two drilled holes on an end plate 4 which is applied with asbestos heatproof processing at the rear part of a combustion chamber 3, heatproof reinforced plastic plates 5 are embedded, and on said heatproof plates 5, two igniters 6 are fixed. Said igniter 6 can be ignited at a same time or at separate times by the conduction of ignit ion conductors 7, and thrust force can be varied. Combustion gas gushes from a nozzle 8, blowing off a nozzle closure 9. COPYRIGHT: (C)1979,JPO&Japio

Solid fuel rocket motor - has combustion chamber split into different chambers by perforated transverse walls

Abstract: The solid fuel rocket motor has a sieve-type body in the combustion chamber in front of the jet opening. The sieve orifices are smaller than the dimensions of the powdrous drive charge but allow a sufficient pressure balance. The drive charge fitted in the combustion chamber consists of leaves or strips. The combustion chamber is split into different chambers (11) by perforated transverse wells (5) whose perforations (10) are smaller than the dimensions of the charge. The transverse walls (5) are fixed on a carrier element (4) which is firmly connected with the sieve-type body (3). A middle tube can act as common carrier element (4) for the transverse walls.

Atmospheric scavenging of solid rocket exhaust effluents

Abstract: Solid propellant rocket exhaust was directly utilized to ascertain raindrop scavenging rates for hydrogen chloride. Two chambers were used to conduct the experiments; a large, rigid walled, spherical chamber stored the exhaust constituents, while the smaller chamber housing all the experiments was charged as required with rocket exhaust HCl. Surface uptake experiments demonstrated an HCl concentration dependence for distilled water. Sea water and brackish water HCl uptake was below the detection limit of the chlorine-ion analysis technique used. Plant life HCl uptake experiments were limited to corn and soybeans. Plant age effectively correlated the HCl uptake data. Metallic corrosion was not significant for single 20 minute exposures to the exhaust HCl under varying relative humidity. Characterization of the aluminum oxide particles substantiated the similarity between the constituents of the small scale rocket and the full size vehicles.

Experimental investigation into the L* instability of double base rocket propellants

Abstract: In a continuous joint research program, the Department of Aerospace Engineering of DUT and the Technological Laboratory TNO investigate unstable combustion of solid rocket propellants. Some results of this research program are discussed here. The oscillatory combustion phenomenon is discussed qualitatively, and the interaction between the acoustics of the cavity and the pyrolysis of the propellants is explained. For the investigation of the low-frequency unstable combustion, a special device, the L* burner, is very suited. A description is given of the two L* burners of different size, which were used in the experiments discussed here. Reproducible ignition is of the utmost importance to conduct these tests. Various ignition methods have been tried and are described together with a very successful one, which was evaluated and finally selected. The experimental results of JPN and ARP propellant experiments are discussed extensively. Oscillatory combustion as well as chuffing has been observed, and it turns out that both phenomena are related to the L* and pressure. As L* burners of different sizes were used, it was possible to estimate the damping, and hence the propellant growth constant. On the basis of the results of these experiments it seems most likely that the damping is volumetric. It has been discovered that for a very large pressure range (∼ 3 MPa) the frequency of the oscillations is linearly dependent on the mean pressure, while at low pressures (<1 MPa) oscillations with a higher frequency, also dependent on the mean pressure, are observed. These observed pressurefrequency correlations are the same for the two L* burners of different size, and hence it may be concluded that these correlations are not affected by the dimensions of the motor cavity.

Solid Rocket Motor Combustion Chamber Aeroacoustics

Abstract: ASOURCE of solid rocket motor vibration is investigated by treating the problem of pressure fluctuations inside the motor cavity. Theoretical aeroacoustics is applied to the cavity gases and imbedded burning metal agglomerates. Several critical experiments are performed to provide numerical input to the theory. Consideration of turbulence, combustion, and entropy noise yields the conclusion that only one cause is dominant for the fluctuations in chamber pressure—that of interaction of turbulence with the exhaust nozzle. Typically, an 0.4% rms pressure fluctuation can be accounted for by this mechanism. Spectral distributions of the noise are presented and a comparison is made of the theory and an actual motor firing. Contents