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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 163 records · Page 9
All sourcesNASA NTRSDOE OSTI
NASA NTRS1986

Spacecraft technology requirements for future NASA missions

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.

Hudson, W. R.↗
NASA NTRS1987

Structural design of an in-line bolted joint for the space shuttle solid rocket motor case segments

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.

Dorsey, John T.↗
NASA NTRS1987

Investigation of the acoustic properties of composite materials

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.

Smith, B. T.↗
NASA NTRS1987

Modal survey of the space shuttle solid rocket motor using multiple input methods

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.

Brillhart, Ralph↗
NASA NTRS1988

Detection of bondline delaminations in multilayer structures with lossy components

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.

Madaras, Eric I.↗
NASA NTRS1988

Space shuttle solid rocket motor Profile Measuring Device (PMD)

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.

Redmon, John↗
NASA NTRS1987

Preliminary 2-D shell analysis of the space shuttle solid rocket boosters

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.

Knight, Norman F., Jr.↗
NASA NTRS1988

Lightweight structural design of a bolted case joint for the Space Shuttle Solid Rocket Motor

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.

Dorsey, John T.↗
NASA NTRS1988

Predicting the velocity and azimuth of fragments generated by the range destruction or random failure of rocket casings and tankage

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.

Eck, Marshall↗
NASA NTRS1988

Lightweight structural design of a bolted case joint for the space shuttle solid rocket motor

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.

Dorsey, John T.↗
NASA NTRS1990

Technical Evaluation Motor no. 5 (TEM-5)

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.

Cook, M.↗
NASA NTRS1990

Circularity measuring system: A shape gauge designed especially for use on large objects

The Circularity Measuring System (CMS) was developed to make an in-situ determination of shape similarity for selected fit large cylinders (RSRM segments). It does this to a repeatable accuracy of 0.10 mm (0.004 inch). This is less that the goal of 0.07 mm (0.003 inch), but was determined adequate because of the addition of an assembly aid that increased the entry chamfer of the clevis side of the joint. The usefulness of the CMS is demonstrated by the application to measurements other than its specified design purpose, such as submarine hull circularity, SRM mid-case circularity, as well as circularity of interfacing SRM tooling, specifically the rounding devices and horizontal disassembly devices. Commercialization of the tool is being pursued, since it is an enhancement of metrology technology for circularity determination. The most accurate in-situ technology it replaces is determined from a template. The CMS is an improvement in accuracy and operation.

Rohrkaste, G. R.↗
NASA NTRS1990

Experimental determination of convective heat transfer coefficients in the separated flow region of the Space Shuttle Solid Rocket Motor

A series of cold flow heat transfer tests was conducted with a 7.5-percent scale model of the Space Shuttle Rocket Motor (SRM) to measure the heat transfer coefficients in the separated flow region around the nose of the submerged nozzle. Modifications were made to an existing 7.5 percent scale model of the internal geometry of the aft end of the SRM, including the gimballed nozzle in order to accomplish the measurements. The model nozzle nose was fitted with a stainless steel shell with numerous thermocouples welded to the backside of the thin wall. A transient 'thin skin' experimental technique was used to measure the local heat transfer coefficients. The effects of Reynolds number, nozzle gimbal angle, and model location were correlated with a Stanton number versus Reynolds number correlation which may be used to determine the convective heating rates for the full scale Space Shuttle Solid Rocket Motor nozzle.

Whitesides, R. Harold↗
NASA NTRS1988

Description of a new full scale lightning simulation facility

A lightning test facility has been constructed in Utah, and the facility has initially been used to test and certify solid rocket motor (SRM) segment shipping containers. The simulator equipment consists of a 17-stage 1.7 MV Marx generator and three current generators. The performance of the Marx is enhanced by use of a distributed peaking capacitor to produce maximum dl/dt's in excess of 4 x 10 to the 11th Amps/sec into large test objects. The current generators consist of the following: (1) a 200 KA underdamped 500 kilojoule high current bank, (2) a critically damped intermediate current bank and (3) an over damped continuing current bank. Total charge transfer for these banks is 270 Coulombs with an action integral in excess of 8 x 10 to the 6th sq Amps-sec. Representative results for the simulator and SRM segment shipping containers are presented.

