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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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NASA NTRS2004

Studies on Hot-Melt Prepregging on PRM-II-50 Polyimide Resin with Graphite Fibers

A second generation PMR (in situ Polymerization of Monomer Reactants) polyimide resin PMR-II-50, has been considered for high temperature and high stiffness space propulsion composites applications for its improved high temperature performance. As part of composite processing optimization, two commercial prepregging methods: solution vs. hot-melt processes were investigated with M40J fabrics from Toray. In a previous study a systematic chemical, physical, thermal and mechanical characterization of these composites indicated the poor resin-fiber interfacial wetting, especially for the hot-melt process, resulted in poor composite quality. In order to improve the interfacial wetting, optimization of the resin viscosity and process variables were attempted in a commercial hot-melt prepregging line. In addition to presenting the results from the prepreg quality optimization trials, the combined effects of the prepregging method and two different composite cure methods, i.e. hot press vs. autoclave on composite quality and properties are discussed.

Shin, E. Eugene↗

NASA NTRS1989

Smart structures research program at Virginia Tech

A review of the smart structures and avionics research and teaching program that started in 1979 at Virginia Tech is described. Current smart structures research include major efforts in the development of embedded and attached optical fiber and acoustic fiber sensors for cure monitoring, in-service lifetime structural testing, nondestructive evaluation, and impact and damage detection and analysis; of gradual material degradation; sensor signal multiplexing, processing and data handling to achieve near real-time distributed structural analysis; and the integration of embedded sensors, actuators and control electronics to achieve controlled structural response. Special campus facilities used for this work include an optical fiber fabrication facility, an autoclave for composite structure fabrication and curing, and laboratories for optical fiber sensor development, materials response and nondestructive evaluation, structural control testing and computer engineering.

Claus, R. O.↗

NASA NTRS1989

Manufacture of composite test specimens for delamination studies

This paper describes the process for manufacturing high-quality test specimens for uses in evaluations of interlaminar tensile strength of laminated composites. The chosen specimen configuration is a curved beam which experiences interlaminar tension in the region of greatest curvature when the beam is subjected to 'opening' forces. The manufacturing process uses a lock-mold tooling approach, the principle of which relies upon the difference in coefficients of thermal expansion between the internal rubber mandrel and the surrounding steel female mold. With this method, compaction pressures above those provided by a typical autoclave can be achieved.

Sumich, M.↗

NASA NTRS1980

Graphite polyimide fabrication research for supersonic cruise aircraft

Advanced fabrication processes and adhesive bonding methods have been developed for the fabrication of full scale fiberglass-polyimide honeycomb stiffened HTS-1 graphite/PMR-15 polyimide panels that meet the design criteria for an upper wing panel of the NASA YF-12 aircraft. Detailed manufacturing, bonding, and autoclave cure procedures are presented. Nondestructive test methods including pulse echo and through-transmission ultrasonic C-scan and laser holography were developed to detect flaws in components and the completed panels. Panels were tested in shear at room temperature and 533 K (500 F) following a variety of thermal exposures to obtain load, deflection and failure analysis data.

Freeman, W. T., Jr.↗

NASA NTRS2001

X-33 Tank Failure During Autoclave Fabrication

During a repair cure cycle on tank #1 of the X-33 liquid hydrogen tanks, a skin to core disbond occurred. Both the inner skin and outer skin of the lobe #1 sandwich panel was noted to have been disbonded and cracked- An investigation was undertaken to determine the cause of this failure. The investigation consisted of reviewing all of the processing data and performing testing on the failed lobe #1, as well as the other lobes, which did not fail during the cure cycle. The tests consisted of residual stress measurements in one of the intact lobes and "plug-pulls" to assess skin to core strength on all of the remaining lobes. Results showed an extremely low bondline strength due to lack of proper filleting of the adhesive, in addition, tests showed a very rapid decrease in strength with increasing temperature, as well as a further decrease in strength with a larger number of cycles. Also, the honeycomb used was not vented so pressure could build up within the cells. All of these factors appeared to be contributors to the failure.

