Engineering topics

Scotti, Stephen J.

Publications and source records attributed to Scotti, Stephen J..

At least 19 records

NASA NTRS2019

Characterization of IM7/8552 Thin-Ply and Hybrid Thin-Ply Composites

Composite materials have increasingly been used for aerospace applications due to improved performance and reduced weight compared to their metallic counterparts. Inclusion of thin-ply material, plies with cured thickness half or less than standard composites, have potential to improve performance and reduce structural weight. Limited characterization of thin-ply IM7/8552 material in 30 and 70 grams per square meter fiber areal weights has been carried out using a series of selected American Society for Testing and Materials (ASTM) tests. Tests included unnotched tension, unnotched compression, v-notched rail shear, open-hole tension, and open-hole compression. Unidirectional, cross-ply, quasiisotropic and hybrid hard laminates were included in the study, and were compared to standard-ply laminates. Properties compared include fiber volume, laminate moduli, and failure strength, with failure modes also being examined. The thin-ply specimens exhibited similar or superior performance to standard ply laminates in many of the cases compared. Improvements in strength for laminates containing thin-ply material were seen for unidirectional laminates under unnotched tension, quasi-isotropic laminates under unnotched tension and compression, and hard laminates under open hole tension. Additional investigation is required to determine appropriate ply stacking rules for hybrids of thin and standard plies to avoid undesirable failure modes such as axial splitting. However, the observed performance improvements demonstrated by the conducted ASTM tests of hybrid thin-ply hard laminates could have benefits for improved structural weight in aircraft.

Lovejoy, Andrew E.↗

NASA NTRS2019

Potential Weight Benefits of IM7/8552 Hybrid Thin-Ply Composites for Aircraft Structures

Composite materials have increasingly been used for aerospace applications due to improved performance and reduced weight compared to their metallic counterparts. Inclusion of thin-ply material, plies with cured thickness half or less than standard-ply composites, have potential to improve performance and reduce structural weight further. The effect of thin-ply material on the weight of aircraft structure was investigated by examining wing cover weight reduction. To minimize the effects on manufacturing due to using thin plies, hybrid laminates were examined that used thin ±45-degree plies to replace their standard-ply counterparts in laminates. Compression after impact (CAI) tests were conducted to examine the possible weight savings that could be gained by increasing the design allowables that were used to size the wing upper cover of a semi-span test article. A large increase in CAI strength was observed for quasi-isotropic hybrid laminates, whereas less improvement was seen for hard hybrid laminates such as found in the wing cover. For laminates design by CAI strength, weight savings of about 13% were found using the hybrid hard laminates compared to the standard-ply laminates. Whether similar weight savings could be expected for structure sized using tension after impact allowables will have to be investigated further. Notched specimens were tested to examine possible weight savings using hybrid laminates in regions that are sized using discrete source damage requirements. As expected, the hybrid laminate had marginal improvements over the standard-ply laminate for compression with a notch present. The hybrid laminate, however, exhibited about 20% lower strength than the standard-ply laminate counterpart for tension with a notch. The failure mode of the hybrid specimens was a brittle, self-similar crack, which differs from the standard-ply specimens that failed by significant amounts of delamination and fiber splitting. In light of the apparent reduction in notched tensile strength, additional investigation is required to assess the use of hybrid laminates for areas containing discrete source damage, and their effect on weight of such regions.

Lovejoy, Andrew E.↗

NASA NTRS2016

Self-Aware Vehicles: Mission and Performance Adaptation to System Health

Advances in sensing (miniaturization, distributed sensor networks) combined with improvements in computational power leading to significant gains in perception, real-time decision making/reasoning and dynamic planning under uncertainty as well as big data predictive analysis have set the stage for realization of autonomous system capability. These advances open the design and operating space for self-aware vehicles that are able to assess their own capabilities and adjust their behavior to either complete the assigned mission or to modify the mission to reflect their current capabilities. This paper discusses the self-aware vehicle concept and associated technologies necessary for full exploitation of the concept. A self-aware aircraft, spacecraft or system is one that is aware of its internal state, has situational awareness of its environment, can assess its capabilities currently and project them into the future, understands its mission objectives, and can make decisions under uncertainty regarding its ability to achieve its mission objectives.

