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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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At least 595 records · Page 33
All sourcesNASA NTRSDOE OSTI
NASA NTRS2017

Assessment of Intralaminar Progressive Damage and Failure Analysis Using an Efficient Evaluation Framework

Reducing the timeline for development and certification for composite structures has been a long standing objective of the aerospace industry. This timeline can be further exacerbated when attempting to integrate new fiber-reinforced composite materials due to the large number of testing required at every level of design. computational progressive damage and failure analysis (PDFA) attempts to mitigate this effect; however, new PDFA methods have been slow to be adopted in industry since material model evaluation techniques have not been fully defined. This study presents an efficient evaluation framework which uses a piecewise verification and validation (V&V) approach for PDFA methods. Specifically, the framework is applied to evaluate PDFA research codes within the context of intralaminar damage. Methods are incrementally taken through various V&V exercises specifically tailored to study PDFA intralaminar damage modeling capability. Finally, methods are evaluated against a defined set of success criteria to highlight successes and limitations.

Hyder, Imran↗
NASA NTRS1992

CFD analysis to optimize a design modification of BSMT

The Bearings, Seals and Materials Tester (BSMT) is a test article being used at MSFC to evaluate the performance of conventional rolling contact bearings. Pressure differentials between the BSMT inlet and exit cavities are found to cause large parasitic axial loads on the bearing-carrier walls. These parasitic loads, besides being detrimental to the life of bearings, make the testing and evaluation of bearing performance very difficult, and need to be eliminated if at all possible. The objectives of this study are to estimate the hydrodynamic loads on the bearings and to recommend feasible design modifications for BSMT to eliminate the parasitic loads. Three-dimensional computational analyses of inlet and exit cavities in their baseline configuration were performed with REFLEQS which is an advanced finite-volume Navier-Stokes code. Computations were performed with and without a 1/4 inch diameter temperature probe included in each of the cavities. The results of the analyses indicate that the temperature probes substantially alter the flow field and reduce the pressure drop/rise in the cavities. The overall pressure drop across the tester compares quite well with the measurements.

Ratcliff, Mark↗
NASA NTRS1984

Alternate nozzle ablative materials program

Four subscale solid rocket motor tests were conducted successfully to evaluate alternate nozzle liner, insulation, and exit cone structural overwrap components for possible application to the Space Shuttle Solid Rocket Motor (SRM) nozzle asasembly. The 10,000 lb propellant motor tests were simulated, as close as practical, the configuration and operational environment of the full scale SRM. Fifteen PAN based and three pitch based materials had no filler in the phenolic resin, four PAN based materials had carbon microballoons in the resin, and the rest of the materials had carbon powder in the resin. Three nozzle insulation materials were evaluated; an aluminum oxide silicon oxide ceramic fiber mat phenolic material with no resin filler and two E-glass fiber mat phenolic materials with no resin filler. It was concluded by MTI/WD (the fabricator and evaluator of the test nozzles) and NASA-MSFC that it was possible to design an alternate material full scale SRM nozzle assembly, which could provide an estimated 360 lb increased payload capability for Space Shuttle launches over that obtainable with the current qualified SRM design.

Kimmel, N. A.↗
NASA NTRS2012

SPRITE: A TPS Test Bed for Ground and Flight

Engineers in the Entry Systems and Technology Division at NASA Ames Research Center developed a fully instrumented, small atmospheric entry probe called SPRITE (Small Probe Reentry Investigation for TPS Engineering). SPRITE, conceived as a flight test bed for thermal protection materials, was tested at full scale in an arc-jet facility so that the aerothermal environments the probe experiences over portions of its flight trajectory and in the arc-jet are similar. This ground-to-flight traceability enhances the ability of mission designers to evaluate margins needed in the design of thermal protection systems (TPS) of larger scale atmospheric entry vehicles. SPRITE is a 14-inch diameter, 45 deg. sphere-cone with a conical aftbody and designed for testing in the NASA Ames Aerodynamic Heating Facility (AHF). The probe is a two-part aluminum shell with PICA (phenolic impregnated carbon ablator) bonded on the forebody and LI-2200 (Shuttle tile material) bonded to the aftbody. Plugs with embedded thermocouples, similar to those installed in the heat shield of the Mars Science Laboratory (MSL), and a number of distributed sensors are integrated into the design. The data from these sensors are fed to an innovative, custom-designed data acquisition system also integrated with the test article. Two identical SPRITE models were built and successfully tested in late 2010-early 2011, and the concept is currently being modified to enable testing of conformable and/or flexible materials.