Nettles, Alan T.↗

NASA NTRS1975

PMR polyimides: Processable high temperature composite matrix resins

Processing reproducibility and versatility were demonstrated for producing addition-cured polyimide/graphite fiber composites using an in situ polymerization of monomeric reactants directly on the fiber surface. The polymers so derived, designated PMR polyimides, can be fabricated into composite structures by laminating, random fiber molding or autoclave curing. Composites were determined to be thermally stable and retain useful properties after extended exposures at 550 F to 650 F. The material and fabrication capability were demonstrated by the fabrication and evaluation of prototype complex fan blades.

Winters, W. E.↗

NASA NTRS1975

PMR polyimides - Processable high temperature composite matrix resins

Processing reproducibility and versatility were demonstrated for producing addition-cured polyimide/graphite fiber composites using a unique in situ polymerization of monomeric reactants directly on the fiber surface. The polymers so derived, designated PMR polyimides, can be fabricated into composite structures by laminating, random fiber molding or autoclave curing. Composites were determined to be thermally stable and retain useful properties after extended exposures at 550 to 650 F. The material and fabrication capability were demonstrated by the fabrication and evaluation of prototype complex fan blades.-

Winters, W. E.↗

NASA NTRS1998

Automated Fabrication Technologies for High Performance Polymer Composites

New fabrication technologies are being exploited for building high graphite-fiber-reinforced composite structure. Stitched fiber preforms and resin film infusion have been successfully demonstrated for large, composite wing structures. Other automatic processes being developed include automated placement of tacky, drapable epoxy towpreg, automated heated head placement of consolidated ribbon/tape, and vacuum-assisted resin transfer molding. These methods have the potential to yield low cost high performance structures by fabricating composite structures to net shape out-of-autoclave.

Shuart , M. J.↗

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 NTRS2025

Improving Adhesive Bondline Time of Flight Predictions During Autoclave Cure Utilizing Machine Learning

Composite materials are increasingly being used in aerospace applications due to their superior strength-to-weight ratio compared to commonly used metals. A current limitation to widespread adoption is the certification of adhesively bonded joints. One approach to improving adhesive bonding in composites is accurately measuring the thickness of adhesive bondlines in composite laminates. Precise bondline thickness control is essential for aerospace applications where adhesive layer thickness directly affects joint fracture properties and structural performance. This study focused on implementing machine learning techniques to determine the ultrasonic time of flight (directly correlated to thickness) in adhesive bondlines throughout autoclave cure cycles. A high-temperature (use up to 180°C) ultrasonic scanning system was deployed in an autoclave to provide time of flight data through composite panels. Three experiments were conducted on the curing of 305 mm × 305 mm unidirectional composite panels. In the first experiment, a piecewise function was fit for the temperature correction factor to account for changing autoclave temperatures. Due to deficiencies in the first calibration experiment, a second experiment was run, and the results were used to train a machine learning model. The revised experiment, in combination with the machine learning model, significantly increased the accuracy of the bondline time of flight predictions (~14% error reduced to <1%). Data was processed using the Regression Learner Application in MATLAB®, with a Support Vector Machine selected for the model. The result was a machine learning algorithm capable of reliably quantifying ultrasonic time of flight through adhesive bondlines. The third experiment provided independent test data for the machine learning model, demonstrating that the model produces accurate predictions from data beyond its training set.

Machine Learning↗

NASA NTRS1981

The challenges of manufacturing graphite-epoxy structural columns for space platforms

A description is given of the manufacturing processes developed for large-quantity production of space structure graphite/epoxy composite tubes. The tubes are to be delivered to orbit by the Space Shuttle and assembled by astronaut-assisted assembly machines. The tooling and manufacturing system uses a novel method of dry fiber placement which carefully controls fiber position and orientation and a non-autoclave cure for the epoxy resin. The means of dimensional tolerance control over differential thermal expansion between metal tooling and graphite tubes, resin shrinkage during cure, hot pressurized resins in close tolerance tooling, and tool wear, are examined in detail. A successful pilot production program is also reported.

