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All sourcesNASA NTRSDOE OSTI
NASA NTRS2006

Nonlinear Spring Finite Elements for Predicting Mode I-Dominated Delamination Growth in Laminated Structure with Through-Thickness reinforcement

One particular concern of polymer matrix composite laminates is the relatively low resistance to delamination cracking, in particular when the dominant type of failure is mode I opening. One method proposed for alleviating this problem involves the insertion pultruded carbon pins through the laminate thickness. The pins, known as z-pins, are inserted into the prepreg laminate using an ultrasonic hammer prior to the curing process, resulting in a field of pins embedded normal to the laminate plane as illustrated in Figure. 1. Pin diameters range between 0.28-mm to 0.5-mm and standard areal densities range from 0.5% to 4%. The z-pins are provided by the manufacturer, Aztex(Registered TradeMark) , in a low-density foam preform, which acts to stabilize orientation of the pins during the insertion process [1-3]. Typical pin materials include boron and carbon fibers embedded in a polymer matrix. A number of methods have been developed for predicting delamination growth in laminates reinforced with z-pins. During a study on the effect of z-pin reinforcement on mode I delamination resistance, finite element analyses of z-pin reinforced double cantilever beam (DCB) specimens were performed by Cartie and Partridge [4]. The z-pin bridging stresses were modeled by applying equivalent forces at the pin locations. Single z-pin pull-out tests were performed to characterize the traction law of the pins under mode I loading conditions. Analytical solutions for delamination growth in z-pin reinforced DCB specimens were independently derived by Robinson and Das [5] and Ratcliffe and O'Brien [6]. In the former case, pin bridging stresses were modeled using a distributed load and in the latter example the bridging stresses were discretely modeled by way of grounded springs. Additionally, Robinson and Das developed a data reduction strategy for calculating mode I fracture toughness, G(sub Ic), from a z-pin reinforced DCB specimen test [5]. In both cases a traction law similar to that adopted by Cartie and Partridge was used to represent z-pin failure under mode I loading conditions. In the current work spring elements available in most commercial finite element codes were used to model z-pins. The traction law used in previous analyses [4-6] was employed to represent z-pin damage. This method is intended for and is limited to simulating z-pins in composite laminate structure containing mode I-dominated delamination cracking. The current technique differs from previous analyses in that spring finite elements (available in commercial codes) are employed for simulating zpins, reducing the complexity of the analysis construction process. Furthermore, the analysis method can be applied to general structure that experiences mode I-dominated delamination cracking, in contrast to existing analytical solutions that are only applicable to coupon DCB specimens.

Ratcliffe, James G.↗
NASA NTRSDate not supplied

Consolidation of Braided Carbon Fiber/ Thermoplastic Tubes by Induction Heating

Thermoplastic composites have experienced renewed enthusiasm within the U.S. aerospace industry, due in part to opportunities for rapidly manufacturing large structures. Unlike the lengthy cure cycles required by thermosets, the thermoplastic matrix prepreg requires only heat and pressure to consolidate consecutive layers of material; with sufficient inter-ply contact time to develop a strong interface. The manufacturing requirements for thermoplastics provide manufacturing flexibility and of particular interest are in-situ consolidation processes; i.e consolidation during layup on a tool. In-situ consolidation removes the time-consuming and costly secondary processing step of autoclave or press consolidation. Process variables which contribute to laminate quality include temperature, heat transfer, inter-ply contact time, and consolidation pressure. Most of the work to date related to in-situ consolidation has been in conjunction with automated fiber placement (AFP), where prepreg tows are consolidated when placed on the tool. Considering this process for a braided structure adds complexity relative to AFP due to the rapid lay-down rate of the braider and simultaneous placement of a multiple prepreg tows. Processes considered for consolidation during AFP could not be applied to a braided part. Alternatively, a fully braided preform could be consolidated on the tool using well established thermoplastic welding techniques. The goal of this effort was to evaluate the effectiveness of induction welding to consolidate braided carbon fiber/thermoplastic composite tubes up to eight plies in thickness. The manufacturing process, thermal characterization, and initial mechanical property data of the consolidated tubes are reported.

