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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 NTRS1986

Sensitivity analysis and approximation methods for general eigenvalue problems

Optimization of dynamic systems involving complex non-hermitian matrices is often computationally expensive. Major contributors to the computational expense are the sensitivity analysis and reanalysis of a modified design. The present work seeks to alleviate this computational burden by identifying efficient sensitivity analysis and approximate reanalysis methods. For the algebraic eigenvalue problem involving non-hermitian matrices, algorithms for sensitivity analysis and approximate reanalysis are classified, compared and evaluated for efficiency and accuracy. Proper eigenvector normalization is discussed. An improved method for calculating derivatives of eigenvectors is proposed based on a more rational normalization condition and taking advantage of matrix sparsity. Important numerical aspects of this method are also discussed. To alleviate the problem of reanalysis, various approximation methods for eigenvalues are proposed and evaluated. Linear and quadratic approximations are based directly on the Taylor series. Several approximation methods are developed based on the generalized Rayleigh quotient for the eigenvalue problem. Approximation methods based on trace theorem give high accuracy without needing any derivatives. Operation counts for the computation of the approximations are given. General recommendations are made for the selection of appropriate approximation technique as a function of the matrix size, number of design variables, number of eigenvalues of interest and the number of design points at which approximation is sought.

Murthy, D. V.↗
NASA NTRS2001

Rock Driller

The next series of planetary exploration missions require a method of extracting rock and soil core samples. Therefore a prototype ultrasonic core driller (UTCD) was developed to meet the constraints of Small Bodies Exploration and Mars Sample Return Missions. The constraints in the design are size, weight, power, and axial loading. The ultrasonic transducer requires a relatively low axial load, which is one of the reasons this technology was chosen. The ultrasonic generator breadboard section can be contained within the 5x5x3 limits and weighs less than two pounds. Based on results attained the objectives for the first phase were achieved. A number of transducer probes were made and tested. One version only drills, and the other will actually provide a small core from a rock. Because of a more efficient transducer/probe, it will run at very low power (less than 5 Watts) and still drill/core. The prototype generator was built to allow for variation of all the performance-effecting elements of the transducer/probe/end effector, i.e., pulse, duty cycle, frequency, etc. The heart of the circuitry is what will be converted to a surface mounted board for the next phase, after all the parameters have been optimized and the microprocessor feedback can be installed.

Peterson, Thomas M.↗
NASA NTRSDate not supplied

A Cost and Benefit Analysis of Orbital Debris Remediation, Mitigation,Tracking, and Characterization

Orbital debris may collide with crewed and robotic spacecraft, placing them at risk. The wide range of debris, from 9,000-kilogram rocket bodies to millions of millimeter-size debris, has led to a similarly wide range of proposed actions for addressing the risks posed by debris. However, the costs and benefits of these actions have historically been unknown. This is a challenge for decision makers who are choosing which actions to support through technology development or policy changes. NASA’s Office of Technology, Policy, and Strategy is addressing these technical and economic uncertainties by building a capability to (1) complete rigorous calculations of the net present value of each action, (2) identify an optimal portfolio of actions to reduce risk, and (3) quantitatively analyze policies related to space sustainability. This report describes our progress toward that capability and to solicit feedback from the space and economic communities. Our previous work(Colvin, Karcz, and Wusk. 2023), referred to here as Phase 1, assessed the costs and benefits of performing debris remediation on operationally relevant timescales. The current paper summarizes the work of Locke and Colvin (2024), which contains major updates to the risk model used in Phase 1 and expands the breadth of actions considered to include mitigating the creation of debris, improving the ability to track debris, and more methods for cleaning up existing debris. We demonstrate that our approach of measuring risks in dollars allows for the effectiveness of seemingly incommensurate actions to be compared and generates insights that other approaches to measuring risk have missed.

