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At least 541 records · Page 30
All sourcesNASA NTRSDOE OSTI
NASA NTRS1999

Thermal Characterization of Adhesive

The current Space Shuttle Reusable Solid Rocket Motor (RSRM) nozzle adhesive bond system is being replaced due to obsolescence. Down-selection and performance testing of the structural adhesives resulted in the selection of two candidate replacement adhesives, Resin Technology Group's Tiga 321 and 3M's EC2615XLW. This paper describes rocket motor testing of these two adhesives. Four forty-pound charge motors were fabricated in configurations that would allow side by side comparison testing of the candidate replacement adhesives and the current RSRM adhesives. The motors provided an environment where the thermal performance of adhesives in flame surface bondlines was compared. Results of the FPC testing show that: 1) The phenolic char depths on radial bond lines is approximately the same and vary depending on the position in the blast tube regardless of which adhesive was used; 2) The adhesive char depth of the candidate replacement adhesives is less than the char depth of the current adhesives; 3) The heat-affected depth of the candidate replacement adhesives is less than the heat-affected depth of the current adhesives; and 4) The ablation rates for both replacement adhesives are slower than that of the current adhesives.

Spomer, Ken A.↗
NASA NTRS2000

NARC Rayon Replacement Program for the Space Shuttle Reusable Solid Rocket Motor Nozzle: Screening Summary

Thiokol Corporation and NASA MSFC are jointly developing a replacement for North American Rayon Corporation (NARC) Aerospace Grade Rayon (1650/720 continuous filament), the precursor for the Carbon Cloth Phenolic (CCP) ablatives used in the Space Shuttle Reusable Solid Rocket Motor (RSRM) Nozzles. NARC discontinued production of Aerospace Grade Rayon in September 1997. NASA maintains a stockpile of NARC Rayon to support RSRM production through the summer of 2005. The program plan for selection and qualification of a replacement for NARC rayon was approved in August 1998. Screening activities began in February 1999. The intent of this paper is to provide a summary of the data generated during the screening phase of the NARC Rayon Replacement Program. Twelve cellulose based fibers (rayon and lyocell) were evaluated. These fibers were supplied by three independent vendors. Many of these fibers were carbonized by two independent carbonizers. Each candidate was tested according to standard acceptance test methods at each step of the manufacturing process. Additional testing was performed with the candidate CCPS, including hot fire tests, Process studies and mechanical and thermal characterization. Six of the twelve fiber candidates tested were dropped at the conclusion of Phase 1. The reasons for the elimination of these candidates included; difficulties in processing the material in the whitegoods, carbon and CCP forms; poor composite mechanical performance; and future availability concerns. The remaining six fibers demonstrated enough promise to merit continued evaluation and optimization of the CCP fabrication process. Note: Certain CCP data falls under the restrictions of US export laws, (ITAR, etc.) and will not be included in this paper.

Cook, R. V.↗
NASA NTRS2000

Upper Stage Flight Experiment 10K Engine Design and Test Results

A 10,000 lbf thrust chamber was developed for the Upper Stage Flight Experiment (USFE). This thrust chamber uses hydrogen peroxide/JP-8 oxidizer/fuel combination. The thrust chamber comprises an oxidizer dome and manifold, catalyst bed assembly, fuel injector, and chamber/nozzle assembly. Testing of the engine was done at NASA's Stennis Space Center (SSC) to verify its performance and life for future upper stage or Reusable Launch Vehicle applications. Various combinations of silver screen catalyst beds, fuel injectors, and combustion chambers were tested. Results of the tests showed high C* efficiencies (97% - 100%) and vacuum specific impulses of 275 - 298 seconds. With fuel film cooling, heating rates were low enough that the silica/quartz phenolic throat experienced minimal erosion. Mission derived requirements were met, along with a perfect safety record.