In-situ consolidation, induction heating/welding, ↗
NASA NTRS1998

Imide Oligomers Containing Pendent and Terminal Phenylethynyl Groups-2

As part of a program to develop high-performance/high-temperature structural resins for aeronautical applications, imide oligomers containing pendent and terminal phenylethynyl groups were prepared, characterized and the cured resins evaluated as composite matrices. The oligomers were prepared at a calculated number-average molecular weight of 5000 g/mol and contained 15-20 mol% pendent phenylethynyl groups. In previous work, an oligomer containing pendent and terminal phenylethynyl groups exhibited a high glass transition temperature (approximately 313 C), and laminates therefrom exhibited high compressive properties, but processability, fracture toughness, microcrack resistance and damage tolerance were less than desired. In an attempt to improve these deficiencies, modifications in the oligomeric backbone involving the incorporation of 1,3-bis(3-aminophenoxy)benzene were investigated as a means of improving processability and toughness without detracting from the high glass transition temperature and high compressive properties. The amide acid oligomeric solutions were prepared in N-methyl-2-pyrrolidinone and were subsequently processed into imide powder, thin films, adhesive tape and carbon fiber prepreg. Neat resin plaques were fabricated from imide powder by compression moulding. The maximum processing pressure was 1.4 MPa and the cure temperature ranged from 350 to 371 C for 1 h for the mouldings, adhesives, films and composites. The properties of the 1,3-bis(3-aniinophenoxy)benzene modified cured imide oligomers containing pendent and terminal phenylethynyl groups are compared with those of previously prepared oligomers containing pendent and terminal phenylethynyl groups of similar composition and molecular weight.

Connell, J. W.↗
NASA NTRS1979

Low density bismaleimide-carbon microballoon composites

A process is described for the preparation of composite laminate.structures of glass cloth preimpregnated with polybismaleimide resin and adhered to a polybismaleimide glass or aromatic polyamide paper honeycomb cell structure that is filled or partially filled with a syntactic foam consisting of a mixture of bismaleimide resin and carbon microballoons. The carbon microballoons are prepared by pyrolyzing phenolic microballoons and subsequently bonded using a 2% bismaleimide solution. The laminate structures are cured for two hours at 477 deg K and are adhered to the honeycomb bismaleimide adhesive using a pressure of 700 KN/sq m pressure at 450 deg K. The laminate composite is then post-cured for two hours at 527 deg K to produce a composite laminate having a density in the range from about 95 kilograms per cubic meter to 130 kilograms per cubic meter.

Kourtides, D. A.↗
NASA NTRS2000

Process of making carbon-carbon composites

A carbon composite structure, for example, an automotive engine piston, is made by preparing a matrix including of a mixture of non crystalline carbon particulate soluble in an organic solvent and a binder that has a liquid phase. The non crystalline particulate also contains residual carbon hydrogen bonding. An uncured structure is formed by combining the matrix mixture, for example, carbon fibers such as graphite dispersed in the mixture and/or graphite cloth imbedded in the mixture. The uncured structure is cured by pyrolyzing it in an inert atmosphere such as argon. Advantageously, the graphite reinforcement material is whiskered prior to combining it with the matrix mixture by a novel method involving passing a gaseous metal suboxide over the graphite surface.

Withers, James C.↗
NASA NTRS1980

Process for the manufacture of low density bis-maleimide-carbon microballoon composites

A process for the preparation of composite laminate structures of glass cloth preimpregnated with polybismaleimide resin and adhered to a polybismaleimide-glass or aromatic polyamide paper honeycomb cell structure filled or partially filled with a syntactic foam consisting of a mixture of bismaleimide resin and carbon microballoons. The carbon microballoons are prepared by pyrolyzing phenolic microballoons and subsequently bonded using a 2% bismaleimide solution. The laminate structures are cured for two hours at 477.degree. K. and are adhered to the honeycomb bismaleimide adhesive using a pressure of 700 kN/m.sup.2 pressure at 450.degree. K. The laminate composite is then post-cured for two hours at 527.degree. K. to produce a composite laminate having a density in the range from about 95 kilograms per cubic meter to 130 kilograms per cubic meter.