Papazian, Peter B.↗
NASA NTRS1990

Ignition transient analysis of solid rocket motor

To predict pressure-time and thrust-time behavior of solid rocket motors, a one-dimensional numerical model is developed. The ignition phase of solid rocket motors (time less than 0.4 sec) depends critically on complex interactions among many elements, such as rocket geometry, heat and mass transfer, flow development, and chemical reactions. The present model solves the mass, momentum, and energy equations governing the transfer processes in the rocket chamber as well as the attached converging-diverging nozzle. A qualitative agreement with the SRM test data in terms of head-end pressure gradient and the total thrust build-up is obtained. Numerical results show that the burning rate in the star-segmented head-end section and the erosive burning are two important parameters in the ignition transient of the solid rocket motor (SRM).

Han, Samuel S.↗
NASA NTRS1990

Transfer orbit stage mechanisms thermal vacuum test

A systems level mechanisms test was conducted on the Orbital Sciences Corp.'s Transfer Orbit Stage (TOS). The TOS is a unique partially reusable transfer vehicle which will boost a satellite into its operational orbit from the Space Shuttle's cargo bay. The mechanical cradle and tilt assemblies will return to earth with the Space Shuttle while the Solid Rocket Motor (SRM) and avionics package are expended. A mechanisms test was performed on the forward cradle and aft tilting assemblies of the TOS under thermal vacuum conditions. Actuating these assemblies under a 1 g environment and thermal vacuum conditions proved to be a complex task. Pneumatic test fixturing was used to lift the forward cradle, and tilt the SRM, and avionics package. Clinometers, linear voltage displacement transducers, and load cells were used in the thermal vacuum chamber to measure the performance and characteristics of the TOS mechanism assembly. Incorporation of the instrumentation and pneumatic system into the test setup was not routine since pneumatic actuation of flight hardware had not been previously performed in the facility. The methods used are presented along with the problems experienced during the design, setup and test phases.

Oleary, Scott T.↗
NASA NTRS1989

An evaluation of the interfacial bond properties between carbon phenolic and glass phenolic composites

The effects of moisture and surface finish on the mechanical and physical properties of the interfacial bond between the carbon/phenolic (C/P) and glass/phenolic (G/P) composite materials are presented. Four flat panel laminates were fabricated using the C/P and G/P materials. Of the four laminates, one panel was fabricated in which the C/P and G/P materials were cured simultaneously. It was identified as the cocure. The remaining laminates were processed with an initial simultaneous cure of the three C/P billets. Two surface finishes, one on each half, were applied to the top surface. Prior to the application and cure of the G/P material to the machined surface of the three C/P panels, each was subjected to the specific environmental conditioning. Types of conditioning included: (1) nominal fabrication environment, (2) a prescribed drying cycle, and (3) a total immersion in water at 160 F. Physical property tests were performed on specimens removed from the C/P materials of each laminate for determination of the specific gravity, residual volatiles and and resin content. Comparisons of results with shuttle solid rocket motor (SRM) nozzle material specifications verified that the materials used in fabricating the laminates met acceptance criteria and were representative of SRM nozzle materials. Mechanical property tests were performed at room temperature on specimens removed from the G/P, the C/P and the interface between the two materials for each laminate. The double-notched shear strength test was used to determine the ultimate interlaminar shear strength. Results indicate no appreciable difference in the C/P material of the four laminates with the exception of the cocure laminate, where 20 percent reduction in the strength was observed. The most significant effect and the ultimate strength was significantly reduced in the wet material. No appreciable variation was noted between the surface finishes in the wet laminate.

Jordan, Kelvin↗
NASA NTRS1991

A computer program to perform flow and thermal analysis during pressurization of the Space Shuttle Solid Rocket Motor field joint

A computational technique for prediction of the flow and thermal environment in the SRM field joint cavities is described. The SRM field joint hardware was tested with a defect in the insulation, and due to this defect, the O-ring gland cavities are pressurized during the early part of the ignition. A computer model is developed to predict the thermal environment and flow through the simulated flaw, during the pressurization of the field joint. The transient mass, momentum, and energy conservation equations in the flow passage together with the thermodynamic equation of state are solved by a fully implicit iterative numerical procedure.

Clayton, J. Louie↗