Orbital Debris↗
NASA NTRS1993

Analysis of multigrid methods on massively parallel computers: Architectural implications

We study the potential performance of multigrid algorithms running on massively parallel computers with the intent of discovering whether presently envisioned machines will provide an efficient platform for such algorithms. We consider the domain parallel version of the standard V cycle algorithm on model problems, discretized using finite difference techniques in two and three dimensions on block structured grids of size 10(exp 6) and 10(exp 9), respectively. Our models of parallel computation were developed to reflect the computing characteristics of the current generation of massively parallel multicomputers. These models are based on an interconnection network of 256 to 16,384 message passing, 'workstation size' processors executing in an SPMD mode. The first model accomplishes interprocessor communications through a multistage permutation network. The communication cost is a logarithmic function which is similar to the costs in a variety of different topologies. The second model allows single stage communication costs only. Both models were designed with information provided by machine developers and utilize implementation derived parameters. With the medium grain parallelism of the current generation and the high fixed cost of an interprocessor communication, our analysis suggests an efficient implementation requires the machine to support the efficient transmission of long messages, (up to 1000 words) or the high initiation cost of a communication must be significantly reduced through an alternative optimization technique. Furthermore, with variable length message capability, our analysis suggests the low diameter multistage networks provide little or no advantage over a simple single stage communications network.

Matheson, Lesley R.↗
NASA NTRS2015

Evolving Design Criteria for Very Large Aperture Space-Based Telescopes and Their Influence on the Need for Integrated Tools in the Optimization Process

NASA's Advanced Mirror Technology Development (AMTD) program has been developing the means to design and build the future generations of space based telescopes. With the nearing completion of the James Webb Space Telescope (JWST), the astrophysics community is already starting to define the requirements for follow on observatories. The restrictions of available launch vehicles and the possibilities of planned future vehicles have fueled the competition between monolithic primaries (with better optical quality) and segmented primaries (with larger apertures, but with diffraction, costs and figure control issues). Regardless of the current shroud sizes and lift capacities, these competing architectures share the need for rapid design tools. As part of the AMTD program a number of tools have been developed and tested to speed up the design process. Starting with the Arnold Mirror Modeler (which creates Finite Element Models (FEM) for structural analysis) and now also feeds these models into thermal stability analyses. They share common file formats and interchangeable results. During the development of the program, numerous trade studies were created for 4 meter and 8 meter monolithic primaries, complete with support systems. Evaluation of these results has led to a better understanding of how the specification drives the results. This paper will show some of the early trade studies for typical specification requirements such as lowest mirror bending frequency and suspension system lowest frequency. The results use representative allowable stress values for each mirror substrate material and construction method and generic material properties. These studies lead to some interesting relationships between feasible designs and the realities of actually trying to build these mirrors. Much of the traditional specifications were developed for much smaller systems, where the mass and volume of the primary where a small portion of the overall satellite. JWST shows us that as the aperture grows, the primary takes up the majority of the mass and volume and the established rules need to be adjusted. For example, a small change in lowest frequency requirement can change the cost by millions of dollars.

Space-based telescopes↗
NASA NTRS2015

Evolving Design Criteria for Very Large Aperture Space Based Telescopes and Their Influence on the Need for Integrated Tools in the Optimization Process