Kourtides, Demetrius A.↗
NASA NTRS2014

PRSEUS Structural Concept Development

A lighter, more robust airframe is one of the key technological advancements necessary for the successful launch of any large next-generation transport aircraft. Such a premise dictates that considerable improvements beyond current state-of-the-art aluminum structures is needed, and that improvements of this magnitude will require an extensive use of composite materials that are not only lightweight, but also economical to produce. To address this challenge, researchers at NASA and The Boeing Company are developing a novel structural concept called the Pultruded Rod Stitched Efficient Unitized Structure (PRSEUS) under the Environmentally Responsible Aviation (ERA) Project. It is an integrally stiffened panel concept that is stitched together and designed to maintain residual load-carrying capabilities under a variety of damage scenarios. In addition to improved structural performance, an important facet of this unique arrangement of stitched carbon fibers is its innovative manufacturing method that has the potential to lower fabrication costs by eliminating fasteners and autoclave cures. The rationale and development status for this new approach forms the basis of the work described in this paper. The test specimens described herein were fabricated, or are currently being fabricated, by The Boeing Company, while the structural analyses and testing tasks are being performed by NASA and Boeing personnel.

Velicki, Alex↗
NASA NTRS2013

RTM370 Polyimide Braided Composites: Characterization and Impact Testing

RTM370 imide oligomer based on 2,3,3',4'-biphenyl dianhydride (a-BPDA), 3,4'-oxydianiline (3,4'-ODA) and terminated with the 4-phenylethynylphthalic (PEPA) endcap has been shown to exhibit a low melt viscosity (10-30 poise) at 280 C with a pot-life of 1-2 h and a high cured glass transition temperature (Tg) of 370 C. RTM370 resin has been successfully fabricated into composites reinforced with T650-35 carbon fabrics by resin transfer molding (RTM). RTM370 composites display excellent mechanical properties up to 327 C (620 F), and outstanding property retention after aging at 288degC (550 F) for 1000 h, and under hot-wet conditions. In ballistic impact testing, RTM370 triaxial braided T650-35 carbon fiber composites exhibited enhanced energy absorption at 288 C (550 F) compared to ambient temperature.

Chuang, Kathy C.↗
NASA NTRS1992

Damage-Tolerant Composites Made By Stitching Carbon Fabrics

Work conducted at NASA Langley Research Center to investigate stitching combined with resin transfer molding to make composites more tolerant of damage and potentially cost competitive with metals. Composite materials tailored for damage tolerance by stitching layers of dry carbon fabric with closely spaced threads to provide reinforcement through thickness. Epoxy resin then infused into stitched preforms, and epoxy was cured. Various stitching patterns and thread materials evaluated by use of flat plate specimens. Also, blade-stiffened structural elements fabricated and tested. Stitched flat laminates showed outstanding damage tolerance, excellent compression strength in notched specimens, and acceptable fatigue behavior. Development of particular interest to aircraft and automotive industries.