Ablation↗
NASA NTRS1994

Analytical Design Package (ADP2): A computer aided engineering tool for aircraft transparency design

The Analytical Design Package (ADP2) is being developed as a part of the Air Force Frameless Transparency Program (FTP). ADP2 is an integrated design tool consisting of existing analysis codes and Computer Aided Engineering (CAE) software. The objective of the ADP2 is to develop and confirm an integrated design methodology for frameless transparencies, related aircraft interfaces, and their corresponding tooling. The application of this methodology will generate high confidence for achieving a qualified part prior to mold fabrication. ADP2 is a customized integration of analysis codes, CAE software, and material databases. The primary CAE integration tool for the ADP2 is P3/PATRAN, a commercial-off-the-shelf (COTS) software tool. The open architecture of P3/PATRAN allows customized installations with different applications modules for specific site requirements. Integration of material databases allows the engineer to select a material, and those material properties are automatically called into the relevant analysis code. The ADP2 materials database will be composed of four independent schemas: CAE Design, Processing, Testing, and Logistics Support. The design of ADP2 places major emphasis on the seamless integration of CAE and analysis modules with a single intuitive graphical interface. This tool is being designed to serve and be used by an entire project team, i.e., analysts, designers, materials experts, and managers. The final version of the software will be delivered to the Air Force in Jan. 1994. The Analytical Design Package (ADP2) will then be ready for transfer to industry. The package will be capable of a wide range of design and manufacturing applications.

Wuerer, J. E.↗
NASA NTRS1977

Ion beam plume and efflux characterization flight experiment study

A flight experiment and flight experiment package for a shuttle-borne flight test of an 8-cm mercury ion thruster was designed to obtain charged particle and neutral particle material transport data that cannot be obtained in conventional ground based laboratory testing facilities. By the use of both ground and space testing of ion thrusters, the flight worthiness of these ion thrusters, for other spacecraft applications, may be demonstrated. The flight experiment definition for the ion thruster initially defined a broadly ranging series of flight experiments and flight test sensors. From this larger test series and sensor list, an initial flight test configuration was selected with measurements in charged particle material transport, condensible neutral material transport, thruster internal erosion, ion beam neutralization, and ion thrust beam/space plasma electrical equilibration. These measurement areas may all be examined for a seven day shuttle sortie mission and for available test time in the 50 - 100 hour period.

Sellen, J. M., Jr.↗
NASA NTRS1994

Design of passive piezoelectric damping for space structures

Passive damping of structural dynamics using piezoceramic electromechanical energy conversion and passive electrical networks is a relatively recent concept with little implementation experience base. This report describes an implementation case study, starting from conceptual design and technique selection, through detailed component design and testing to simulation on the structure to be damped. About 0.5kg. of piezoelectric material was employed to damp the ASTREX testbed, a 500kg structure. Emphasis was placed upon designing the damping to enable high bandwidth robust feedback control. Resistive piezoelectric shunting provided the necessary broadband damping. The piezoelectric element was incorporated into a mechanically-tuned vibration absorber in order to concentrate damping into the 30 to 40 Hz frequency modes at the rolloff region of the proposed compensator. A prototype of a steel flex-tensional motion amplification device was built and tested. The effective stiffness and damping of the flex-tensional device was experimentally verified. When six of these effective springs are placed in an orthogonal configuration, strain energy is absorbed from all six degrees of freedom of a 90kg. mass. A NASTRAN finite element model of the testbed was modified to include the six-spring damping system. An analytical model was developed for the spring in order to see how the flex-tensional device and piezoelectric dimensions effect the critical stress and strain energy distribution throughout the component. Simulation of the testbed demonstrated the damping levels achievable in the completed system.