NASA's Advanced Mirror Technology Development (AMTD) program has been developing the means to design and build the future generations of space based telescopes. With the nearing completion of the James Webb Space Telescope (JWST), the astrophysics community is already starting to define the requirements for follow-on observatories. The restrictions of available launch vehicles and the possibilities of planned future vehicles have fueled the competition between monolithic primaries (with better optical quality) and segmented primaries (with larger apertures, but with diffraction, costs and figure control issues). Regardless of the current shroud sizes and lift capacities, these competing architectures share the need for rapid design tools. As part of the AMTD program a number of tools have been developed and tested to speed up the design process. Starting with the Arnold Mirror Modeler (which creates Finite Element Models (FEM) for structural analysis) and now also feeds these models into thermal stability analyses. They share common file formats and interchangeable results. During the development of the program, numerous trade studies were created for 4-meter and 8-meter monolithic primaries, complete with support systems. Evaluation of these results has led to a better understanding of how the specification drives the results. This paper will show some of the early trade studies for typical specification requirements such as lowest mirror bending frequency and suspension system lowest frequency. The results use representative allowable stress values for each mirror substrate material and construction method and generic material properties. These studies lead to some interesting relationships between feasible designs and the realities of actually trying to build these mirrors. Much of the traditional specifications were developed for much smaller systems, where the mass and volume of the primary where a small portion of the overall satellite. JWST shows us that as the aperture grows, the primary takes up the majority of the mass and volume and the established rules need to be adjusted. For example, a small change in lowest frequency requirement can change the cost by millions of dollars. The paper uses numerous trade studies created during the software development phase of the Arnold Mirror Modeler to illustrate the influences of system specifications on the design space. The future telescopes will require better performance, stability and documented feasibility to meet the hurdles of today's budget and schedules realities. AMTD is developing the tools, but the basic system planning mentality also has to adopt to the requirements of these very large and complex physical structures.

Arnold, William R., Sr.↗
NASA NTRS2006

The Effectiveness of TAG or Guard-Gates in SET Suppression Using Delay and Dual-Rail Configurations at 0.35 microns

Design options for decreasing the susceptibility of integrated circuits to Single Event Upset (SEU) fall into two categories: (1) increasing the critical charge to cause an upset at a particular node, and (2) employing redundancy to mask or correct errors. With decreasing device sizes on an Integrated Circuit (IC), the amount of charge required to represent a logic state has steadily reduced. Critical charge methods such as increasing drive strength or increasing the time required to change state as in capacitive or resistive hardening or delay based approaches extract a steadily increasing penalty as a percentage of device resources and performance. Dual redundancy is commonly assumed only to provide error detection with Triple Modular Redundancy (TMR) required for correction, but less well known methods employ dual redundancy to achieve full error correction by voting two inputs with a prior state to resolve ambiguity. This requires special circuits such as the Whitaker latch [1], or the guard-gate [2] which some of us have called a Transition AND Gate (TAG) [3]. A 2-input guard gate is shown in Figure 1. It is similar to a Muller Completion Element [4] and relies on capacitance at node "out" to retain the prior state when inputs disagree, while eliminating any output buffer which would be susceptible to radiation strikes. This paper experimentally compares delay based and dual rail flip-flop designs wherein both types of circuits employ guard-gates to optimize layout and performance, and draws conclusions about design criteria and suitability of each option. In both cases a design goal is protection against Single Event Transients (SET) in combinational logic as well as SEU in the storage elements. For the delay based design, it is also a goal to allow asynchronous clear or preset inputs on the storage elements, which are often not available in radiation tolerant designs.

Shuler, Robert L.↗
NASA NTRS1980

Determination of the complex refractive index and size distribution of atmospheric particulates from bistatic-monostatic lidar and solar radiometer measurements

A method is presented for inferring both the size distribution and the complex refractive index of atmospheric particulates from combined bistatic-monostatic lidar and solar radiometer observations. The basic input measurements are spectral optical depths at several visible and near-infrared wavelengths as obtained with a solar radiometer and backscatter and angular scatter coefficients as obtained from a biostatic-monostatic lidar. The spectral optical depth measurements obtained from the radiometer are mathematically inverted to infer a columnar particulate size distribution. Advantage is taken of the fact that the shape of the size distribution obtained by inverting the particulate optical depth is relatively insensitive to the particle refractive index assumed in the inversion. Bistatic-monostatic angular scatter and backscatter lidar data are then processed to extract an optimum value for the particle refractive index subject to the constraint that the shape of the particulate size distribution be the same as that inferred from the solar radiometer data. Specifically, the scattering parameters obtained from the bistatic-monostatic lidar data are compared with corresponding theoretical computations made for various assumed refractive index values. That value which yields best agreement, in a weighted least squares sense, is selected as the optimal refractive index estimate. The results of this procedure applied to a set of simulated measurements as well as to measurements collected on two separate days are presented and discussed.