Vaughn, R. L.↗

NASA NTRS1979

Development and fabrication of bismaleimide-graphite composites

The successful fabrication of high temperature resistant composites depends mainly on the processability of the resin binder matrix. For two new bismaleimide type resins the processing of graphite fabric prepregs to composites is described. One resin coded M 751 has to be processed from N-Methylpyrrolidone, the other resin evaluated is a so-called hot melt solvent-less system. Commercial T300/3000 Graphite fabrics were used as reinforcement. The M 751 - Resin is a press grade material and laminates are therefore moulded in high pressure conditions (400 N/sq cm). The solvent-less resin system H 795 is an autoclave grade material and can be cured at 40 N/sq cm. The cure cycles for both the press grade and the autoclave grade material (Fiberite W 143 fabric prepregs) are provided and the mechanical properties of laminates at low (23 C) and high (232 C) temperatures were measured. For comparison, the neat resin flexural properties are also presented. The water absorption for the neat resins and the graphite fabric laminates after a 1000 hour period was evaluated.

Stenzenberger, H.↗

NASA NTRS1994

Resin impregnation process for producing a resin-fiber composite

Process for vacuum impregnation of a dry fiber reinforcement with a curable resin to produce a resin-fiber composite, by drawing a vacuum to permit flow of curable liquid resin into and through a fiber reinforcement to impregnate same and curing the resin-impregnated fiber reinforcement at a sufficient temperature and pressure to effect final curing. Both vacuum and positive pressure, e.g. autoclave pressure, are applied to the dry fiber reinforcement prior to application of heat and prior to any resin flow to compact the dry fiber reinforcement, and produce a resin-fiber composite of reduced weight, thickness and resin content, and improved mechanical properties. Preferably both a vacuum and positive pressure, e.g. autoclave pressure, are also applied during final curing.

Palmer, Raymond J.↗

NASA NTRS1986

Dynamic dielectric analysis for nondestructive cure monitoring and process control

Dynamic dielectric analysis (DDA) is an effective in situ NDE method that can monitor the reaction status in thermosets and the phase changes in thermoplastics, including slow reactions occuring late in the cure cycle and recrystallization during annealing. The effects of moisture and resin history on reaction rate can also be determined, as can ionic and dipolar contributions. The ionic mobility parameter is noted to be an excellent monitor of viscosity above the glass transition temperature. The ability of DDA to monitor cure rate variations in a thick section during autoclaving has been demonstrated.

Kranbuehl, D. E.↗

NASA NTRSDate not supplied

In-situ Inspection of Reflowable-Interface Composite Joints During Cure in an Autoclave

In structural bonds, the interface between adherend and adhesive is nearly two-dimensional making it susceptible to minute quantities of contamination, which can cause weak bonds. Regulatory organizations such as the Federal Aviation Administration (FAA) often require redundant load paths in secondary-bonded, primary-structures to alleviate concerns with bonded performance. Under NASA’s Convergent Aeronautics Solutions (CAS) project, the Adhesive Free Bonding of Composites (AERoBOND) project is investigating reformulated aerospace epoxy-matrix resins to enable reflow and diffusion of the resin at the joint interface during a secondary bonding and cure process. The reflow and intermixing of the matrix resin during assembly can eliminate the material discontinuity at the interface, thereby removing the dependence of bond performance on adhesion at a nearly two-dimensional boundary. AERoBOND process development has evaluated a multitude of parameters including the materials used, the stoichiometric offset of the epoxy resins, time and temperature of cure, and thickness of each ply. Without in-situ process monitoring, the condition of a test article is unknown until mechanical testing is completed. This presentation describes the use of an in-situ ultrasonic inspection system to monitor the joining of two composite parts assembled using the AERoBOND technique. This work quantifies the elimination of the interface by measuring the wave reflection, or lack thereof, at the joint throughout the cure cycle. In addition, the results indicate when reflow and cure of the epoxy resin at the joint occurred. By using the recently developed in-situ inspection with a mobile ultrasonic transducer, localized results can be obtained across a large portion of the joint with high resolution.

