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Thrust-terminating system conceived to reduce thrust of solid-propellant rocket motor in controlled manner such that thrust loads not increased or decreased beyond predictable levels. Concept involves explosively cutting opposing venting pairs in case of rocket motor above nozzles to initiate venting of chamber and reduction of thrust. Vents sized and numbered to control amount and rate of reduction in thrust.
The high specific strength and the high specific modules made graphite epoxy laminate an expedient material substitute for the Shuttle Solid Rocket Motor steel case to substantially increase the payload performance without increasing the composite case axial growth during thrust build up which was constrained to minimize liftoff excitation effects on existing structural elements and interfaces. Parameters associated with axial growth were identified for quality and manufacturing controls. Included is an innovative method for experimentally verifying extensional elastic properties on a laminate pressurized test bottle.
An automated inspection system has been developed to inspect the o-ring sealing surfaces on the Space Shuttle reusable solid rocket motor (RSRM) case segment joints. A laser digitizing system is used to create a three dimensional map of the o-ring sealing surfaces. This data is analyzed for any irregularities, which are noted for further inspection and disposition. This paper describes both the hardware and the software used to gather data as well as the methods developed to analyze the data. The RSRM is assembled from four casting segments. Each of the field joints between these segments is sealed with three o-rings. These o-rings are critical for the proper operation of the motor. After launch the booster segments are recovered and reused. As part of the refurbishment process the o-ring sealing surfaces are inspected for any irregularities. The system developed uses a combination of commercial and customized hardware and software. The system uses two computer systems in a real time environment to control a laser, an XYZ precision table, and case rotation and position information. The system is capable of inspecting both the tang end and the clevis end of the RSRM field joint. The o-ring grooves and flat sealing surfaces are inspected.
Orbital behavior of echo i satellite and its rocket casing during its first 500 days
Viscoelastic analysis of structural integrity of accelerated ablating rocket
Gaseous film cooling of rocket motor with injection near throat
Flame propagation inside of cracks in polyester ammonium perchlorate composite propellants
The application of laser holographic interferometric techniques for nondestructive inspection of material structures commonly used in aerospace works is investigated. Two types of structures, composite plate and solid fuel rocket engine motor casing, were examined. In conducting the experiments, both CW HeNe gas lasers and Q-switched ruby lasers were used as light sources for holographic recording setups. Different stressing schemes were investigated as to their effectiveness in generating maximum deformation at regions of structural weakness such as flaws and disbonds. Experimental results on stressing schemes such as thermal stressing, pressurized stressing, transducer excitation, and mechanical impact are presented and evaluated.
Recent developments in the technology of carbon or graphite filaments now provide the designer with greatly improved materials offering high specific strength and modulus. Besides these advantages are properties which are distinctly useful for space applications and which provide feasibility for missions not obtainable by other means. Current applications include major and secondary structures of communications satellites. A number of R & D projects are exploring carbon-fiber application to rocket engine motor cases, advanced antenna systems, and space shuttle components. Future system studies are being made, based on the successful application of carbon fibers for orbiting space telescope assemblies, orbital transfer vehicles, and very large deployable energy generation systems. Continued technology development is needed in analysis, material standards, and advanced structural concepts to exploit the full potential of carbon filaments in composite materials.
Qualification of the full-scale process and design changes for elimination of redesigned solid rocket motor tang nitrile butadiene rubber insulation edge separations and voids was performed from 24 March to 3 December 1990. The objectives of this test were: to qualify design and process changes on flight hardware using a tie ply between the redesigned solid rocket motor steel case and the nitrile butadiene rubber insulation over the tang capture features; to qualify the use of methyl ethyl ketone in the tang flap region to reduce voids; and to determine if holes in the separator film reduce voids in the tang flap region. The tie ply is intended to aid insulation flow during the insulation cure process, and thus reduce or eliminate edge unbonds. Methyl ethyl ketone is intended to reduce voids in the tang flap area by providing better tacking characteristics. The perforated film was intended to provide possible vertical breathe paths to reduce voids in the tang area. Tang tie ply testing consisted of 270 deg of the tang circumference using a new layup method and 90 deg of the tang circumference using the current layup methods. Tie ply process success was defined as a reduction of insulation unbonds. Lack of any insulation edge unbonds on the tang area where the new process was used, and the presence of 17 unbonds with the current process, proves the test to be a success. Successful completion of this test has qualified the new processes.
Dynamic, flexible ceramic seals developed for use at high temperatures in high-performance, variable-geometry, hypersonic airplane engines. Stacked ceramic wafers pressed against stationary sidewall by pressure in one or more metal bellows. Seals also used in hypersonic engines, two-dimensional convergent/divergent and vectored-thrust exhaust nozzles, airframes of reentry vehicles, casings of rocket motors furnaces, and other applications.
The purposes of the tests and analyses described in this report are as follows: (1) determine the lightning current survivability of five alternative changed designs of the bond straps which electrically bond the solid rocket booster (SRB) systems tunnel to the solid rocket motor (SRM) case; (2) determine the amount of reduction in induced voltages on operational flight (OF) tunnel cables obtained by a modified design of tunnel bond straps (both tunnel cover-to-cover and cover-to-motor case); (3) determine the contribution of coupling to the OF tunnel cables by ground electrical and instrumentation (GEI) cables which enter the systems tunnel from unshielded areas on the surfaces of the motor case; and (4) develop a model (based on test data) and calculate the voltage levels at electronic 'black boxes' connected to the OF cables that run in the systems tunnel.