Hagood, Nesbitt W., IV↗
NASA NTRS1982

Development of high strength, high temperature ceramics

Improvement in the high-pressure turbopumps, both fuel and oxidizer, in the Space Shuttle main engine were considered. The operation of these pumps is limited by temperature restrictions of the metallic components used in these pumps. Ceramic materials that retain strength at high temperatures and appear to be promising candidates for use as turbine blades and impellers are discussed. These high strength materials are sensitive to many related processing parameters such as impurities, sintering aids, reaction aids, particle size, processing temperature, and post thermal treatment. The specific objectives of the study were to: (1) identify and define the processing parameters that affect the properties of Si3N4 ceramic materials, (2) design and assembly equipment required for processing high strength ceramics, (3) design and assemble test apparatus for evaluating the high temperature properties of Si3N4, and (4) conduct a research program of manufacturing and evaluating Si3N4 materials as applicable to rocket engine applications.

Hall, W. B.↗
NASA NTRS2001

Laser High-Cycle Thermal Fatigue of Pulse Detonation Engine Combustor Materials Tested

Pulse detonation engines (PDE's) have received increasing attention for future aerospace propulsion applications. Because the PDE is designed for a high-frequency, intermittent detonation combustion process, extremely high gas temperatures and pressures can be realized under the nearly constant-volume combustion environment. The PDE's can potentially achieve higher thermodynamic cycle efficiency and thrust density in comparison to traditional constant-pressure combustion gas turbine engines (ref. 1). However, the development of these engines requires robust design of the engine components that must endure harsh detonation environments. In particular, the detonation combustor chamber, which is designed to sustain and confine the detonation combustion process, will experience high pressure and temperature pulses with very short durations (refs. 2 and 3). Therefore, it is of great importance to evaluate PDE combustor materials and components under simulated engine temperatures and stress conditions in the laboratory. In this study, a high-cycle thermal fatigue test rig was established at the NASA Glenn Research Center using a 1.5-kW CO2 laser. The high-power laser, operating in the pulsed mode, can be controlled at various pulse energy levels and waveform distributions. The enhanced laser pulses can be used to mimic the time-dependent temperature and pressure waves encountered in a pulsed detonation engine. Under the enhanced laser pulse condition, a maximum 7.5-kW peak power with a duration of approximately 0.1 to 0.2 msec (a spike) can be achieved, followed by a plateau region that has about one-fifth of the maximum power level with several milliseconds duration. The laser thermal fatigue rig has also been developed to adopt flat and rotating tubular specimen configurations for the simulated engine tests. More sophisticated laser optic systems can be used to simulate the spatial distributions of the temperature and shock waves in the engine. Pulse laser high-cycle thermal fatigue behavior has been investigated on a flat Haynes 188 alloy specimen, under the test condition of 30-Hz cycle frequency (33-msec pulse period and 10-msec pulse width including a 0.2-msec pulse spike; ref. 4). Temperature distributions were calculated with one-dimensional finite difference models. The calculations show that that the 0.2-msec pulse spike can cause an additional 40 C temperature fluctuation with an interaction depth of 0.08 mm near the specimen surface region. This temperature swing will be superimposed onto the temperature swing of 80 C that is induced by the 10-msec laser pulse near the 0.53-mm-deep surface interaction region.

Zhu, Dong-Ming↗
NASA NTRS1993

Space Station technology testbed: 2010 deep space transport

A space station in a crew-tended or permanently crewed configuration will provide major R&D opportunities for innovative, technology and materials development and advanced space systems testing. A space station should be designed with the basic infrastructure elements required to grow into a major systems technology testbed. This space-based technology testbed can and should be used to support the development of technologies required to expand our utilization of near-Earth space, the Moon and the Earth-to-Jupiter region of the Solar System. Space station support of advanced technology and materials development will result in new techniques for high priority scientific research and the knowledge and R&D base needed for the development of major, new commercial product thrusts. To illustrate the technology testbed potential of a space station and to point the way to a bold, innovative approach to advanced space systems' development, a hypothetical deep space transport development and test plan is described. Key deep space transport R&D activities are described would lead to the readiness certification of an advanced, reusable interplanetary transport capable of supporting eight crewmembers or more. With the support of a focused and highly motivated, multi-agency ground R&D program, a deep space transport of this type could be assembled and tested by 2010. Key R&D activities on a space station would include: (1) experimental research investigating the microgravity assisted, restructuring of micro-engineered, materials (to develop and verify the in-space and in-situ 'tuning' of materials for use in debris and radiation shielding and other protective systems), (2) exposure of microengineered materials to the space environment for passive and operational performance tests (to develop in-situ maintenance and repair techniques and to support the development, enhancement, and implementation of protective systems, data and bio-processing systems, and virtual reality and telepresence/kinetic processes), (3) subsystem tests of advanced nuclear power, nuclear propulsion and communication systems (using boom extensions, remote station-keeping platforms and mobile EVA crew and robots), and (4) logistics support (crew and equipment) and command and control of deep space transport assembly, maintenance, and refueling (using a station-keeping platform).