Tyler B Hudson↗

NASA NTRSDate not supplied

In-situ Inspection of Reflowable-Interface Composite Joints During Cure in an Autoclave

Tyler B Hudson↗

NASA NTRSDate not supplied

In-situ Inspection of Reflowable-Interface Composite Joints During Cure in an Autoclave

In structural bonds, the interface between adherend and adhesive is nearly two-dimensional making it susceptible to minute quantities of contamination, which can cause weak bonds. Regulatory organizations such as the Federal Aviation Administration (FAA) often require redundant load paths in secondary-bonded primary-structures to alleviate the inability to certify bonded performance. To address this issue, the NASA Convergent Aeronautics Solutions (CAS): Adhesive Free Bonding of Composites (AERoBOND) project is investigating reformulated aerospace epoxy-matrix resins to enable reflow and diffusion of the resin at the joint interface during a secondary bonding and cure process. The reflow and intermixing of the matrix resin during assembly can eliminate the material discontinuity at the interface, thereby removing the dependence of bond performance on adhesion at a nearly two-dimensional boundary. AERoBOND process development has evaluated a multitude of parameters including the materials used, the stoichiometric offset of the epoxy resins, time and temperature of cure, and thickness of each ply. Without in-situ process monitoring, the condition of a test article is unknown until mechanical testing is completed. This paper describes the use of an in-situ ultrasonic inspection system to monitor the joining of two composite parts assembled using the AERoBOND technique. This work quantifies the interface by measuring the wave reflection, or lack thereof, at the joint throughout the cure cycle. In addition, the results indicate when reflow and cure of the epoxy resin at the joint occurred. By using the recently developed in-situ inspection method with a mobile ultrasonic transducer, localized results can be obtained across a large portion of the joint with high resolution.

Tyler B Hudson↗

NASA NTRS2024

Offset-Stoichiometric Reflowable Composite Bonding Method with Adhesive for Mitigating Strict Faying Surface Tolerances

Inherent susceptibility of adhesive bonds to miniscule quantities of contamination can cause undetectable weakened bonds. For this reason, the Federal Aviation Administration (FAA) places strict regulations on adhesively bonded joints in primary aircraft structures. To meet certification requirements aircraft manufactures resort to redundant load paths in the form of fasteners which inherently add weight to the structure and increase manufacturing time. In prior work, a secondary bonding technique called AERoBOND was developed, which utilized off-stoichiometric epoxymatrix resins to facilitate reflow and diffusion of the resin within the joint interface during a secondary bonding/cure process, thus achieving a bond similar to a co-cured joint. However, the AERoBOND process required tight spatial tolerances between the two parts being joined. This study examined the utilization of conventional adhesive with the AERoBOND method to act as a filler in the joint line, effectively reducing the need for tight tolerances on the joining parts and serving as a flexible alternative for existing manufacturing processes. Ultrasonic inspection, optical microscopy, and ASTM International standard tests were performed to analyze the joints for defects and to quantify the mode-I and mode-II interlaminar fracture toughness and short beam strength of the proposed methodology with varying manufacturing parameters. The comprehensive results indicate that the AERoBOND+ method with and without surface preparation performs comparably to co-cured and conventional, adhesively bonded joints when secondary cured in an autoclave with 791 kPa of pressure. As an example, the AERoBOND+ panel without surface preparation bonded with 791 kPa of pressure (referred to as AB+3 throughout paper) had a mode-I fracture toughness (G Ic ) of 0.643 kJ/m 2 and a mode-II fracture toughness (G IIc ) of 4.000 kJ/m 2 in the non-precracked condition and 4.218 kJ/m 2 in the precracked condition at the adhesive-to-prepreg interface. These results were 96%, 142%, and 217%, respectively, of a co-cured baseline panel (referred to as C1 throughout paper).

Composites↗