Dow, Marvin B.↗
NASA NTRS2007

Fabrication of Polyimide-Matrix/Carbon and Boron-Fiber Tape

The term HYCARB denotes a hybrid composite of polyimide matrices reinforced with carbon and boron fibers. HYCARB and an improved process for fabricating dry HYCARB tapes have been invented in a continuing effort to develop lightweight, strong composite materials for aerospace vehicles. Like other composite tapes in this line of development, HYCARB tapes are intended to be used to build up laminated structures having possibly complex shapes by means of automated tow placement (ATP) - a process in which a computer-controlled multiaxis machine lays down prepreg tape or tows. The special significance of the present process for making dry HYCARB for ATP is that it contributes to the reduction of the overall cost of manufacturing boron-reinforced composite-material structures while making it possible to realize increased compression strengths. The present process for making HYCARB tapes incorporates a "wet to dry" process developed previously at Langley Research Center. In the "wet to dry" process, a flattened bundle of carbon fiber tows, pulled along a continuous production line between pairs of rollers, is impregnated with a solution of a poly(amide acid) in N-methyl-2-pyrrolidinone (NMP), then most of the NMP is removed by evaporation in hot air. In the present case, the polyamide acid is, more specifically, that of LaRC. IAX (or equivalent) thermoplastic polyimide, and the fibers are, more specifically, Manganite IM7 (or equivalent) polyacrylonitrile- based carbon filaments that have a diameter of 5.2 m and are supplied in 12,000-filament tows. The present process stands in contrast to a prior process in which HYCARB tape was made by pressing boron fibers into the face of a wet carbon-fiber/ poly(amide acid) prepreg tape . that is, a prepreg tape from which the NMP solvent had not been removed. In the present process, one or more layer(s) of side-by-side boron fibers are pressed between dry prepreg tapes that have been prepared by the aforementioned gwet to dry h process. The multilayer tape is then heated to imidize the matrix material and remove most of the remaining solvent, and is pressed to consolidate the multiple layers into a dense tape. For tests, specimens of HYCARB tapes and laminated composite panels made from HYCARB tape were prepared as follows: HYCARB tapes were fabricated as described above. Each panel was made by laying down ten layers of tape, containing, variously, one, two, or three boron-fiber plies and the remainder carbon- fiber-only plies (see figure). Each panel was made by laying down ten layers of tape. Each panel was then cured by heating to a temperature of 225 C for 15 minutes, then pressing at 200 psi (A1.4 MPa) while heating to 371 C, holding at 371 C for 1 hour, then continuing to hold pressure during cooling. Control specimens that were otherwise identical except that they did not contain boron fibers also were prepared. In room-temperature flexural tests, the HYCARB specimens performed comparably to the control specimens; in room-temperature, open-hole compression tests, the HYCARB specimens performed slightly better, by amounts that increased with boron content.

Belvin, Harry L.↗
NASA NTRSDate not supplied

Process Development and Testing for Lunar Habitats

Non-metallic composites are appealing for Lunar habitats due to their low specific weight and contribution to passive radiation shielding. This application of composite materials presents notable challenges, however: flammability risks are higher due to elevated oxygen levels expected in habitat structures and absorbed radiation can cause damage long-term. The Materials and Processes Laboratory at Marshall Space Flight Center awarded funding for an internal research and development study which included fabrication and testing of composite materials for Lunar habitats. Composite layups used carbon fiber or ultra-high molecular weight polyethylene reinforcement and a variety of polymer matrices. Pre-impregnated fabric was autoclave cured; however, novel coatings were incorporated during select layups. Samples from each panel were used for screening-level flammability testing (ASTM D 2863) and combined vacuum ultraviolet and electron irradiation followed by tension testing (ASTM D5766). Two of the tested composite materials exceeded a 37% limiting oxygen index. Material property degradation for a 10 year simulated radiation dose was minimal in most cases.

