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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.
Electret devices were used to collect samples of rocket exhaust products. The samples were analyzed by X-ray spectroscopy. The performance of the electrets was compared with that of Thiokol fixed flow samplers during the static test firing of a solid rocket motor (SRM) demonstration model. A vertical profile of wind speed, wind direction, and air temperature, used as inputs to the multilayer diffusion model for the SRM static test is presented.
The computational procedure for signal propagation in the presence of an exhaust plume is presented. Comparisons with well-known analytic diffraction solutions indicate that accuracy suffers when mesh spacing is inadequate to resolve the first unobstructed Fresnel zone at the plume edge. Revisions to the procedure to improve its accuracy without requiring very large arrays are discussed. Comparisons to field measurements during a shuttle solid rocket motor (SRM) test firing suggest that the plume is sharper edged than one would expect on the basis of time averaged electron density calculations. The effects, both of revisions to the computational procedure and of allowing for a sharper plume edge, are to raise the signal level near tail aspect. The attenuation levels then predicted are still high enough to be of concern near SRM burnout for northerly launches of the space shuttle.
Details of the design, operation, testing and recovery procedures of the reusable solid rocket boosters (SRB) are given. Using a composite PBAN propellant, they will provide the primary thrust (six million pounds maximum at 20 s after ignition) within a 3 g acceleration constraint, as well as thrust vector control for the Space Shuttle. The drogues were tested to a load of 305,000 pounds, and the main parachutes to 205,000. Insulation in the solid rocket motor (SRM) will be provided by asbestos-silica dioxide filled acrylonitrile butadiene rubber ('asbestos filled NBR') except in high erosion areas (principally in the aft dome), where a carbon-filled ethylene propylene diene monomer-neopreme rubber will be utilized. Furthermore, twenty uses for the SRM nozzle will be allowed by its ablative materials, which are principally carbon cloth and silica cloth phenolics.
A pattern search technique is presented, which is utilized in a computer program that minimizes the sum of the squares of the differences, at various times, between a desired thrust-time trace and that calculated with a special mathematical internal ballistics model of a solid propellant rocket motor. The program is demonstrated by matching the thrust-time trace obtained from static tests of the first Space Shuttle SRM starting with input values of 10 variables which are, in general, 10% different from the as-built SRM. It is concluded that an excellent match is obtained.
The Space Transportation (STS) launch vehicles are more complex than other launch vehicles. A comprehensive analysis of the lift-off event is required so that the responses of the vehicle and payloads can be predicted. The transient overpressure induced by ignition of the solid rocket motors (SRM's) is one of the critical design factors assessed for lift-off. The STS-1 ignition overpressure design environment was developed from 6.4-percent scale-model tests in which the Tomahawk solid rocket motors were used to simulate the start-up process of the SRM's. The overpressures measured during STS-1 lift-off were much more severe than predicted. The reasons for this anomaly are discussed in this paper. The 6.4-percent scale model was redesigned and used as a tool to develop an effective ignition overpressure suppression system for STS-2 and subsequent flights. Also presented are advancements in subscale-model simulation and theoretical understanding of this transient overpressure phenomenon that led to the successful development of the fix.
An analysis of the Shuttle SRM thrust imbalance during the steady-state and tailoff portions of the boost phase of flight are presented. Results from flights STS-1 through STS-13 are included. A statistical analysis of the observed thrust imbalance data is presented. A 3 sigma thrust imbalance history versus time was generated from the observed data and is compared to the vehicle design requirements. The effect on Shuttle thrust imbalance from the use of replacement SRM segments is predicted. Comparisons of observed thrust imbalances with respect to predicted imbalances are presented for the two space shuttle flights which used replacement aft segments (STS-9 and STS-13).
Missions selected by NASA as representative of the most critical technology issues are described and the key technology requirements are discussed. The following are considered: (1) the Large Deployable Reflector (LDR), (2) the Sample Return Missions (SRM), (3) the Earth Observing System (EOS), and (4) the Geostationary Communications Platform (GCP). The LDR, designed for infrared and submillimeter astronomical measurements, will contain the following components for its instruments: indium antimonide charge-coupled arrays, extrinsic photodetector arrays, and heterodyne photoconductors. New technology developed for SRM includes an X-band transponder, a fiber-optic rotation sensor, and long-life batteries. The EOS will consist of instruments designed to measure the earth's atmosphere, interior, and surface, as well as an information system to support scientific research. Inherent in the platform of the GCP is the ability to effectively utilize both the geostationary arc and the electromagnetic spectrum; the GCP can also reduce overall satellite communications system costs.
Results of a structural design study of an in-line bolted joint concept which can be used to assemble Space Shuttle Solid Rocket Motor (SRM) case segments are presented. Numerous parametric studies are performed to characterize the in-line bolted joint behavior as major design variables are altered, with the primary objective always being to keep the inside of the joint (where the O-rings are located) closed during the SRM firing. The resulting design has 180 1-inch studs, an eccentricity of -0.5 inch, a flange thickness of 3/4 inch, a bearing plate thickness of 1/4 inch, and the studs are subjected to a preload which is 70% of ultimate. The mass penalty per case segment joint for the in-line design is 346 lbm more than the weight penalty for the proposed capture tang fix.
An effort to provide inspection of the bondlines in the Space Shuttle Solid Rocket Motor (SRM) is discussed. Investigation of the bondlines in the membrane region of the SRM has centered on representative samples provided by Morton Thiokol. These samples have pull-tabs, or in one case grafoil inserts, to simulate delaminations at various points in the multilayer structure of steel case, liner, insulation, liner, and propellant. Discussed here is the continuing approach to examining these defects which is predicated on a model of an ultrasonic wave propagating in a lossy multilayer resonant structure, the experimental arrangements used, and the results of these experiments.