Reagan, J. A.↗
NASA NTRS2005

Advanced Method to Estimate Fuel Slosh Simulation Parameters

The nutation (wobble) of a spinning spacecraft in the presence of energy dissipation is a well-known problem in dynamics and is of particular concern for space missions. The nutation of a spacecraft spinning about its minor axis typically grows exponentially and the rate of growth is characterized by the Nutation Time Constant (NTC). For launch vehicles using spin-stabilized upper stages, fuel slosh in the spacecraft propellant tanks is usually the primary source of energy dissipation. For analytical prediction of the NTC this fuel slosh is commonly modeled using simple mechanical analogies such as pendulums or rigid rotors coupled to the spacecraft. Identifying model parameter values which adequately represent the sloshing dynamics is the most important step in obtaining an accurate NTC estimate. Analytic determination of the slosh model parameters has met with mixed success and is made even more difficult by the introduction of propellant management devices and elastomeric diaphragms. By subjecting full-sized fuel tanks with actual flight fuel loads to motion similar to that experienced in flight and measuring the forces experienced by the tanks these parameters can be determined experimentally. Currently, the identification of the model parameters is a laborious trial-and-error process in which the equations of motion for the mechanical analog are hand-derived, evaluated, and their results are compared with the experimental results. The proposed research is an effort to automate the process of identifying the parameters of the slosh model using a MATLAB/SimMechanics-based computer simulation of the experimental setup. Different parameter estimation and optimization approaches are evaluated and compared in order to arrive at a reliable and effective parameter identification process. To evaluate each parameter identification approach, a simple one-degree-of-freedom pendulum experiment is constructed and motion is induced using an electric motor. By applying the estimation approach to a simple, accurately modeled system, its effectiveness and accuracy can be evaluated. The same experimental setup can then be used with fluid-filled tanks to further evaluate the effectiveness of the process. Ultimately, the proven process can be applied to the full-sized spinning experimental setup to quickly and accurately determine the slosh model parameters for a particular spacecraft mission. Automating the parameter identification process will save time, allow more changes to be made to proposed designs, and lower the cost in the initial design stages.

Schlee, Keith↗
NASA NTRS1998

Mechanical Design of High Lift Systems for High Aspect Ratio Swept Wings

The NASA Ames Research Center is working to develop a methodology for the optimization and design of the high lift system for future subsonic airliners with the involvement of two partners. Aerodynamic analysis methods for two dimensional and three dimensional wing performance with flaps and slats deployed are being developed through a grant with the aeronautical department of the University of California Davis, and a flap and slat mechanism design procedure is being developed through a contract with PKCR, Inc., of Seattle, WA. This report documents the work that has been completed in the contract with PKCR on mechanism design. Flap mechanism designs have been completed for seven (7) different mechanisms with a total of twelve (12) different layouts all for a common single slotted flap configuration. The seven mechanisms are as follows: Simple Hinge, Upside Down/Upright Four Bar Linkage (two layouts), Upside Down Four Bar Linkages (three versions), Airbus A330/340 Link/Track Mechanism, Airbus A320 Link/Track Mechanism (two layouts), Boeing Link/Track Mechanism (two layouts), and Boeing 767 Hinged Beam Four Bar Linkage. In addition, a single layout has been made to investigate the growth potential from a single slotted flap to a vane/main double slotted flap using the Boeing Link/Track Mechanism. All layouts show Fowler motion and gap progression of the flap from stowed to a fully deployed position, and evaluations based on spanwise continuity, fairing size and number, complexity, reliability and maintainability and weight as well as Fowler motion and gap progression are presented. For slat design, the options have been limited to mechanisms for a shallow leading edge slat. Three (3) different layouts are presented for maximum slat angles of 20 deg, 15 deg and 1O deg all mechanized with a rack and pinion drive similar to that on the Boeing 757 airplane. Based on the work of Ljungstroem in Sweden, this type of slat design appears to shift the lift curve so that higher lift is achieved with the deployed slat with no increase in angle of attack. The layouts demonstrate that these slat systems can be designed with no need for slave links, and an experimental test program is outlined to experimentally validate the lift characteristics of the shallow slat.