Ross, R.↗
NASA NTRS2000

Development of Thermal Barriers For Solid Rocket Motor Nozzle Joints

Joints in the Space Shuttle solid rocket motors are sealed by O-rings to contain combustion gases inside the rocket that reach pressures of up to 900 psi and temperatures of up to 5500 F. To provide protection for the O-rings, the motors are insulated with either phenolic or rubber insulation. Gaps in the joints leading up to the O-rings are filled with polysulfide joint-fill compounds as an additional level of protection. The current RSRM nozzle-to-case joint design incorporating primary, secondary, and wiper O-rings experiences gas paths through the joint-fill compound to the innermost wiper O-ring in about one out of every seven motors. Although this does not pose a safety hazard to the motor, it is an undesirable condition that NASA and rocket manufacturer Thiokol want to eliminate. Each nozzle-to-case joint gas path results in extensive reviews and evaluation before flights can be resumed. Thiokol and NASA Marshall are currently working to improve the nozzle-to-case joint design by implementing a more reliable J-leg design that has been used successfully in the field and igniter joint. They are also planning to incorporate the NASA Glenn braided carbon fiber thermal barrier into the joint. The thermal barrier would act as an additional level of protection for the O-rings and allow the elimination of the joint-fill compound from the joint.

Steinetz, Bruce M.↗
NASA NTRS1994

Aerothermodynamics of Pyrolizing Surfaces in Hypersonic Rarefied Flows

Direct simulation Monte Carlo (DSMC) calculations of rarefied flows about entry bodies typically employ a fixed surface temperature or a radiative-equilibrium energy balance to compute that temperature. Such boundary conditions neglect any effects of heat capacitance and heat conduction in the spacecraft heat shield and, therefore, provide an upper bound for the surface temperature. Such calculations also neglect pyrolysis from the heat shield which can be significant for a high-energy incident flow at very low densities. Accurate prediction of both heating and aerodynamic forces requires including pyrolysis and surface heat transfer in the models for gas-surface interaction employed in DSMC methods. Although these physical models have long appeared in various continuum flow calculation codes, they have only recently appeared in DSMC codes which are required to simulate rarefied flows during entry at high altitudes. In the current implementation, routines from the widely distributed Charring Material Thermal Response and Ablation (CMA) program are coupled into a DSMC code to calculate the one-dimensional heat transfer into the carbon phenolic heat shield at each point on a vehicle surface. Temperature-dependent material properties, surface re-radiation, and in-depth pyrolysis were included in the calculation, but surface ablation was neglected. Sample calculations for entry of the Galileo probe into the atmosphere of Jupiter demonstrate that including pyrolysis in the model leads to significant differences in predicted aerodynamics. Granted, the drag coefficient does not depend strongly on the surface temperature which can itself be significantly below the radiative equilibrium value during entry. However, the surface mass flux due to pyrolysis of the material is significant once the probe drops to altitudes characterized by transition flow. This leads to a noticeable increase in drag and a decrease in heating compared to a body without pyrolysis.

Haas, Brian L.↗
NASA NTRS2001

Method of Preparing Polymers with Low Melt Viscosity

This invention is an improvement in standard polymerizations procedures, i.e., addition-type and step-growth type polymerizations, wherein monomers are reacted to form a growing polymer chain. The improvement includes employing an effective amount of a trifunctional monomer (such as a trifunctional amine anhydride, or phenol) in the polymerization procedure to form a mixture of polymeric materials consisting of branced polymers, star-shaped polymers, and linear polymers. This mixture of polymeric materials has a lower melt temperature and a lower melt viscosity than corresponding linear polymeric materials of equivalent molecular weight.

