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Structural design synthesis of recoverable launch vehicle structures
Systems requirements for complex spacecraft drive design requirements that lead to structures, components, and/or enclosures of a multi-material and multifunctional design. The varying physical properties of aluminum, tungsten, Invar, or other high-grade aerospace metals when utilized in conjunction with lightweight composites multiply system level solutions. These multi-material designs are largely dependent upon effective joining techAn improved method of joining metal components to matrix/fiber composite material structures has been invented. The method is particularly applicable to equipping such thin-wall polymer-matrix composite (PMC) structures as tanks with flanges, ceramic matrix composite (CMC) liners for high heat engine nozzles, and other metallic-to-composite attachments. The method is oriented toward new architectures and distributing mechanical loads as widely as possible in the vicinities of attachment locations to prevent excessive concentrations of stresses that could give rise to delaminations, debonds, leaks, and other failures. The method in its most basic form can be summarized as follows: A metal component is to be joined to a designated attachment area on a composite-material structure. In preparation for joining, the metal component is fabricated to include multiple studs projecting from the aforementioned face. Also in preparation for joining, holes just wide enough to accept the studs are molded into, drilled, or otherwise formed in the corresponding locations in the designated attachment area of the uncured ("wet') composite structure. The metal component is brought together with the uncured composite structure so that the studs become firmly seated in the holes, thereby causing the composite material to become intertwined with the metal component in the joining area. Alternately, it is proposed to utilize other mechanical attachment schemes whereby the uncured composite and metallic parts are joined with "z-direction" fasteners. The resulting "wet" assembly is then subjected to the composite-curing heat treatment, becoming a unitary structure. It should be noted that this new art will require different techniques for CMC s versus PMC's, but the final architecture and companion curing philosophy is the same. For instance, a chemical vapor infiltration (CVI) fabrication technique may require special integration of the pre-form and
Carbon nanotubes are cylindrical molecules composed of carbon atoms in a regular hexagonal arrangement. If nanotubes can be uniformly dispersed in a supporting matrix to form structural materials, the resulting structures could be significantly lighter and stronger than current aerospace materials. Work is currently being done to develop an electrolyte-based self-assembly process that produces a Carbon Nanotube/Nickel composite material with high specific strength. This process is expected to produce a lightweight metal matrix composite material, which maintains it's thermal and electrical conductivities, and is potentially suitable for applications such as advanced structures, space based optics, and cryogenic tanks.
A new microsensor fabrication approach has been demonstrated based upon the use of nanostructures as templates. The fundamental idea is that existing nanostructures, such as carbon nano tubes or biological structures, have a material structure that can be used advantageously in order to provide new sensor systems but lack the advantages of some materials to, for example, operate at high temperatures. The approach is to start with a template using nanostructures such as a carbon nanotube. This template can then be coated by an oxide material with higher temperature capabilities. Upon heating in air, the carbon nanotube template is burned off, leaving only the metal oxide nanostructure. The resulting structure has a combination of the crystal structure and surface morphology of the carbon nanotube, combined with the material durability and hightemperature- sensing properties of the metal oxide. Further, since the metal oxide nanocrystals are deposited on the carbon nanotube, after burn-off what is left is a metal oxide porous nanostructure. This makes both the interior and the exterior of this nano structured sensor available for gas species detection. This, in effect, increases the surface area available for sensing, which has been shown in the past to significantly increase sensor performance.
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In an improved method of fabricating composite-material structures by laying up prepreg tapes (tapes of fiber reinforcement impregnated by uncured matrix materials) and then curing them, one cures the layups by use of beams of electrons having kinetic energies in the range of 200 to 300 keV. In contrast, in a prior method, one used electron beams characterized by kinetic energies up to 20 MeV. The improved method was first suggested by an Italian group in 1993, but had not been demonstrated until recently. With respect to both the prior method and the present improved method, the impetus for the use of electron- beam curing is a desire to avoid the high costs of autoclaves large enough to effect thermal curing of large composite-material structures. Unfortunately, in the prior method, the advantages of electron-beam curing are offset by the need for special walls and ceilings on curing chambers to shield personnel from x rays generated by impacts of energetic electrons. These shields must be thick [typically 2 to 3 ft (about 0.6 to 0.9 m) if made of concrete] and are therefore expensive. They also make it difficult to bring large structures into and out of the curing chambers. Currently, all major companies that fabricate composite-material spacecraft and aircraft structures form their layups by use of automated tape placement (ATP) machines. In the present improved method, an electron-beam gun is attached to an ATP head and used to irradiate the tape as it is pressed onto the workpiece. The electron kinetic energy between 200 and 300 keV is sufficient for penetration of the ply being laid plus one or two of the plies underneath it. Provided that the electron-beam gun is properly positioned, it is possible to administer the required electron dose and, at the same time, to protect personnel with less shielding than is needed in the prior method. Adequate shielding can be provided by concrete walls 6 ft (approximately equal to 1.8 m) high and 16 in. (approximately equal to 41 cm) thick, without a ceiling. The success of the present method depends on the use of a cationic epoxy as the matrix material in the prepreg tape, heating the prepreg tape to a temperature of 50 C immediately prior to layup, and exposing the workpiece to an electron-beam dose of approximately 2 Mrad. Experiments have shown that structures fabricated by the present method have the same mechanical properties as those of nominally identical structures fabricated by the prior method with electron beams of 3 to 4 MeV.