manufacturing↗
NASA NTRS2020

AERoBOND Project Summary

Under NASA’s Convergent Aeronautics Solutions (CAS) project, the Adhesive-Free Bonding of Complex Composites (AERoBOND) project investigated off-stoichiometric epoxy polymers for fast, reliable assembly of epoxy matrix composite structures. The project goal was to demonstrate feasibility of the AERoBOND joining method by demonstrating mechanical properties greater than 80% of conventional co-cured materials while reducing structure weight by 1%. The project consisted of three convergent research areas: material and process development, systems analysis, and material and process modeling. Material and process development was the largest component of AERoBOND with approximately 6 FTE and 1WYE of support to formulate and characterize new resins, prepare carbon fiber prepregs, fabricate laminates, measure mechanical properties, analyze failure results, and select material and process improvements. The systems analysis activity estimated the potential reduction in part count and aircraft weight by comparing models of composite wing boxes with no fasteners (co-cured structure), fasteners in major joints (co-cured stringers), and fasteners in all joints. The materials and process modeling activity included a molecular model of the AERoBOND materials system to predict mechanical properties of resins with offset stoichiometry and a process model to predict the effect of resin formulation and processing conditions on the extent of mixing and degree of cure in a finished joint. As the number of airline passenger trips doubles in the next 20 years (IATA/Tourism Economics Air Passenger Forecasts, April 2019), the increased demand for new commercial aircraft is now the single greatest technical challenge to the airframe manufacturing industry. To meet efficiency requirements, new aircraft must be fabricated primarily from high performance structural composites, but manufacturing processes are inherently slow with the largest bottleneck attributed to assembly and installation of fasteners (NASA/TM–2019-220428). Manufactures of commercial transport aircraft are compelled to install more than 100,000 redundant fasteners into bonded joints to prevent failures due to unpredictable weak bonds. In structural adhesive bonds, the interface between adherend and adhesive is nearly two-dimensional making it susceptible to minute quantities of contamination, which can cause weak bonds. Currently, bond strength assessment is only possible through destructive testing (i.e., breaking the joint). For these reasons, regulatory organizations such as the Federal Aviation Administration (FAA) often require redundant load paths in secondary-bonded, primary-structures to alleviate concerns with bond performance. The AERoBOND process enables reflow of matrix resin during assembly to eliminate the material discontinuity at the interface, thereby eliminating the dependence of mechanical performance on interfacial adhesion. The AERoBOND joint is equivalent to the interlaminar region obtained during a co-cure process, so joint performance depends on the cohesive properties of the matrix resin. Conventional co-cured structures, although too costly and complex for large-scale manufacturing, are trusted by manufacturers and regulators, and are certified for flight with few or no redundant fasteners.Systems analysis performed on a composite wing model at the scale of a single-aisle commercial transport aircraft indicated that >20,000 redundant fasteners per wing could be eliminated by implementing the AERoBOND joining method. A total weight reduction of 15% was predicted in a wing box by eliminating fasteners and thinning components that must no longer support localized fastener loads and accommodate fastener dimensions. Interlaminar shear fracture toughness measured by the end-notched flexure test was greater than 1 kJ/m2 (nearly 140% of the co-cured benchmark property), which is greatly in excess of the project goals for mechanical properties. Testing was planned to measure interlaminar tensile fracture toughness as well as interlaminar tensile and shear strengths using the same AERoBOND configuration, but was delayed due to closure of LaRC facilities during the COVID-19 pandemic. The AERoBOND process model is partially validated and available for experimental use. It allows the user to input AERoBOND process parameters such as material composition, laminate configuration, and cure cycle to predict the final cure state of the AERoBOND joint. A preliminary, multi-scale material model was developed to predict AERoBOND joint mechanical properties (stiffness and strength) based on the cure state of the joint provided by the process model. The timing for transition of this technology within NASA is excellent as NASA initiates new enduring projects to address composites manufacturing rate challenges. AERoBOND technology is well suited to AAVP/AATT objectives for rapid manufacturing of a composite wing. A minimal effort (1 FTE/$15k procurement/0 WYE) is proposed in FY21 to continue a minor mechanical testing effort and maintain a SAA with ASX composites to develop commercial quality prepreg material. An RFI with the composites industry is suggested to quantify the technology gap between the current TRL and the TRL needed for transition to industry. A moderate effort [3-4 FTE/$150k/1 WYE (~$115k)] is proposed in FY22 for the “high rate composites manufacturing” project currently in planning. The partnership with ASX Composites will be expanded to produce material for sub-element/element-scale “panel-off” activities. Industry partnerships with airframe manufacturers is an expected component to explore damage tolerance and environmental stability. Further development of multi-scale modeling tools (process model, meso-scale model, and molecular model) is planned to enhance and deliver tools for rapid manufacturing infusion.