Holt, Alan C.↗
NASA NTRS2009

Advanced Design Heat PumpRadiator for EVA Suits

Absorption cooling using a LiCl/water heat pump can enable lightweight and effective thermal control for EVA suits without venting water to the environment. The key components in the system are an absorber/radiator that rejects heat to space and a flexible evaporation cooling garment that absorbs heat from the crew member. This paper describes progress in the design, development, and testing of the absorber/radiator and evaporation cooling garment. New design concepts and fabrication approaches will significantly reduce the mass of the absorber/radiator. We have also identified materials and demonstrated fabrication approaches for production of a flexible evaporation cooling garment. Data from tests of the absorber/radiator s modular components have validated the design models and allowed predictions of the size and weight of a complete system.

Izenson, Michael G.↗
NASA NTRS2003

PEM-INST-001: Instructions for Plastic Encapsulated Microcircuit (PEM) Selection, Screening, and Qualification

Potential users of plastic encapsulated microcircuits (PEMs) need to be reminded that unlike the military system of producing robust high-reliability microcircuits that are designed to perform acceptably in a variety of harsh environments, PEMs are primarily designed for use in benign environments where equipment is easily accessed for repair or replacement. The methods of analysis applied to military products to demonstrate high reliability cannot always be applied to PEMs. This makes it difficult for users to characterize PEMs for two reasons: 1. Due to the major differences in design and construction, the standard test practices used to ensure that military devices are robust and have high reliability often cannot be applied to PEMs that have a smaller operating temperature range and are typically more frail and susceptible to moisture absorption. In contrast, high-reliability military microcircuits usually utilize large, robust, high-temperature packages that are hermetically sealed. 2. Unlike the military high-reliability system, users of PEMs have little visibility into commercial manufacturers proprietary design, materials, die traceability, and production processes and procedures. There is no central authority that monitors PEM commercial product for quality, and there are no controls in place that can be imposed across all commercial manufacturers to provide confidence to high-reliability users that a common acceptable level of quality exists for all PEMs manufacturers. Consequently, there is no guaranteed control over the type of reliability that is built into commercial product, and there is no guarantee that different lots from the same manufacturer are equally acceptable. And regarding application, there is no guarantee that commercial products intended for use in benign environments will provide acceptable performance and reliability in harsh space environments. The qualification and screening processes contained in this document are intended to detect poor-quality lots and screen out early random failures from use in space flight hardware. However, since it cannot be guaranteed that quality was designed and built into PEMs that are appropriate for space applications, users cannot screen in quality that may not exist. It must be understood that due to the variety of materials, processes, and technologies used to design and produce PEMs, this test process may not accelerate and detect all failure mechanisms. While the tests herein will increase user confidence that PEMs with otherwise unknown reliability can be used in space environments, such testing may not guarantee the same level of reliability offered by military microcircuits. PEMs should only be used where due to performance needs there are no alternatives in the military high-reliability market, and projects are willing to accept higher risk.

Teverovsky, Alexander↗
NASA NTRS2026

Artemis Suit Material Optical Property Testing of Dust Exposed Fabrics

This paper highlights one aspect of NASA’s ongoing technology-infusion effort to design, fabricate, and test a next-generation outer shell fabric for a lunar Extravehicular Activity (EVA) space suit, a critical component of sustained lunar exploration. Managing thermal loads on the Moon is essential for astronaut safety and suit performance. The suit’s exterior fabric directly influences heat gain and loss through its optical properties: low solar absorptivity minimizes sunlight absorption, while high infrared emissivity aids radiative cooling. Lunar regolith complicates this balance. Its fine, abrasive particles possess unique optical behavior that can lower reflectivity and raise emissivity when embedded in or adhered to fabric surfaces, degrading thermal control and increasing the risk of overheating or cooling inefficiency. To quantify these effects, the Artemis Suit Materials (ASM) team measured solar absorptance and infrared emissivity of clean and dust-soiled Ortho Fabric, establishing beginning-of-life (BOL) and end-of-life (EOL) benchmarks. EOL conditions were simulated with a rotary tumbler abrasion process using lunar dust simulant and ceramic media to reproduce cumulative wear expected during surface operations. Tests also included unmodified fabrics and a fabric/film laminate system containing titanium dioxide to evaluate potential improvements in dust resistance and optical performance. Results from these evaluations provide critical insight into how lunar dust alters fabric thermal behavior and inform the design of bespoke suit materials that maintain required optical properties throughout mission life, supporting safe and effective long-duration EVA on the lunar surface.