Rudolph, Peter K. C.↗
NASA NTRSDate not supplied

Demonstration of How Manufacturing Innovations Challenge Conventional Structural Design

For almost 100 years, commercial aircraft have been fabricated using riveted aluminum alloy structures. Aside from refining aircraft designs for better aerodynamic efficiency, improving alloy compositions, and automating assembly steps, fuselage construction remains largely unchanged today. The intent of the lightweight metallic fuselage prototype undertaken in NASA’s Advanced Air Transport Technology project was to demonstrate innovative forming and joining processes that advance the design paradigm, i.e. achieve high-rate manufacturing and reduce assembly time/costs. Over the past decade, researchers at NASA Langley Research Center have evaluated the potential to modify the flow forming process to produce near-net shaped cylinders, integrally stiffened along the cylinder axis, for space launch vehicles. The production of integral stiffeners in a formed aluminum cylinder, the size of the Space Shuttle external tank, replaced machined thick plate and welding steps to eliminate > 500,000 pounds of machining chips and ~ 0.5 miles of welds. The flow forming method, called the Integrally Stiffened Cylinder (ISC) process, was successfully demonstrated up to the 10-foot diameter scale, which laid the groundwork for the current investigation of metallic fuselage structures. Using the ISC process as the basis for the re-design of a metallic fuselage eliminates hundreds of thousands of holes and rivets, while significantly reducing assembly time and crack initiation sites in the integral structure. However, the ISC process is not currently configured to fabricate circumferential ring frames for carrying fuselage internal pressure loads. Consequently, a trade study was performed to assess existing and advanced manufacturing processes, including additive manufacturing, forming, and welding. The approach with the lowest barriers to success, while simultaneously improving manufacturing time and cost, was to use formed ring frame segments attached to the ISC via Refill Friction Stir Spot Welding (RFSSW). The RFSSW process is five times faster than drilling, reaming, and riveting and provides similar mechanical performance. Finishing the fuselage structure with conventional windows, floor beams, and floor panels can then be accomplished using incumbent assembly methods, thereby maximizing reuse of existing infrastructure for aircraft construction. Structural analyses were performed to assess and optimize the geometric variables of integrated skin and stiffener configurations. A cost benefit and manufacturing rate analysis was also performed to compare against the current state-of-the-art for single aisle transport class aircraft. The resulting structure offers a weight reduction that rivals current graphite-epoxy composite fuselage structures. The projected manufacturing rate is close to double current metallic fuselages and six times faster than current composite manufacturing practices. The damage tolerance properties of a monocoque fuselage structure have yet to be assessed, but past integral airframe structural work has exploited geometric features to blunt or turn cracks. This concept offers promise that integrated structures can meet stringent aircraft durability specifications. An important benefit is that such aluminum fuselage structures may be inspected and repaired using established practices and existing expertise. Finally, pursuit of advanced manufacturing processes for future aluminum fuselages minimizes waste and the structure is 100% recyclable at the end-of-life for maximum sustainability.

aluminum↗
NASA NTRS2005

Buckling of Carbon Nanotube-Reinforced Polymer Laminated Composite Materials Subjected to Axial Compression and Shear Loadings