Frank Louis Palmieri↗
NASA NTRS1992

An infiltration/cure model for manufacture of fabric composites by the resin infusion process

A 1-D infiltration/cure model was developed to simulate fabrication of advanced textile composites by the resin film infusion process. The simulation model relates the applied temperature and pressure processing cycles, along with the experimentally measured compaction and permeability characteristics of the fabric preforms, to the temperature distribution, the resin degree of cure and viscosity, and the infiltration flow front position as a function of time. The model also predicts the final panel thickness, fiber volume fraction, and resin mass for full saturation as a function of compaction pressure. Composite panels were fabricated using the RTM (Resin Transfer Molding) film infusion technique from knitted, knitted/stitched, and 2-D woven carbon preforms and Hercules 3501-6 resin. Fabric composites were fabricated at different compaction pressures and temperature cycles to determine the effects of the processing on the properties. The composites were C-scanned and micrographed to determine the quality of each panel. Advanced cure cycles, developed from the RTM simulation model, were used to reduce the total cure cycle times by a factor of 3 and the total infiltration times by a factor of 2.

Weideman, Mark H.↗
NASA NTRS2011

Infusion Processing of Phenylethynyl Terminated Imides by High Temperature RTM and VARTM

Fabrication of composite structures using infusion processes such as resin transfer molding (RTM) and vacuum assisted resin transfer molding (VARTM) is generally more affordable than conventional autoclave techniques. Recent efforts have focused on adapting both technologies for the fabrication of high temperature (HT) resistant composites. Due to their low melt viscosity and long melt stability, certain phenylethynyl terminated imides (PETI) can be processed into composites using these high temperature out-of-autoclave processes. In the current study, two PETI resins, LARC(TradeMark) PETI-330 and LARC(TradeMark) PETI-8 have been used to make test specimens using both RTM and VARTM. For aerospace applications, a void fraction of less than 2% is desired. Traditionally, RTM has had the advantage over VARTM for generating composites with low void content. However, the process is limited in terms of size. Work at NASA LaRC has incorporated modifications to the thermal cycle used in laminate fabrication that have reduced the void content significantly (typically 1-3%) using the current HT-VARTM process. For composite fabrication by both RTM and VARTM, the resins were infused into three carbon fiber preforms (T650-35-3k 5HS, IM7-6k 5HS, and IM7-6k Uniweave) at 316 C and 260 C respectively and cured up to 371 C. The details of the RTM processing carried out at the University of Akron are discussed in this work along with a brief description of the HT-VARTM processing carried out at NASA-LaRC. Photomicrographs of the panels were taken and void contents were determined by acid digestion. Mechanical properties (short beam shear, SBS) of the panels fabricated by both infusion processes were determined at room temperature as well as at various elevated temperatures. The results of this work are presented herein.