Jensen, Brian J.↗
NASA NTRS1994

Light Weight Ceramic Ablators for Mars Follow-on Mission Vehicle Thermal Protection System

New Light Weight Ceramic Ablators (LCA) were produced by using ceramic and carbon fibrous substrates, impregnated with silicone and phenolic resins. The special infiltration techniques (patent pending) were developed to control the amount of organic resins in the highly porous fiber matrices so that the final densities of LCA's range from 0.22 to 0.24 g/cc. This paper presents the thermal and ablative performance of the Silicone Impregnated Reusable Ceramic Ablators (SIRCA) in simulated entry conditions for Mars-Pathfinder in the Ames 60 MW Interaction Heating Facility (I HF). Arc jet test results yielded no evidence of char erosion and mass loss at high stagnation pressures to 0.25 atm. Minimal silica melt was detected on surface char at a stagnation pressure of 0.31 atm. Four ceramic substrates were used in the production of SIRCA's to obtain the effective of boron oxide present in substrate so the thermal performance of SIRCA's. A sample of SIRCA was also exposed to the same heating condition for five cycles and no significant mass loss or recession was observed. Tensile testing established that the SIRCA tensile strength is about a factor of two higher than that of the virgin substrates. Thermogravimetric Analysis (TGA) of the char in nitrogen and air showed no evidence of free carbon in the char. Scanning Electron Microscopy of the post test sample showed that the char surface consists of a fibrous structure that was sealed with a thin layer of silicon oxide melt.

Tran, Huy K.↗
NASA NTRS2001

Determining CDOM Absorption Spectra in Diverse Aquatic Environments Using a Multiple Pathlength, Liquid Core Waveguide System

We evaluated the accuracy, sensitivity and precision of a multiple pathlength, liquid core waveguide (MPLCW) system for measuring colored dissolved organic matter (CDOM) absorption in the UV-visible spectral range (370-700 nm). The MPLCW has four optical paths (2.0, 9.8, 49.3, and 204 cm) coupled to a single Teflon AF sample cell. Water samples were obtained from inland, coastal and ocean waters ranging in salinity from 0 to 36 PSU. Reference solutions for the MPLCW were made having a refractive index of the sample. CDOM absorption coefficients, aCDOM, and the slope of the log-linearized absorption spectra, S, were compared with values obtained using a dual-beam spectrophotometer. Absorption of phenol red secondary standards measured by the MPLCW at 558 nm were highly correlated with spectrophotometer values and showed a linear response across all four pathlengths. Values of aCDOM measured using the MPLCW were virtually identical to spectrophotometer values over a wide range of concentrations. The dynamic range of aCDOM for MPLCW measurements was 0.002 - 231.5 m-1. At low CDOM concentrations spectrophotometric aCDOM were slightly greater than MPLCW values and showed larger fluctuations at longer wavelengths due to limitations in instrument precision. In contrast, MPLCW spectra followed an exponential to 600 nm for all samples.

Miller, Richard L.↗
NASA NTRS1994

Vapor Phase Catalytic Ammonia Reduction

This paper discusses the development of a Vapor Phase Catalytic Ammonia Reduction (VPCAR) teststand and the results of an experimental program designed to evaluate the potential of the technology as a water purification process. In the experimental program the technology is evaluated based upon product water purity, water recovery rate, and power consumption. The experimental work demonstrates that the technology produces high purity product water and attains high water recovery rates at a relatively high specific power consumption. The experimental program was conducted in 3 phases. In phase I an Igepon(TM) soap and water mixture was used to evaluate the performance of an innovative Wiped-Film Rotating-Disk evaporator and associated demister. In phase II a phenol-water solution was used to evaluate the performance of the high temperature catalytic oxidation reactor. In phase III a urine analog was used to evaluate the performance of the combined distillation/oxidation functions of the processor.