A multi-scale method to predict the stiffness and stability properties of carbon nanotube-reinforced laminates has been developed. This method is used in the prediction of the buckling behavior of laminated carbon nanotube-polyethylene composites formed by stacking layers of carbon nanotube-reinforced polymer with the nanotube alignment axes of each layer oriented in different directions. Linking of intrinsic, nanoscale-material definitions to finite scale-structural properties is achieved via a hierarchical approach in which the elastic properties of the reinforced layers are predicted by an equivalent continuum modeling technique. Solutions for infinitely long symmetrically laminated nanotube-reinforced laminates with simply-supported or clamped edges subjected to axial compression and shear loadings are presented. The study focuses on the influence of nanotube volume fraction, length, orientation, and functionalization on finite-scale laminate response. Results indicate that for the selected laminate configurations considered in this study, angle-ply laminates composed of aligned, non-functionalized carbon nanotube-reinforced lamina exhibit the greatest buckling resistance with 1% nanotube volume fraction of 450 nm uniformly-distributed carbon nanotubes. In addition, hybrid laminates were considered by varying either the volume fraction or nanotube length through-the-thickness of a quasi-isotropic laminate. The ratio of buckling load-to-nanotube weight percent for the hybrid laminates considered indicate the potential for increasing the buckling efficiency of nanotube-reinforced laminates by optimizing nanotube size and proportion with respect to laminate configuration.

Riddick, J. C.↗
NASA NTRS2015

Open-Source Conceptual Sizing Models for the Hyperloop Passenger Pod

Hyperloop is a new mode of transportation proposed as an alternative to California's high speed rail project, with the intended benefits of higher performance at lower overall costs. It consists of a passenger pod traveling through a tube under a light vacuum and suspended on air bearings. The pod travels up to transonic speeds resulting in a 35 minute travel time between the intended route from Los Angeles and San Francisco. Of the two variants outlined, the smaller system includes a 1.1 meter tall passenger capsule traveling through a 2.2 meter tube at 700 miles per hour. The passenger pod features water-based heat exchangers as well as an on-board compression system that reduces the aerodynamic drag as it moves through the tube. Although the original proposal looks very promising, it assumes that tube and pod dimensions are independently sizable without fully acknowledging the constraints of the compressor system on the pod geometry. This work focuses on the aerodynamic and thermodynamic interactions between the two largest systems; the tube and the pod. Using open-source toolsets, a new sizing method is developed based on one-dimensional thermodynamic relationships that accounts for the strong interactions between these sub-systems. These additional considerations require a tube nearly twice the size originally considered and limit the maximum pod travel speed to about 620 miles per hour. Although the results indicate that Hyperloop will need to be larger and slightly slower than originally intended, the estimated travel time only increases by approximately five minutes, so the overall performance is not dramatically affected. In addition, the proposed on-board heat exchanger is not an ideal solution to achieve reasonable equilibrium air temperatures within the tube. Removal of this subsystem represents a potential reduction in weight, energy requirements and complexity of the pod. In light of these finding, the core concept still remains a compelling possibility, although additional engineering and economic analyses are markedly necessary before a more complete design can be developed.

design analysis↗
NASA NTRS2010

From Shuttle Main Engine to the Human Heart: A Presentation to the Federal Lab Consortium for Technology Transfer

A NASA engineer received a heart transplant performed by Drs. DeBakey and Noon after suffering a serious heart attack. 6 months later that engineer returned to work at NASA determined to use space technology to help people with heart disease. A relationship between NASA and Drs. DeBakey and Noon was formed and the group worked to develop a low cost, low power implantable ventricular assist device (VAD). NASA patented the method to reduce pumping damage to red blood cells and the design of a continuous flow heart pump (#5,678,306 and #5,947,892). The technology and methodology were licensed exclusively to MicroMed Technology, Inc.. In late 1998 MicroMed received international quality and electronic certifications and began clinical trials in Europe. Ventricular assist devices were developed to bridge the gap between heart failure and transplant. Early devices were cumbersome, damaged red blood cells, and increased the risk of developing dangerous blood clots. Application emerged from NASA turbopump technology and computational fluid dynamics analysis capabilities. To develop the high performance required of the Space Shuttle main engines, NASA pushed the state of the art in the technology of turbopump design. NASA supercomputers and computational fluid dynamics software developed for use in the modeling analysis of fuel and oxidizer flow through rocket engines was used in the miniaturization and optimization of a very small heart pump. Approximately 5 million people worldwide suffer from chronic heart failure at a cost of 40 billion dollars In the US, more than 5000 people are on the transplant list and less than 3000 transplants are performed each year due to the lack of donors. The success of ventricular assist devices has led to an application as a therapeutic destination as well as a bridge to transplant. This success has been attributed to smaller size, improved efficiency, and reduced complications such as the formation of blood clots and infection.