Ghose, Sayata↗
NASA NTRS2000

TPS Materials and Costs for Future Reusable Launch Vehicles

There is considerable interest in developing new reusable launch vehicles (RLVs) for reducing the cost of transporting payload to and from orbit. This work reviews thirteen candidate thermal protection system (TPS) options currently available for RLVs. It is useful to begin with the current Shuttle TPS layout as a reference. The nose cap and wing leading edge , which reach the highest temperatures, are made of reinforced carbon-carbon (RCC) that is protected from oxidation by an external coating (about 0.020" thick) of silicon-carbide. Most of the windward surface is 9 lb/cubic ft ceramic tiles (LI-900) with a thin (about 0.012") coating of Reaction Cured Glass (RCG). The leeward side of the vehicle is covered largely by AFRSI, a quilted ceramic blanket, and FRSI, a polyamide felt. These four materials can be considered first generation reusable TPS. Since the time of the Shuttle design, considerable progress has been made advancing TPS technologies in terms of thermal performance, robustness, and cost. For each of the major systems, a second generation ceramic TPS has been developed, tested, and characterized. Metallic-based systems have also been developed. For applications requiring RCC in the past, advanced carbon-carbon (ACC) is now available. This material has better mechanical properties, somewhat higher temperature capability to 2900F and greatly increased oxidation resistance. New carbon fiber reinforced silicon-carbide matrix composites (C/SiCs) have shown additional improvement in properties over ACC with use temperatures to 3000F and above. For rigid tiles, NASA Ames has made two significant advancements. The first is a tile substrate called Alumina Enhanced Thermal Barrier, or AETB, that incorporates alumina fibers for improved dimensional stability at high temperatures, to 2600F and above. This material can be made to densities as low as 8 lb/cubic ft. The second is a coating preparation called Toughened Uni-piece Fibrous Insulation, or TUFT, that penetrates about 0.1 in. into the tile substrate. The resulting composite, with a functionally gradient density near the surface, provides orders of magnitude increased damage resistance compared with RCG coated LI-900, with only a small weight increase. The TPS that combines these two developments is called AETB-8/TUFI and has been adopted for high damage areas on the Shuttles. Two notable developments have occur-red in flexible ceramic blanket technology. The first is aluminoborosilicate-based fibers with use temperatures of 2200F and above," in comparison to quartz and silica fiber used in AFRSI which have multi-use temperature limits of 1200 to 1400F. Blankets incorporating these new high temperature fibers are referred to as AFRSI-HT. The second is an integral weaving techniques that produces a fluted core blanket with a smoother surface and greater resistance to aero-acoustic noise, to levels as high as 170 dB. This Ames innovation is called Tailorable Advanced Blanket Insulation, or TABI. Finally, for felt-based TPS, Boeing is developing Polybenzimidazole Blanket Insulation, or PBI, with a multi-use temperature limit of 1000F and above, in contrast to Shuttle FRSI which has a multi-use temperature limit of about 700F. 1.6 NASA Langley and BF Goodrich (formerly Rohr Corp.) have led the development of metallic-based TPS. This activity uses essentially three approaches: metallic tiles which encase a fibrous ceramic batting in a box fabricated largely from metallic honeycombs, typically Nickel based alloys; metallic honeycomb sheets, made of Nickel-based alloys, incorporating a fibrous back-side insulation encapsulated in a metallic foil bag, providing reduced weight; and metallic multi-wall, which is comprised of dimpled Titanium metal sheets, which are stacked and then diffusion bonded at contact points to form the TPS. The Nickel-based systems can be used up to temperatures of about 1800F, and the Titanium system to about 100F. These thirteen TPS materials have pros and cons to their usage in terms of temperature capability, weight, initial cost, and maintenance. Carbon-carbon and C/SiC systems have the highest temperature capability but are relatively expensive and heavy, requiring significant time, expertise, and costly facilities and tools for design and fabrication. Second generation ceramic tiles are relatively light, durable, simple to fabricate and easy to install; however, waterproofing is a concern. Blankets and felts are light, simple, inexpensive, and easy to install over curved vehicle surfaces, but durability and waterproofing are concerns. Metallics are robust and appear to have eliminated waterproofing, but they tend to be heavy and relatively expensive, requiring costly facilities and tools. If thin metal sheets are used to reduce weights, then issues arise from possible metal fatigue and corrosion caused by thermal cycling, pressure oscillations, and environmental exposure. For application to future RLVs, system analyses show that a significant component of the vehicle life cycle cost is from the TPS; however, it is difficult to quantify and to compare the potential savings of advanced systems without performing full vehicle designs using each of the different options. Because this entails a considerable effort and also tends to submerge TPS cost impacts under unrelated vehicle design assumptions, there is a clear need for a simpler quantitative method to evaluate the cost impact of different TPS options. To this end, this work introduces a TPS life-cycle cost parameter which is easily computed and applicable to generic RLVs.