space↗
NASA NTRS2026

Artemis Suit Material Optical Property Testing of Dust Exposed Fabrics

This paper highlights one aspect of NASA’s ongoing technology-infusion effort to design, fabricate, and test a next-generation outer shell fabric for a lunar Extravehicular Activity (EVA) space suit, a critical component of sustained lunar exploration. Managing thermal loads on the Moon is essential for astronaut safety and suit performance. The suit’s exterior fabric directly influences heat gain and loss through its optical properties: low solar absorptivity minimizes sunlight absorption, while high infrared emissivity aids radiative cooling. Lunar regolith complicates this balance. Its fine, abrasive particles possess unique optical behavior that can lower reflectivity and raise emissivity when embedded in or adhered to fabric surfaces, degrading thermal control and increasing the risk of overheating or cooling inefficiency. To quantify these effects, the Artemis Suit Materials (ASM) team measured solar absorptance and infrared emissivity of clean and dust-soiled Ortho Fabric, establishing beginning-of-life (BOL) and end-of-life (EOL) benchmarks. EOL conditions were simulated with a rotary tumbler abrasion process using lunar dust simulant and ceramic media to reproduce cumulative wear expected during surface operations. Tests also included unmodified fabrics and a fabric/film laminate system containing titanium dioxide to evaluate potential improvements in dust resistance and optical performance. Results from these evaluations provide critical insight into how lunar dust alters fabric thermal behavior and inform the design of bespoke suit materials that maintain required optical properties throughout mission life, supporting safe and effective long-duration EVA on the lunar surface.

Textile↗
NASA NTRS2004

Particulate Titanium Matrix Composites Tested--Show Promise for Space Propulsion Applications

Uniformly distributed particle-strengthened titanium matrix composites (TMCs) can be manufactured at lower cost than many types of continuous-fiber composites. The innovative manufacturing technology combines cold and hot isostatic pressing procedures to produce near-final-shape components. Material stiffness is increased up to 26-percent greater than that of components made with conventional titanium materials at no significant increase in the weight. The improved mechanical performance and low-cost manufacturing capability motivated an independent review to assess the improved properties of ceramic titanium carbide (TiC) particulate-reinforced titanium at elevated temperature. Researchers at the NASA Glenn Research Center creatively designed and executed deformation and durability tests to reveal operating regimes where these materials could lower the cost and weight of space propulsion systems. The program compares the elevated-temperature performance of titanium alloy Ti-6Al-4V matrix material to an alloy containing 10 wt% of TiC particles. Initial experiments showed that at these relatively low particle concentrations the material stiffness of the TMC was improved 20 percent over that of the plain Ti-6Al-4V alloy when tested at 427 C. The proportional limit and ultimate strength of the composite in tension are 21- and 14-percent greater than those of the plain alloy. Compression tests showed that the proportional limit is about 30 percent greater for TMC than for the plain alloy. The enhanced deformation resistance of the TMC was also evident in a series of tensile and compressive stress relaxation tests that were made. Specimens were subjected to tensile or compressive strain amplitudes of 0.75 percent for 24 hr followed by a return to zero strain imposed for 24 hr. The stress relaxation data were normalized with respect to the maximum stress for each case and plotted as a function of time in the following graph. Tensile stresses relaxed 19 percent for the TMC and 25 percent for the plain Ti-6Al-4V alloy. Compressive stresses relaxed 25 percent for the TMC and 39 percent for the plain Ti-6Al-4V alloy. The superior deformation resistance of the TMC extends to a creep rate that is 28-percent slower for the TMC when it is loaded to stress levels that are 26-percent higher than for the plain Ti-6Al-4V alloy.