Fogarty, Jennifer A.↗
NASA NTRS2016

Subscale Carbon-Carbon Nozzle Extension Development and Hot Fire Testing in Support of Upper Stage Liquid Rocket Engines

Upper stage and in-space liquid rocket engines are optimized for performance through the use of high area ratio nozzles to fully expand combustion gases to low exit pressures increasing exhaust velocities. Due to the large size of such nozzles and the related engine performance requirements, carbon-carbon (C/C) composite nozzle extensions are being considered for use in order to reduce weight impacts. NASA and industry partner Carbon-Carbon Advanced Technologies (C-CAT) are working towards advancing the technology readiness level of large-scale, domestically-fabricated, C/C nozzle extensions. These C/C extensions have the ability to reduce the overall costs of extensions relative to heritage metallic and composite extensions and to decrease weight by 50%. Material process and coating developments have advanced over the last several years, but hot fire testing to fully evaluate C/C nozzle extensions in relevant environments has been very limited. NASA and C-CAT have designed, fabricated and hot fire tested multiple subscale nozzle extension test articles of various C/C material systems, with the goal of assessing and advancing the manufacturability of these domestically producible materials as well as characterizing their performance when subjected to the typical environments found in a variety of liquid rocket and scramjet engines. Testing at the MSFC Test Stand 115 evaluated heritage and state-of-the-art C/C materials and coatings, demonstrating the capabilities of the high temperature materials and their fabrication methods. This paper discusses the design and fabrication of the 1.2k-lbf sized carbon-carbon nozzle extensions, provides an overview of the test campaign, presents results of the hot fire testing, and discusses potential follow-on development work.

Gradl, Paul↗
NASA NTRS2016

Successful Completion of the JWST OGSE2 Cryogenic Test at JSC Chamber-A While Managing Numerous Challenges

The James Webb Space Telescope (JWST) Optical Telescope Element (OTE) assembly is the largest optically stable infrared-optimized telescope currently being manufactured and assembled, and scheduled for launch in 2018. The JWST OTE, including the primary mirrors, secondary mirror, and the Aft Optics Subsystems (AOS) are designed to be passively cooled and operate at near 45 degrees Kelvin. Due to the size of its large sunshield in relation to existing test facilities, JWST cannot be optically or thermally tested as a complete observatory-level system at flight temperatures. As a result, the telescope portion along with its instrument complement will be tested as a single unit very late in the program, and on the program schedule critical path. To mitigate schedule risks, a set of 'pathfinder' cryogenic tests will be performed to reduce program risks by demonstrating the optical testing capabilities of the facility, characterizing telescope thermal performance, and allowing project personnel to learn valuable testing lessons off-line. This paper describes the 'pathfinder' cryogenic test program, focusing on the recently completed second test in the series called the Optical Ground Support Equipment 2 (OGSE2) test. The JWST OGSE2 was successfully completed within the allocated project schedule while faced with numerous conflicting thermal requirements during cool-down to the final cryogenic operational temperatures, and during warm-up after the cryo-stable optical tests. The challenges include developing a pre-test cool-down and warm-up profiles without a reliable method to predict the thermal behaviors in a rarified helium environment, and managing the test article hardware safety driven by the project Limits and Constraints (L&C's). Furthermore, OGSE2 test included the time critical Aft Optics Subsystem (AOS), a part of the flight Optical Telescope Element that would need to be placed back in the overall telescope assembly integrations. The OGSE2 test requirements included the strict adherence of the project contamination controls due to the presence of the contamination sensitive flight optical elements. The test operations required close coordination of numerous personnel while they being exposed and trained for the 'final' combined OTE and instrument cryo-test in 2017. This paper will also encompass the OGSE2 thermal data look-back review.