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The North American Rayon Corporation (NARC) precursor was developed, qualified, and characterized for Space Shuttle nozzle carbon-cloth phenolic ablative materials in three distinct phases. The characterization phase includes thermal and structural material property analysis and comparisons. This report documents the thermal and structural material property characterization performed by Southern Research Institute (SRI) on the two NARC baseline and two crossover materials.
The results of experimental studies of the anisotropy of elastic and strength characteristics of various structural materials, including pressure worked metals and alloys, laminated fiberglass plastics, and laminated wood plastics, are correlated and classified. Strength criteria under simple and complex stresses are considered as applied to anisotropic materials. Practical application to determining the strength of machine parts and structural materials is discussed.
The objective of this educational exercise is to have students build and evaluate simple wing structures, and in doing so, learn about materials choices and lightweight construction methods. A list of equipment and supplies and the procedure for the experiment are presented.
An experimental investigation was undertaken to quantify damage tolerance and resistance in composite materials impacted using the drop-weight method. Tests were conducted on laminates of several different carbon-fiber composite systems, such as epoxies, modified epoxies, and amorphous and semicrystalline thermoplastics. Impacted composite specimens were examined using destructive and non-destructive techniques to establish the characteristic damage states. Specifically, optical microscopy, ultrasonic, and scanning electron microscopy techniques were used to identify impact induced damage mechanisms. Damage propagation during post impact compression was also studied.
A family of metal/ceramic composite materials has been developed that are relatively inexpensive, lightweight alternatives to structural materials that are typified by beryllium, aluminum, and graphite/epoxy composites. These metal/ceramic composites were originally intended to replace beryllium (which is toxic and expensive) as a structural material for lightweight mirrors for aerospace applications. These materials also have potential utility in automotive and many other terrestrial applications in which there are requirements for lightweight materials that have high strengths and other tailorable properties as described below. The ceramic component of a material in this family consists of hollow ceramic spheres that have been formulated to be lightweight (0.5 g/cm3) and have high crush strength [40.80 ksi (.276.552 MPa)]. The hollow spheres are coated with a metal to enhance a specific performance . such as shielding against radiation (cosmic rays or x rays) or against electromagnetic interference at radio and lower frequencies, or a material to reduce the coefficient of thermal expansion (CTE) of the final composite material, and/or materials to mitigate any mismatch between the spheres and the matrix metal. Because of the high crush strength of the spheres, the initial composite workpiece can be forged or extruded into a high-strength part. The total time taken in processing from the raw ingredients to a finished part is typically 10 to 14 days depending on machining required.
A fabrication feasibility demonstration component for the Enabling Propulsion Materials (EPM) program was evaluated under prototypical engine loading conditions at the Structural Benchmark Test Facility at the NASA Lewis Research Center. The purpose for this test was to verify EPM casting, joining, coating, and life-prediction methods. Electron beam welding techniques developed in the EPM program were used to join two large superalloy cast sections of an exhaust nozzle flap to fabricate the demonstration component. After the joints were inspected, the component was coated with an oxidation-resistant barrier coating and was sent to Lewis for testing. The special test fixture shown in the photo (the Structural Benchmark Test Facility) was designed and built at Lewis to produce a biaxial bending condition similar to the loading condition this part would encounter during engine operation. Several finite element analyses were conducted to validate the mechanical test method. A floating furnace was then designed to provide prototypical thermal profiles in the component. An isothermal low-cycle fatigue test was used to evaluate the component at a cyclic load of 13 kN (maximum) to 1 kN (minimum) at a frequency of 1 Hz. Component failure was defined as a 30-percent increase in the component's compliance. On the basis of this definition, the low-cycle fatigue life of this component would be 35,000 cycles.