Lerch, Bradley A.↗
NASA NTRS2016

Fracture Mechanics Testing of Titanium 6AL-4V in AF-M315E

The Green Propellant Infusion Mission (GPIM) will demonstrate the performance of AF-M315E monopropellant on orbit. Flight certification requires a safe-life analysis of the titanium alloy fuel tank to ensure inherent processing flaws will not cause failure during the design life of the tank. Material property inputs for this analysis require testing to determine the stress intensity factor for environment-assisted cracking (KEAC) of Ti 6Al-4V in combination with the AF-M315E monopropellant. Testing of single-edge notched, or SE(B), specimens representing the bulk tank membrane and weld material were performed in accordance with ASTM E1681. Specimens with fatigue pre-cracks were loaded into test fixtures so that the crack tips were exposed to AF-M315E at 50 C for a duration of 1,000 hours. Specimens that did not fail during exposure were opened to inspect the crack surfaces for evidence of crack growth. The threshold stress intensity value, KEAC, is the highest applied stress intensity that produced neither a failure of the specimen during the exposure nor showed evidence of crack growth. The threshold stress intensity factor for environment-assisted cracking of the Ti 6Al-4V forged tank material was found to be at least 22 ksi√in and at least 31 ksi√in for the weld material when exposed to AF-M315E monopropellant.

AF-M315E↗
NASA NTRS2014

Affordable Development and Optimization of CERMET Fuels for NTP Ground Testing

CERMET fuel materials for Nuclear Thermal Propulsion (NTP) are currently being developed at NASA's Marshall Space Flight Center. The work is part of NASA's Advanced Space Exploration Systems Nuclear Cryogenic Propulsion Stage (NCPS) Project. The goal of the FY12-14 project is to address critical NTP technology challenges and programmatic issues to establish confidence in the affordability and viability of an NTP system. A key enabling technology for an NCPS system is the fabrication of a stable high temperature nuclear fuel form. Although much of the technology was demonstrated during previous programs, there are currently no qualified fuel materials or processes. The work at MSFC is focused on developing critical materials and process technologies for manufacturing robust, full-scale CERMET fuels. Prototypical samples are being fabricated and tested in flowing hot hydrogen to understand processing and performance relationships. As part of this initial demonstration task, a final full scale element test will be performed to validate robust designs. The next phase of the project will focus on continued development and optimization of the fuel materials to enable future ground testing. The purpose of this paper is to provide a detailed overview of the CERMET fuel materials development plan. The overall CERMET fuel development path is shown in Figure 2. The activities begin prior to ATP for a ground reactor or engine system test and include materials and process optimization, hot hydrogen screening, material property testing, and irradiation testing. The goal of the development is to increase the maturity of the fuel form and reduce risk. One of the main accomplishmens of the current AES FY12-14 project was to develop dedicated laboratories at MSFC for the fabrication and testing of full length fuel elements. This capability will enable affordable, near term development and optimization of the CERMET fuels for future ground testing. Figure 2 provides a timeline of the development and optimization tasks for the AES FY15-17 follow on program.

Hickman, Robert R.↗
NASA NTRS1990

Conceptual design of a Moving Belt Radiator (MBR) shuttle-attached experiment

The conceptual design of a shuttle-attached Moving Belt Radiator (MBR) experiment is presented. The MBR is an advanced radiator concept in which a rotating belt is used to radiate thermal energy to space. The experiment is developed with the primary focus being the verification of the dynamic characteristics of a rotating belt with a secondary objective of proving the thermal and sealing aspects in a reduced gravity, vacuum environment. The mechanical design, selection of the belt material and working fluid, a preliminary test plan, and program plan are presented. The strategy used for selecting the basic sizes and materials of the components are discussed. Shuttle and crew member requirements are presented with some options for increasing or decreasing the demands on the STS. An STS carrier and the criteria used in the selection process are presented. The proposed carrier for the Moving Belt Radiator experiment is the Hitchhiker-M. Safety issues are also listed with possible results. This experiment is designed so that a belt can be deployed, run at steady state conditions, run with dynamic perturbations imposed, verify the operation of the interface heat exchanger and seals, and finally be retracted into a stowed position for transport back to earth.

Aguilar, Jerry L.↗