Flynn, Michael T.↗
NASA NTRS2001

Sulfate- and Sulfur-Reducing Bacteria as Terrestrial Analogs for Microbial Life on Jupiter's Satellite Io

Observations from the Voyager and Galileo spacecraft have revealed Jupiter's moon Io to be the most volcanically active body of our Solar System. The Galileo Near Infrared Imaging Spectrometer (NIMS) detected extensive deposits of sulfur compounds, elemental sulfur and SO2 frost on the surface of Io. There are extreme temperature variations on Io's surface, ranging from -130 C to over 2000 C at the Pillan Patera volcanic vent. The active volcanoes, fumaroles, calderas, and lava lakes and vast sulfur deposits on this frozen moon indicate that analogs of sulfur- and sulfate-reducing bacteria might inhabit Io. Hence Io may have great significance to Astrobiology. Earth's life forms that depend on sulfur respiration are members of two domains: Bacteria and Archaea. Two basic links of the biogeochemical sulfur cycle of Earth have been studied: 1) the sulfur oxidizing process (occurring at aerobic conditions) and 2) the process of sulfur-reduction to hydrogen sulfide (anaerobic conditions). Sulfate-reducing bacteria (StRB) and sulfur-reducing bacteria (SrRB) are responsible for anaerobic reducing processes. At the present time the systematics of StRB include over 112 species distributed into 35 genera of Bacteria and Archaea. Moderately thermophilic and mesophilic SrRB belong to the Bacteria. The hyperthermophilic SrRB predominately belong to the domain Archaea and are included in the genera: Pyrodictium, Thermoproteus, Pyrobaculum, Thermophilum, Desulfurococcus, and Thermodiscus. The StRB and SrRB use a wide spectrum of substrates as electron donors for lithotrophic and heterotrophic type nutrition. The electron acceptors for the StRB include: sulfate, thiosulfate, sulfite, sulfur, arsenate, dithionite, tetrathionate, sulfur monoxide, iron, nitrite, selenite, fumarate, oxygen, carbon dioxide, and chlorine-containing phenol compounds. The Sulfate- and Sulfur-reducing bacteria are widely distributed in anaerobic ecosystems, including extreme environments like hot springs, deepsea hydrothermal vents, soda and high salinity lakes, and cryo-environments. Furthermore, the StRB and SrRB have Astrobiological significance as these anaerobic extremophiles may represent the dominant relic life forms that inhabited our planet during the extensive volcanic activity in the Earth's early evolutionary period.

Pikuta, Elena V.↗
NASA NTRS2001

Thermal Characterization of Epoxy Adhesive by Hotfire Testing

This paper describes subscale solid-rocket motor hot-fire testing of epoxy adhesives in flame surface bondlines to evaluate heat-affected depth, char depth and ablation rate. Hot-fire testing is part of an adhesive down-selection program on the Space Shuttle Solid Rocket Motor Nozzle to provide additional confidence in the down-selected adhesives. The current nozzle structural adhesive bond system is being replaced due to obsolescence. Prior to hot-fire testing, adhesives were tested for chemical, physical and mechanical properties, which resulted in the selection of two potential replacement adhesives, Resin Technology Group's TIGA 321 and 3M's EC2615XLW. Hot-fire testing consisted of four forty-pound charge (FPC) motors fabricated in configurations that would allow side-by-side comparison testing of the candidate replacement adhesives with the current RSRM adhesives. Results of the FPC motor testing show that: 1) the phenolic char depths on radial bondlines is approximately the same and vary depending on the position in the blast tube regardless of which adhesive was used, 2) the replacement candidate adhesive char depths are equivalent to the char depths of the current adhesives, 3) the heat-affected depths of the candidate and current adhesives are equivalent, and 4) the ablation rates for both replacement adhesives were equivalent to the current adhesives.