Gregory, Irene M.↗

NASA NTRS2016

A Study of Flexible Composites for Expandable Space Structures

Payload volume for launch vehicles is a critical constraint that impacts spacecraft design. Deployment mechanisms, such as those used for solar arrays and antennas, are approaches that have successfully accommodated this constraint, however, providing pressurized volumes that can be packaged compactly at launch and expanded in space is still a challenge. One approach that has been under development for many years is to utilize softgoods - woven fabric for straps, cloth, and with appropriate coatings, bladders - to provide this expandable pressure vessel capability. The mechanics of woven structure is complicated by a response that is nonlinear and often nonrepeatable due to the discrete nature of the woven fiber architecture. This complexity reduces engineering confidence to reliably design and certify these structures, which increases costs due to increased requirements for system testing. The present study explores flexible composite materials systems as an alternative to the heritage softgoods approach. Materials were obtained from vendors who utilize flexible composites for non-aerospace products to determine some initial physical and mechanical properties of the materials. Uniaxial mechanical testing was performed to obtain the stress-strain response of the flexible composites and the failure behavior. A failure criterion was developed from the data, and a space habitat application was used to provide an estimate of the relative performance of flexible composites compared to the heritage softgoods approach. Initial results are promising with a 25% mass savings estimated for the flexible composite solution.

Scotti, Stephen J.↗

NASA NTRS2012

Combination structural support and thermal protection system

A combination structural support and thermal protection system is provided. A cured preceramic polymer matrix reinforced with carbon fibers can be used alone or in a multiple component system.

Rivers, H. Kevin↗

NASA NTRS2012

Compression After Impact on Honeycomb Core Sandwich Panels With Thin Facesheets: Experiments - Part 1

McQuigg, Thomas D.↗

NASA NTRS2012

Accounting for Proof Test Data in a Reliability Based Design Optimization Framework

This paper investigates the use of proof (or acceptance) test data during the reliability based design optimization of structural components. It is assumed that every component will be proof tested and that the component will only enter into service if it passes the proof test. The goal is to reduce the component weight, while maintaining high reliability, by exploiting the proof test results during the design process. The proposed procedure results in the simultaneous design of the structural component and the proof test itself and provides the designer with direct control over the probability of failing the proof test. The procedure is illustrated using two analytical example problems and the results indicate that significant weight savings are possible when exploiting the proof test results during the design process.

Ventor, Gerharad↗

NASA NTRS2012

Compression After Impact on Honeycomb Core Sandwich Panels with Thin Facesheets, Part 2: Analysis

A two part research study has been completed on the topic of compression after impact (CAI) of thin facesheet honeycomb core sandwich panels. The research has focused on both experiments and analysis in an effort to establish and validate a new understanding of the damage tolerance of these materials. Part 2, the subject of the current paper, is focused on the analysis, which corresponds to the CAI testings described in Part 1. Of interest, are sandwich panels, with aerospace applications, which consist of very thin, woven S2-fiberglass (with MTM45-1 epoxy) facesheets adhered to a Nomex honeycomb core. Two sets of materials, which were identical with the exception of the density of the honeycomb core, were tested in Part 1. The results highlighted the need for analysis methods which taken into account multiple failure modes. A finite element model (FEM) is developed here, in Part 2. A commercial implementation of the Multicontinuum Failure Theory (MCT) for progressive failure analysis (PFA) in composite laminates, Helius:MCT, is included in this model. The inclusion of PFA in the present model provided a new, unique ability to account for multiple failure modes. In addition, significant impact damage detail is included in the model. A sensitivity study, used to assess the effect of each damage parameter on overall analysis results, is included in an appendix. Analysis results are compared to the experimental results for each of the 32 CAI sandwich panel specimens tested to failure. The failure of each specimen is predicted using the high-fidelity, physicsbased analysis model developed here, and the results highlight key improvements in the understanding of honeycomb core sandwich panel CAI failure. Finally, a parametric study highlights the strength benefits compared to mass penalty for various core densities.

Mcquigg, Thomas D.↗

NASA NTRS2012

Flexible Plug Repair for Shuttle Wing Leading Edge

In response to the Columbia Accident Investigation Board report, a plug repair kit has been developed to enable astronauts to repair the space shuttle's wing leading edge (WLE) during orbit. The plug repair kit consists of several 17.78- cm-diameter carbon/silicon carbide (C/SiC) cover plates of various curvatures that can be attached to the refractory carbon-carbon WLE panels using a TZM refractory metal attach mechanism. The attach mechanism is inserted through the damage in the WLE panel and, as it is tightened, the cover plate flexes to conform to the curvature of the WLE panel within 0.050 mm. An astronaut installs the repair during an extravehicular activity (EVA). After installing the plug repair, edge gaps are checked and the perimeter of the repair is sealed using a proprietary material, developed to fill cracks and small holes in the WLE.