The second annual report on the development and application of contamination technology for MSFC managed space systems covering the period from 10 Dec. 1992 to 10 Dec. 1993 is presented. During this time period, studies were concluded which proved that in-process corrosion protection is not required during Redesigned Solid Rocket Motor (RSRM) case processing. Also completed were a series of tests evaluating the effects of environmental exposure and contamination on 2219-T87 aluminum (Space Shuttle External Tank) OSEE response and bonding properties. Correlations were developed between OSEE response, contamination type/level, and primer adhesion. The results showed that the wet tape and water break free tests currently employed during ET processing may not detect bond affecting levels of some potential contaminants; however, the contaminants were detected with OSEE analysis. Finally, exposure/contamination studies were initiated with HP9-4-30 steel. HP9-4-30 was selected for evaluation because it represents a class of metals commom to MSFC managed space flight systems which are less prone to oxidation than D6AC steel or aluminum.
This presentation describes extraction and integration of requirements information and safety information in visualizations to support early review of completeness, correctness, and consistency of lengthy and diverse system safety analyses. Software tools have been developed and extended to perform the following tasks: 1) extract model parts and safety information from text in interface requirements documents, failure modes and effects analyses and hazard reports; 2) map and integrate the information to develop system architecture models and visualizations for safety analysts; and 3) provide model output to support virtual system integration testing. This presentation illustrates the methods and products with a rocket motor initiation case.
An all-new sounding rocket was designed at NASA's Marshall Space Flight Center that featured an aft finocyl, aluminized solid propellant grain and silica-filled ethylene-propylene-diene monomer (SFEPDM) internal insulation. Upon the initial static firing of the first of this new design, the solid rocket motor (SRM) case failed thermally just upstream of the aft closure early in the burn time. Subsequent fluid modeling indicated that the high-velocity combustion-product jets emanating from the fin-slots in the propellant grain were likely inducing a strongly swirling flow, thus substantially increasing the severity of the convective environment on the exposed portion of the SFEPDM insulation in this region. The aft portion of the fin-slots in another of the motors were filled with propellant to eliminate the possibility of both direct jet impingement on the exposed SFEPDM and the appearance of strongly swirling flow in the aft region of the motor. When static-fired, this motor's case still failed in the same axial location, and, though somewhat later than for the first static firing, still in less than 1/3rd of the desired burn duration. These results indicate that the extreme material decomposition rates of the SFEPDM in this application are not due to gas-phase convection or shear but rather to interactions with burning aluminum or alumina slag. Further comparisons with between SFEPDM performance in this design and that in other hot-fire tests provide insight into the mechanisms of SFEPDM decomposition in SRM aft domes that can guide the upcoming redesign effort, as well as other future SRM designs. These data also highlight the current limitations of modeling elastomeric insulators solely with diffusion-controlled, gas-phase thermochemistry in SRM regions with significant viscous shear and/or condense-phase impingement or flow.
Using three different current injection methods and a simple voltage probe, transfer impedances for Solid Rocket Motor (SRM) joints, wire meshes, aluminum foil, Thorstrand and a graphite composite motor case were measured. In all cases, the surface current distribution for the particular current injection device was calculated analytically or by finite difference methods. The results of these calculations were used to generate a geometric factor which was the ratio of total injected current to surface current density. The results were validated in several ways. For wire mesh measurements, results showed good agreement with calculated results for a 14 by 18 Al screen. SRM joint impedances were independently verified. The filiment wound case measurement results were validated only to the extent that their curve shape agrees with the expected form of transfer impedance for a homogeneous slab excited by a plane wave source.
The pin retainer band is a 1-1/4 in. wide by about 1/32 in. thick steel band that is stretched over the pins circumferentially at the joints of the solid rocket booster (SRB) motor case segments. Its purpose is to hold the retaining pins in place during flight. Since the band is thin and is mounted stretched, or under tension, over the pins which are greased in place, it is necessary to protect it thermally. Three typical areas of the band location on the SRB motor case joints were chosen for verifying the cork thermal prection system (TPS) performance on the band: the joint in front of the attach ring, at the aft skirt joint in front of the kick ring and in an unperturbed area at the forward segments of the SRB motor cases. The kick ring area pin retainer band cork TPS passed the test requirements in the region where there are no phenolic splice caps. In the splice cap areas where the band TPS is closed out around the splice with K5NA, additional amounts of K5NA material is required to adequately protect the steel band. TPS requirements were easily met in the one location where the ends of the steel band are joined together after installation and tensioning. The attack ring area pin retainer band TPS did not quite pass the heat load requirements with the .375 in. thick cork TPS as used for the kick ring area band. It is recommended that the cork thickness here be increased to .50 in. The unperturbed area band TPS presented no problems.
A technique for determining the average erosion rate versus Mach number of candidate internal insulation materials was developed for flight motor applications in 12 inch I.D. test firing hardware. The method involved the precision mounting of a mechanical measuring tool within a conical test cartridge fabricated from either a single insulation material or two non-identical materials each of which constituted one half of the test cartridge cone. Comparison of the internal radii measured at nine longitudinal locations and between eight to thirty two azimuths, depending on the regularity of the erosion pattern before and after test firing, permitted calculation of the average erosion rate and Mach number. Systematic criteria were established for identifying erosion anomalies such as the formation of localized ridges and for excluding such anomalies from the calculations. The method is discussed and results presented for several asbestos-free materials developed in-house for the internal motor case insulation in solid propellant rocket motors.