JWST OGSE2 Cryogenic Test↗
NASA NTRSDate not supplied

A Multiphysics Study to Improve Specific Energy of Primary Batteries for Low Temperature Operation for Deep Space Missions

Several lander missions on the outer planets such as Europa, Enceladus, and Titan require electrical power to operate scientific and communication equipment. The traditional power generation methods, such as a photovoltaic array, are not feasible as their efficiency drops significantly at these vast distances. The novel radioisotope power systems are not practical today based on current lander designs and the effectiveness of these systems. To perform in situ science on distant planets, a high specific energy battery (>700 Wh/kg) needs to operate for about 480 hours under cold temperatures (-40C or 0C) [1]. While a primary battery such as Li-CFx can provide high specific energy at room temperature, its specific capacity decreases significantly at low temperatures. One of the causes for this drop is low ion and electrical conductivity, and slower reaction kinetics. Slower transport and facile kinetics lead to an increase in the battery’s resistance and higher voltage drops during the cell operation, thus reducing specific capacity. Both the transport and kinetics show a strong dependence on temperature. Thus, a small temperature rise can lead to an increase in the reaction rate and ion conductivity; since the temperature, cell resistance, and specific capacity are interdependent. A conventional battery model accounts for ohmic, thermodynamic, and, electrochemical, and chemical decomposition heating. The ohmic heating can be controlled by designing a resistive microstructure and varying the ratios of the active materials [2]. The kinetics can be improved by increasing the surface area, reducing the particle size, or adding a catalyst. These parameters are often optimized to achieve high specific energy at room temperatures. A similar optimization study is not available at low temperatures and for a primary (high specific energy) battery. For this presentation, we will explore the effect of geometrical, microstructural, and material properties on optimal specific capacity at low temperatures through multiphysics simulations. The ion transport resistance depends on the porosity and the tortuosity of an electrode and the separator.

M. Mehta↗
NASA NTRS2009

Using Maximal Isometric Force to Determine the Optimal Load for Measuring Dynamic Muscle Power

Maximal power output occurs when subjects perform ballistic exercises using loads of ~30-50% of one-repetition maximum (1-RM). However, performing 1-RM testing prior to power measurement requires considerable time, especially when testing involves multiple exercises. Maximal isometric force (MIF), which requires substantially less time to measure than 1-RM, might be an acceptable alternative for determining the optimal load for power testing. PURPOSE: To determine the optimal load based on MIF for maximizing dynamic power output during leg press and bench press exercises. METHODS: Twenty healthy volunteers (12 men and 8 women; mean +/- SD age: 31+/-6 y; body mass: 72 +/- 15 kg) performed isometric leg press and bench press movements, during which MIF was measured using force plates. Subsequently, subjects performed ballistic leg press and bench press exercises using loads corresponding to 20%, 30%, 40%, 50%, and 60% of MIF presented in randomized order. Maximal instantaneous power was calculated during the ballistic exercise tests using force plates and position transducers. Repeated-measures ANOVA and Fisher LSD post hoc tests were used to determine the load(s) that elicited maximal power output. RESULTS: For the leg press power test, six subjects were unable to be tested at 20% and 30% MIF because these loads were less than the lightest possible load (i.e., the weight of the unloaded leg press sled assembly [31.4 kg]). For the bench press power test, five subjects were unable to be tested at 20% MIF because these loads were less than the weight of the unloaded aluminum bar (i.e., 11.4 kg). Therefore, these loads were excluded from analysis. A trend (p = 0.07) for a main effect of load existed for the leg press exercise, indicating that the 40% MIF load tended to elicit greater power output than the 60% MIF load (effect size = 0.38). A significant (p . 0.05) main effect of load existed for the bench press exercise; post hoc analysis indicated that the effect of load on power output was: 30% > 40% > 50% = 60%. CONCLUSION: Loads of 40% and 30% of MIF elicit maximal power output during dynamic leg presses and bench presses, respectively. These findings are similar to those obtained when loading is based on 1-RM.

Spiering, Barry A.↗