Spomer, Ken A.↗
NASA NTRS1995

Galileo Heat Shield Response Experiment Data Reduction and Analyses

The Galileo heat shield response experiment consists of 10 Analog Resistance Ablation Detectors (ARADs) embedded in the carbon phenolic heat shield. As the vehicle descends into the atmosphere of Jupiter, these ARADs will in effect measure the recession as a function of time at various locations on the heat shield. This recession data will be used to reconstruct the time-dependent shape of the vehicle. The shape data is of critical importance to the Atmospheric Reconstruction Experiment which must know mass and drag coefficient of the vehicle in order to determine the structure of the Jovian atmosphere. The data is also intrinsically useful as a database for evaluating the accuracy of our numerical codes and methodology for predicting highly radiating reentry flowfields and coupled heat shield material response. This paper will document the reduction of the data from the ARAD sensors, reconstruction of the heat shield shape history, and comparison of the actual heat shield response with previous calculations and experiments. If possible, we will also include results from new calculations which couple the shock layer flow with the transient material response.

Milos, Frank S.↗
NASA NTRS1995

Estimates of Heat Shielding Requirements for Saturn Probe Missions

Entry trajectories, decelerations, heating and heat shielding requirements of probes entering the atmosphere of Saturn at relative entry velocities ranging from 26.5 - 33.0 km/sec were studied. Much of the Galileo Jupiter probe technology was applicable to these outer planet probes; therefore, the same configuration was used. A wide range entry flight path angles were used in the trajectory analysis in order to satisfy the complete range of arrival conditions and trajectory requirements. Various entry probe sizes and a range of entry masses from 32 - 258 kg were considered. The carbon phenolic heat shield material used on the Galileo and the successful Pioneer-Venus probes was selected. Turbulent boundary layer convection dominated the heating of the Saturn probes; radiative heating was negligible by comparison at the entry velocities considered. The peak heating rates were found to be about 4 kW per square centimeter, or approximately 17% of the Galileo probe's expected maximum value. The forebody heat shield mass fractions for the Saturn probes were found to vary from about 0.18 to 0.28. By comparison, the Pioneer-Venus probes' heat shield mass fractions ranged from 0.10 to 0.13 and the Galileo probe's forebody value is 0.43.

Wercinski, P.↗
NASA NTRS1997

TPS Materials for Planetary Exploration

This paper will review the current material options, and near term developments, for thermal protection systems (TPS) of future space craft involving atmospheric entry. TPS is a very important element in the design of these vehicles, and can have large impact on the project risk and cost. The new family of light weight ceramic ablators illustrated by SIRCA and PIC will be discussed, along with state-of-the-art ablators including SLA-561V, Avcoat-5026, carbon-phenolic, and carbon-carbon.

Rasky, Daniel J.↗
NASA NTRS1996

PICA Forebody Heatshield Qualification for the Stardust Discovery Class Mission

This paper presents the qualification of the light weight Phenolic Impregnated Carbon Ablators (PICA) as the forebody heatshield for the Stardust Discovery Class Mission. The Stardust spacecraft will be launched in early 1999 and fly by Comet Wild-2 to collect cometary and interstellar dust and return them back to earth in the Sample Return Capsule (SRC). This earth re-entry will be the fastest to date, at 12.6 km/s, and therefore requires a heatshield that can withstand very high heating rates and stagnation pressures, as well as provide the necessary insulation to the vehicle structure. The PICA material was developed as part of the Lightweight Ceramic Ablators program at NASA Ames Research Center, and was baselined as the forebody heatshield because of its low density and superior ablation and thermal performance at severe aerothermodynamic conditions. Under a Small Business Innovative Research (SBIR) program with NASA Ames, Fiber Materials, Inc. developed a process to manufacture a single-piece PICA heatshield for the forebody of the SRC, along with witness material for the fabrication of the test models. The test models were fabricated and instrumented by the staff of Lockheed Martin Astronautics in Denver, Colorado. Full body preliminary aerothermal CFD calculations were performed at NASA Ames to determine the heating and stagnation pressure conditions. The Heat shield sizing was also performed at NASA Ames by using a new material response code that accounts for the highly porous characteristics of the PICA material. The ablation and thermal performance of PICA was qualified in the NASA Ames Interaction Heating Arc Jet Facility. A total of 24 models and four test conditions were used to qualify PICA at the predicted peak heat flux, heat load, shear, and stagnation pressure conditions. Surface and in-depth temperatures were measured using optical pyrometers and thermocouples. Surface recession was measured by using a template and a height gage. Several models were tested to evaluate repair procedures, and two models were cold soaked in liquid nitrogen, prior th testing, to investigate the effect of the cold space environment on the performance of the material. In addition, material cored from a demonstration single-piece heatshield was tested to verify that the PICA process can be successfully completed on a large, complex heatshield shape.