Camarda, Charles J.↗

NASA NTRS2012

Transpiration-Cooled Spacecraft-Insulation-Repair Fasteners

Transpiration-cooled fasteners are proposed that operate like an open-loop heat pipe (self-tapping screws, bolts, and spikes) for use in on-orbit repair of thermal- insulation of a space shuttle or other spacecraft. By limiting the temperature rise of such a fastener and of the adjacent repair material and thermal protection system, the transpiration cooling would contribute to the ability of the repair to retain its strength and integrity in the high-heat-flux, oxidizing environment of reentry into the atmosphere of the Earth. A typical fastener according to the proposal would include a hollow refractory-metal, refractory-composite, or ceramic screw or bolt, the central cavity of which would be occupied by a porous refractory- metal or ceramic plug that would act as both a reservoir and a wick for a transpirant liquid. The plug dimensions, the plug material, and the sizes of the pores would be chosen in conjunction with the transpirant liquid so that (1) capillary pumping could be relied upon to transport the liquid to the heated surface, where the liquid would be vaporized, and (2) the amount of liquid would suffice for protecting against the anticipated heat flux and integrated heat load.

Camarda, Charles J.↗

NASA NTRS2011

Multi-functional annular fairing for coupling launch abort motor to space vehicle

An annular fairing having aerodynamic, thermal, structural and acoustic attributes couples a launch abort motor to a space vehicle having a payload of concern mounted on top of a rocket propulsion system. A first end of the annular fairing is fixedly attached to the launch abort motor while a second end of the annular fairing is attached in a releasable fashion to an aft region of the payload. The annular fairing increases in diameter between its first and second ends.

Camarda, Charles J.↗

NASA NTRS2011

Response of Honeycomb Core Sandwich Panel with Minimum Gage GFRP Face-Sheets to Compression Loading After Impact

A compression after impact study has been conducted to determine the residual strength of three sandwich panel constructions with two types of thin glass fiber reinforced polymer face-sheets and two hexagonal honeycomb Nomex core densities. Impact testing is conducted to first determine the characteristics of damage resulting from various impact energy levels. Two modes of failure are found during compression after impact tests with the density of the core precipitating the failure mode present for a given specimen. A finite element analysis is presented for prediction of the residual compressive strength of the impacted specimens. The analysis includes progressive damage modeling in the face-sheets. Preliminary analysis results were similar to the experimental results; however, a higher fidelity core material model is expected to improve the correlation.

McQuigg, Thomas D.↗

NASA NTRS2007

Representing Functions in n Dimensions to Arbitrary Accuracy

A method of approximating a scalar function of n independent variables (where n is a positive integer) to arbitrary accuracy has been developed. This method is expected to be attractive for use in engineering computations in which it is necessary to link global models with local ones or in which it is necessary to interpolate noiseless tabular data that have been computed from analytic functions or numerical models in n-dimensional spaces of design parameters.

Scotti, Stephen J.↗

NASA NTRS2003

METAShield: Hot Metallic Aeroshell Concept for RLV/SOV

An innovative fuselage design approach that combines many desirable operational features with a simple and efficient structural approach is being developed by NASA. The approach, named METAShield for MEtallic TransAtmospheric Shield, utilizes lightly loaded, hot aeroshell structures surrounding integral propellant tanks that carry the primary structural loads. The aeroshells are designed to withstand the local pressure loads, transmitting them to the tanks with minimal restraint of thermal growth. No additional thermal protection system protects the METAShield, and a fibrous or multilayer insulation blanket, located in the space between the aeroshell and the tanks, serves as both high temperature and cryogenic insulation for the tanks. The concept is described in detail, and the performance and operational features are highlighted. Initial design results and analyses of the structural, thermal, and thermal-structural performance are described. Computational results evaluating resistance to hypervelocity impact damage, as well as some supporting aerothermal wind tunnel results. are also presented. Future development needs are summarized.