The ability to accurately characterize propellant in a finite element model is a concern of engineers tasked with studying the dynamic response of the Space Shuttle Solid Rocket Motor (SRM). THe uncertainties arising from propellant characterization through specimem testing led to the decision to perform a model survey and model correlation of a single segment of the Shuttle SRM. Multiple input methods were used to excite and define case/propellant modes of both an inert segment and, later, a live propellant segment. These tests were successful at defining highly damped, flexible modes, several pairs of which occured with frequency spacing of less than two percent.
The detection of bondline delaminations in multilayer structures using ultrasonic reflection techniques is a generic problem in adhesively bonded composite structures such as the Space Shuttles's Solid Rocket Motors (SRM). Standard pulse echo ultrasonic techniques do not perform well for a composite resonator composed of a resonant layer combined with attenuating layers. Excessive ringing in the resonant layer tends to mask internal echoes emanating from the attenuating layers. The SRM is made up of a resonant steel layer backed by layers of adhesive, rubber, liner and fuel, which are ultrasonically attenuating. The structure's response is modeled as a lossy ultrasonic transmission line. The model predicts that the acoustic response of the system is sensitive to delaminations at the interior bondlines in a few narrow frequency bands. These predictions are verified by measurements on a fabricated system. Successful imaging of internal delaminations is sensitive to proper selection of the interrogating frequency. Images of fabricated bondline delaminations are presented based on these studies.
The SRM PMD is an electromechanical tool used for measuring and recording the profile and diameters of the solid rocket motor segments, both Tang and Clevis ends. This system consists of a crossbeam assembly that mounts to the SRM segment using the existing assembly pin holes. The mounting configuration is such that the tool can be used to measure Clevis up/Tang down or Clevis up/Tang down. The testing and calibration of the PMD is described.
A two-dimensional shell model of an entire solid rocket booster (SRB) has been developed using the STAGSC-1 computer code and executed on the Ames CRAY computer. The purpose of these analyses is to calculate the overall deflection and stress distributions for the SRB when subjected to mechanical loads corresponding to critical times during the launch sequence. The mechanical loading conditions for the full SRB arise from the external tank (ET) attachment points, the solid rocket motor (SRM) pressure load, and the SRB hold down posts. The ET strut loads vary with time after the Space Shuttle main engine (SSME) ignition. The SRM internal pressure varies axially by approximately 100 psi. Static analyses of the full SRB are performed using a snapshot picture of the loads. The field and factory joints are modeled by using equivalent stiffness joints instead of detailed models of the joint. As such, local joint behavior cannot be obtained from this global model.
This paper presents the structural design of a bolted joint with a static face seal which can be used to join Space Shuttle Solid Rocket Motor (SRM) case segments. Results from numerous finite element parametric studies indicate that the bolted joint meets the design requirement of preventing joint opening at the O-ring locations during SRM pressurization. A final design recommended for further development has the following parameters: 180 1-inch-diameter studs, stud center line offset of .5 inches radially inward from the shell wall center line, flange thickness of 0.75 inches, bearing plate thickness of 0.25 inches, studs prestressed to 70 percent of ultimate load, and the intermediate alcove. The design has a mass penalty of 1096 lbm, which is 164 lbm greater than the currently proposed capture tang redesign.
A calculational method is described which provides a powerful tool for predicting solid rocket motor (SRM) casing and liquid rocket tankage fragmentation response. The approach properly partitions the available impulse to each major system-mass component. It uses the Pisces code developed by Physics International to couple the forces generated by an Eulerian-modeled gas flow field to a Lagrangian-modeled fuel and casing system. The details of the predictive analytical modeling process and the development of normalized relations for momentum partition as a function of SRM burn time and initial geometry are discussed. Methods for applying similar modeling techniques to liquid-tankage-overpressure failures are also discussed. Good agreement between predictions and observations are obtained for five specific events.
The structural design of a bolted joint with a static face seal which can be used to join Space Shuttle Solid Rocket Motor (SRM) case segments is given. Results from numerous finite element parametric studies indicate that the bolted joint meets the design requirement of preventing joint opening at the O-ring locations during SRM pressurization. A final design recommended for further development has the following parameters: 180 one-in.-diam. studs, stud centerline offset of 0.5 in radially inward from the shell wall center line, flange thickness of 0.75 in, bearing plate thickness of 0.25 in, studs prestressed to 70 percent of ultimate load, and the intermediate alcove. The design has a mass penalty of 1096 lbm, which is 164 lbm greater than the currently proposed capture tang redesign.
Technical Evaluation Motor No. 5 (TEM-5) was static test fired at the Thiokol Corporation Static Test Bay T-97. TEM-5 was a full scale, full duration static test fire of a high performance motor (HPM) configuration solid rocket motor (SRM). The primary purpose of TEM static tests is to recover SRM case and nozzle hardware for use in the redesigned solid rocket motor (RSRM) flight program. Inspection and instrumentation data indicate that the TEM-5 static test firing was successful. The ambient temperature during the test was 41 F and the propellant mean bulk temperature (PMBT) was 72 F. Ballistics performance values were within the specified requirements. The overall performance of the TEM-5 components and test equipment was nominal. Dissembly inspection revealed that joint putty was in contact with the inner groove of the inner primary seal of the ignitor adapter-to-forward dome (inner) joint gasket; this condition had not occurred on any previous static test motor or flight RSRM. While no qualification issues were addressed on TEM-5, two significant component changes were evaluated. Those changes were a new vented assembly process for the case-to-nozzle joint and the installation of two redesigned field joint protection systems. Performance of the vented case-to-nozzle joint assembly was successful, and the assembly/performance differences between the two field joint protection system (FJPS) configurations were compared.