Rasky, Dan J.↗
NASA NTRS2007

Sequestration of Single-Walled Carbon Nanotubes in a Polymer

Sequestration of single-walled carbon nanotubes (SWCNs) in a suitably chosen polymer is under investigation as a means of promoting the dissolution of the nanotubes into epoxies. The purpose of this investigation is to make it possible to utilize SWCNs as the reinforcing fibers in strong, lightweight epoxy-matrix/carbon-fiber composite materials. SWCNs are especially attractive for use as reinforcing fibers because of their stiffness and strength-to-weight ratio: Their Young s modulus has been calculated to be 1.2 TPa, their strength has been calculated to be as much as 100 times that of steel, and their mass density is only one-sixth that of steel. Bare SWCNs cannot be incorporated directly into composite materials of the types envisioned because they are not soluble in epoxies. Heretofore, SWCNS have been rendered soluble by chemically attaching various molecular chains to them, but such chemical attachments compromise their structural integrity. In the method now under investigation, carbon nanotubes are sequestered in molecules of poly(m-phenylenevinylene-co-2,5-dioctyloxy-p-phenylenevinylene) [PmPV]. The strength of the carbon nanotubes is preserved because they are not chemically bonded to the PmPV. This method exploits the tendency of PmPV molecules to wrap themselves around carbon nanotubes: the wrapping occurs partly because there exists a favorable interface between the conjugated face of a nanotube and the conjugated backbone of the polymer and partly because of the helical molecular structure of PmPV. The constituents attached to the polymer backbones (the side chains) render the PmPV-wrapped carbon nanotubes PmPV soluble in organic materials that, in turn, could be used to suspend the carbon nanotubes in epoxy precursors. At present, this method is being optimized: The side chains on the currently available form of PmPV are very nonpolar and unable to react with the epoxy resins and/or hardeners; as a consequence, SWCN/PmPV composites have been observed to precipitate out of epoxies while the epoxies were being cured. If the side chains of the PmPV molecules were functionalized to make them capable of reacting with the epoxy matrices, it might be possible to make progress toward practical applications. By bonding the side chains of the PmPV molecules to an epoxy matrix, one would form an PmPV conduit between the epoxy matrix and the carbon nanotubes sequestered in the PmPV. This conduit would transfer stresses from the epoxy matrix to the nanotubes. This proposed load-transfer mode is similar to that of the current practice in which silane groups are chemically bonded to both the epoxy matrices and the fibers (often glass fibers) in epoxymatrix/fiber composites.

Bley, Richard A.↗
NASA NTRS1992

Removal of surface loop from stitched composites can improve compression and compression-after-impact strengths

Stitching through-the-thickness (TTT) of composite materials produces a surface loop of yarn between successive penetrations. The surface loop is pressed into the surface layers of the composite material during the curing of the laminate, kinking the in-plane fibers near the surface of the material. The compression strength and compression-after-impact (CAI) strengths of carbon-epoxy specimens were measured with and without the surface loop. Removal of the surface loop had no influence on failure mode or failure mechanism, but did significantly increase the compression and CAI strengths.

Farley, Gary L.↗
NASA NTRS2002

Structural Health Monitoring of Composite Wound Pressure Vessels

The increasing use of advanced composite materials in the wide range of applications including Space Structures is a great impetus to the development of smart materials. Incorporating these FBG sensors for monitoring the integrity of structures during their life cycle will provide valuable information about viability of the usage of such material. The use of these sensors by surface bonding or embedding in this composite will measure internal strain and temperature, and hence the integrity of the assembled engineering structures. This paper focuses on such a structure, called a composite wound pressure vessel. This vessel was fabricated from the composite material: TRH50 (a Mitsubishi carbon fiber with a 710-ksi tensile strength and a 37 Msi modulus) impregnated with an epoxy resin from NEWPORT composites (WDE-3D-1). This epoxy resin in water dispersed system without any solvents and it cures in the 240-310 degrees F range. This is a toughened resin system specifically designed for pressure applications. These materials are a natural fit for fiber sensors since the polyimide outer buffer coating of fiber can be integrated into the polymer matrix of the composite material with negligible residual stress. The tank was wound with two helical patterns and 4 hoop wraps. The order of winding is: two hoops, two helical and two hoops. The wall thickness of the composite should be about 80 mil or less. The tank should burst near 3,000 psi or less. We can measure the actual wall thickness by ultrasonic or we can burst the tank and measure the pieces. Figure 1 shows a cylinder fabricated out of carbon-epoxy composite material. The strain in different directions is measured with a surface bonded fiber Bragg gratings and with embedded fiber Bragg gratings as the cylinder is pressurized to burst pressures. Figure 2 shows the strain as a function of pressure of carbon-epoxy cylinder as it is pressurized with water. Strain is measured in different directions by multiple gratings oriented in both axial and hoops directions.

Joseph Grant↗