Tran, Huy K.↗
NASA NTRS2002

Thermal Barrier/Seal for Extreme Temperature Applications

Large solid rocket motors, as found on the Space Shuttle, are fabricated in segments for manufacturing considerations, bolted together, and sealed using conventional Viton O-ring seals. Similarly the nine large solid rocket motor nozzles are assembled from several different segments, bolted together, and sealed at six joint locations using conventional O-ring seals. The 5500 F combustion gases are generally kept a safe distance away from the seals by thick layers of phenolic or rubber insulation. Joint-fill compounds, including RTV (room temperature vulcanized compound) and polysulfide filler, are used to fill the joints in the insulation to prevent a direct flow-path to the O-rings. Normally these two stages of protection are enough to prevent a direct flow-path of the 900-psi hot gases from reaching the temperature-sensitive O-ring seals. However, in the current design 1 out of 15 Space Shuttle solid rocket motors experience hot gas effects on the Joint 6 wiper (sacrificial) O-rings. Also worrisome is the fact that joints have experienced heat effects on materials between the RTV and the O-rings, and in two cases O-rings have experienced heat effects. These conditions lead to extensive reviews of the post-flight conditions as part of the effort to monitor flight safety. We have developed a braided carbon fiber thermal barrier to replace the joint fill compounds in the Space Shuttle solid rocket motor nozzles to reduce the incoming 5500 F combustion gas temperature and permit only cool (approximately 100 F) gas to reach the temperature-sensitive O-ring seals. Implementation of this thermal barrier provides more robust, consistent operation with shorter turn around times between Shuttle launches.

Steinetz, Bruce M.↗
NASA NTRS2002

NARC Rayon Replacement Program

This viewgraph presentation provides information on the search for a replacement product for the aerospace grade rayon formerly used in the reusable solid rocket motor (RSRM). The rayon used in RSRMs is no longer produced by the North American Rayon Corporation (NARC), though NASA has stockpiled rayon to be used until a suitable replacement is found to be used in the RSRM nozzle carbon cloth phenolic (CCP). The primary considerations are performance, process ability, variability, long term availability, and cost. The preference will be given to commercial off-the-shelf products as opposed to those which are custom made. The candidate materials have been tested in previous phases of this selection process, and Enka Textile Rayon has emerged as a possible fabric, though its availability is still in question.

Wilson, Kenneth P.↗
NASA NTRS1998

A Fast Code for Jupiter Atmospheric Entry

A fast code was developed to calculate the forebody heating environment and heat shielding that is required for Jupiter atmospheric entry probes. A carbon phenolic heat shield material was assumed and, since computational efficiency was a major goal, analytic expressions were used, primarily, to calculate the heating, ablation and the required insulation. The code was verified by comparison with flight measurements from the Galileo probe's entry; the calculation required 3.5 sec of CPU time on a work station. The computed surface recessions from ablation were compared with the flight values at six body stations. The average, absolute, predicted difference in the recession was 12.5% too high. The forebody's mass loss was overpredicted by 5.5% and the heat shield mass was calculated to be 15% less than the probe's actual heat shield. However, the calculated heat shield mass did not include contingencies for the various uncertainties that must be considered in the design of probes. Therefore, the agreement with the Galileo probe's values was considered satisfactory, especially in view of the code's fast running time and the methods' approximations.

Tauber, Michael E.↗