The accurate determination of inert strength is important in reliable life prediction of structural ceramic components. At ambient temperature, the inert strength of a brittle material is typically regarded as free of the effects of slow crack growth due to stress corrosion. Therefore, the inert strength can be determined either by eliminating active species, especially moisture, with an appropriate inert medium, or by using a very high test rate. However, at elevated temperatures, the concept or definition of the inert strength of brittle ceramic materials is not clear, since temperature itself is a degrading environment, resulting in strength degradation through slow crack growth and/or creep. Since the mechanism to control strength is rate-dependent viscous flow, the only conceivable way to determine the inert strength at elevated temperatures is to utilize a very fast test rate that either minimizes the time for or eliminates slow crack growth. Few experimental studies have measured the elevated-temperature, inert (or "ultra"-fast fracture) strength of advanced ceramics. At the NASA Lewis Research Center, an experimental study was initiated to better understand the "ultra"-fast fracture strength behavior of advanced ceramics at elevated temperatures. Fourteen advanced ceramics - one alumina, eleven silicon nitrides, and two silicon carbides - have been tested using constant stress-rate (dynamic fatigue) testing in flexure with a series of stress rates including the "ultra"-fast stress rate of 33 000 MPa/sec with digitally controlled test frames. The results for these 14 advanced ceramics indicate that, notwithstanding possible changes in flaw populations as well as flaw configurations because of elevated temperatures, the strength at 33 000 MPa/sec approached the room-temperature strength or reached a higher value than that determined at the conventional test rate of 30 MPa/sec. On the basis of the experimental data, it can be stated that the elevated-temperature, inert strength of an advanced ceramic material can be defined as the strength where no slow crack growth takes place at the temperature.
Non-linear analysis methods were adapted and incorporated in a finite element based DIAL code. These methods are necessary to evaluate the global response of a stiffened structure under combined in-plane and out-of-plane loading. These methods include the Arc Length method and target point analysis procedure. A new interface material model was implemented that can model elastic-plastic behavior of the bond adhesive. Direct application of this method is in skin/stiffener interface failure assessment. Addition of the AML (angle minus longitudinal or load) failure procedure and Hasin's failure criteria provides added capability in the failure predictions. Interactive Stiffened Panel Analysis modules were developed as interactive pre-and post-processors. Each module provides the means of performing self-initiated finite elements based analysis of primary structures such as a flat or curved stiffened panel; a corrugated flat sandwich panel; and a curved geodesic fuselage panel. This module brings finite element analysis into the design of composite structures without the requirement for the user to know much about the techniques and procedures needed to actually perform a finite element analysis from scratch. An interactive finite element code was developed to predict bolted joint strength considering material and geometrical non-linearity. The developed method conducts an ultimate strength failure analysis using a set of material degradation models.
Existing nondestructive evaluation (NDE) technologies are inherently limited by the physical response of the structural material being inspected and are therefore not generally effective at the identification of small discontinuities, making the detection of incipient damage extremely difficult. One innovative solution to this problem is to enhance or complement the NDE signature of structural materials to dramatically improve the ability of existing NDE tools to detect damage. To address this need, a multifunctional metallic material has been developed that can be used in structural applications. The material is processed to contain second phase sensory particles that significantly improve the NDE response, enhancing the ability of conventional NDE techniques to detect incipient damage both during and after flight. Ferromagnetic shape-memory alloys (FSMAs) are an ideal material for these sensory particles as they undergo a uniform and repeatable change in both magnetic properties and crystallographic structure (martensitic transformation) when subjected to strain and/or temperature changes which can be detected using conventional NDE techniques. In this study, the use of a ferromagnetic shape memory alloy (FSMA) as the sensory particles was investigated.
Non-metallic solutions for Lunar bases are appealing due to their light weight and ability to block many forms of radiation. Composite materials made from different combinations of these non-metallic materials enable effective mating of dissimilar materials for strong structures such as pressurized habitats. Inserting and manufacturing composite structural materials on the Moon presents several challenges, however: the oxygen levels inside pressurized habitats will be relatively high, radiation that is blocked causes damage and the lack of atmosphere results in a high-magnitude negative pressure environment. In the past two years, the Materials and Processes Lab (EM01) invested in two internal research and development (IRAD) proposals for testing and procurement of composite materials. One proposal endeavors to determine flammability and radiation damage of various composite materials with a focus on carbon fiber and ultra-high molecular weight polyethylene (UHWMPE). The other proposal enables procurement of chopped fiber spray equipment materials which will be used to develop a novel technique for chopped UHMWPE fiber layup, similar to the chopped fiberglass layups used for making large structures such a boat hulls and hot tubs.