Scotti, Stephen J.↗

NASA NTRS2003

Structures and Materials Technologies for Extreme Environments Applied to Reusable Launch Vehicles

This paper provides an overview of the evolution of structures and materials technology approaches to survive the challenging extreme environments encountered by earth-to-orbit space transportation systems, with emphasis on more recent developments in the USA. The evolution of technology requirements and experience in the various approaches to meeting these requirements has significantly influenced the technology approaches. While previous goals were primarily performance driven, more recently dramatic improvements in costs/operations and in safety have been paramount goals. Technologies that focus on the cost/operations and safety goals in the area of hot structures and thermal protection systems for reusable launch vehicles are presented. Assessments of the potential ability of the various technologies to satisfy the technology requirements, and their current technology readiness status are also presented.

Scotti, Stephen J.↗

NASA NTRS1997

Transpiration Cooling Experiment

The transpiration cooling method was considered for a scram-jet engine to accommodate thermally the situation where a very high heat flux (200 Btu/sq. ft sec) from hydrogen fuel combustion process is imposed to the engine walls. In a scram-jet engine, a small portion of hydrogen fuel passes through the porous walls of the engine combustor to cool the engine walls and at the same time the rest passes along combustion chamber walls and is preheated. Such a regenerative system promises simultaneously cooling of engine combustor and preheating the cryogenic fuel. In the experiment, an optical heating method was used to provide a heat flux of 200 Btu/sq. ft sec to the cylindrical surface of a porous stainless steel specimen which carried helium gas. The cooling efficiencies by transpiration were studied for specimens with various porosity. The experiments of various test specimens under high heat flux have revealed a phenomenon that chokes the medium flow when passing through a porous structure. This research includes the analysis of the system and a scaling conversion study that interprets the results from helium into the case when hydrogen medium is used.

Song, Kyo D.↗

NASA NTRS1997

Transpiring Cooling of a Scram-Jet Engine Combustion Chamber

The peak cold-wall heating rate generated in a combustion chamber of a scram-jet engine can exceed 2000 Btu/sq ft sec (approx. 2344 W/sq cm). Therefore, a very effective heat dissipation mechanism is required to sustain such a high heating load. This research focused on the transpiration cooling mechanism that appears to be a promising approach to remove a large amount of heat from the engine wall. The transpiration cooling mechanism has two aspects. First, initial computations suggest that there is a reduction, as much as 75%, in the heat flux incident on the combustion chamber wall due to the transpirant modifying the combustor boundary layer. Secondly, the heat reaching the combustor wall is removed from the structure in a very effective manner by the transpirant. It is the second of these two mechanisms that is investigated experimentally in the subject paper. A transpiration cooling experiment using a radiant heating method, that provided a heat flux as high as 200 Btu/sq ft sec ( approx. 234 W/sq cm) on the surface of a specimen, was performed. The experiment utilized an arc-lamp facility (60-kW radiant power output) to provide a uniform heat flux to a test specimen. For safety reasons, helium gas was used as the transpirant in the experiments. The specimens were 1.9-cm diameter sintered, powdered-stainless-steel tubes of various porosities and a 2.54cm square tube with perforated multi-layered walls. A 15-cm portion of each specimen was heated. The cooling effectivenes and efficiencies by transpiration for each specimen were obtained using the experimental results. During the testing, various test specimens displayed a choking phenomenon in which the transpirant flow was limited as the heat flux was increased. The paper includes a preliminary analysis of the transpiration cooling mechanism and a scaling conversion study that translates the results from helium tests into the case when a hydrogen medium is used.

Choi, Sang H.↗

NASA NTRS1995

Structural design using equilibrium programming formulations

Solutions to increasingly larger structural optimization problems are desired. However, computational resources are strained to meet this need. New methods will be required to solve increasingly larger problems. The present approaches to solving large-scale problems involve approximations for the constraints of structural optimization problems and/or decomposition of the problem into multiple subproblems that can be solved in parallel. An area of game theory, equilibrium programming (also known as noncooperative game theory), can be used to unify these existing approaches from a theoretical point of view (considering the existence and optimality of solutions), and be used as a framework for the development of new methods for solving large-scale optimization problems. Equilibrium programming theory is described, and existing design techniques such as fully stressed design and constraint approximations are shown to fit within its framework. Two new structural design formulations are also derived. The first new formulation is another approximation technique which is a general updating scheme for the sensitivity derivatives of design constraints. The second new formulation uses a substructure-based decomposition of the structure for analysis and sensitivity calculations. Significant computational benefits of the new formulations compared with a conventional method are demonstrated.

Scotti, Stephen J.↗