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NASA NTRS1985

Cell wall elasticity: I. A critique of the bulk elastic modulus approach and an analysis using polymer elastic principles

The traditional bulk elastic modulus approach to plant cell pressure-volume relations is inconsistent with its definition. The relationship between the bulk modulus and Young's modulus that forms the basis of their usual application to cell pressure-volume properties is demonstrated to be physically meaningless. The bulk modulus describes stress/strain relations of solid, homogeneous bodies undergoing small deformations, whereas the plant cell is best described as a thin-shelled, fluid-filled structure with a polymer base. Because cell walls possess a polymer structure, an alternative method of mechanical analysis is presented using polymer elasticity principles. This initial study presents the groundwork of polymer mechanics as would be applied to cell walls and discusses how the matrix and microfibrillar network induce nonlinear stress/strain relationships in the cell wall in response to turgor pressure. In subsequent studies, these concepts will be expanded to include anisotropic expansion as regulated by the microfibrillar network.

Plants/chemistry/cytology/ultrastructure↗

NASA NTRS2016

A First Look at the Upcoming SISO Space Reference FOM

Spaceflight is difficult, dangerous and expensive; human spaceflight even more so. In order to mitigate some of the danger and expense, professionals in the space domain have relied, and continue to rely, on computer simulation. Simulation is used at every level including concept, design, analysis, construction, testing, training and ultimately flight. As space systems have grown more complex, new simulation technologies have been developed, adopted and applied. Distributed simulation is one those technologies. Distributed simulation provides a base technology for segmenting these complex space systems into smaller, and usually simpler, component systems or subsystems. This segmentation also supports the separation of responsibilities between participating organizations. This segmentation is particularly useful for complex space systems like the International Space Station (ISS), which is composed of many elements from many nations along with visiting vehicles from many nations. This is likely to be the case for future human space exploration activities. Over the years, a number of distributed simulations have been built within the space domain. While many use the High Level Architecture (HLA) to provide the infrastructure for interoperability, HLA without a Federation Object Model (FOM) is insufficient by itself to insure interoperability. As a result, the Simulation Interoperability Standards Organization (SISO) is developing a Space Reference FOM. The Space Reference FOM Product Development Group is composed of members from several countries. They contribute experiences from projects within NASA, ESA and other organizations and represent government, academia and industry. The initial version of the Space Reference FOM is focusing on time and space and will provide the following: (i) a flexible positioning system using reference frames for arbitrary bodies in space, (ii) a naming conventions for well-known reference frames, (iii) definitions of common time scales, (iv) federation agreements for common types of time management with focus on time stepped simulation, and (v) support for physical entities, such as space vehicles and astronauts. The Space Reference FOM is expected to make collaboration politically, contractually and technically easier. It is also expected to make collaboration easier to manage and extend.

Mueller, Bjorn↗

NASA NTRSDate not supplied

Preliminary Medical Risk Estimates and Clinical Capability Needs for Artemis IV

BACKGROUND Human exploration spaceflight missions to the Moon and Mars present unprecedented challenges for in- mission medical care. Compared with the ISS, the greater distance from Earth will mean increased mission durations, communication delays, limited to no resupply opportunities, and significant limitations on the evacuation of ill or injured crew. Spacecraft mass, volume, and power will be curtailed while higher demands will be placed on the crew’s knowledge, skills, and abilities. In this higher risk environment, it is important to: a) quantitatively estimate human system risk attributable to medical conditions, a process known as Probabilistic Risk Analysis, and b) use these estimates to inform medical system design. IMPACT (Informing Mission Planning via Analysis of Complex Tradespaces) is a PRA and medical trade space analysis tool developed by NASA to advance exploration mission medical system design. IMPACT v1.0 improves upon and will soon replace NASA’s existing tool, the Integrated Medical Model, with: a novel evidence base baselined to exploration environments; an expanded list of 119 medical conditions; a significant increase in the number of medical resources that can be utilized and in the flexibility of their use; and the modelling of time lost performing mission-specific tasks due to medical conditions. METHODOLOGY: This abstract will present IMPACT estimates of medical system risk and clinical capability needs for the Artemis IV mission. Artemis IV is currently scheduled for 2026 and will visit the Gateway space station in lunar orbit prior to the second lunar landing of the Artemis program. The baseline Artemis IV mission that was modeled was 28 days in duration with phases including Orion outbound, 4 days on the Gateway space station in lunar orbit, 2 crew on the surface of the Moon for approximately one week, an additional 5 days on Gateway, and then return to Earth. This baseline was compared to two alternative 34-day design reference missions (DRMs) that shifted the lunar sortie earlier or later in the mission profile. Assumptions included 2 female and 2 male crew and a notional medical system mass of 25 kg. Medical system risk estimates include loss of crew life (LOCL), consideration of medical evacuation (known as return to definitive care – RTDC), and an estimate of crew time lost due to medical conditions (Task Time Lost – TTL). The presentation will also describe the medical conditions that are the greatest drivers of risk as well as the clinical capabilities and resources that have the largest effect on risk. RESULTS: All three DRMs had very low probability of LOCL from medical conditions, primarily due to short duration. RTDC was also similar across the DRMs. In contrast, TTL was higher in the early lunar sortie DRM due to earlier occurrence of EVA-related medical conditions. Taken as a whole, there was no clinically significant difference in medical risk across the three missions. Results for clinical capabilities and an example medical equipment list will be discussed but were similar across DRMs.

D Hilmers↗

NASA NTRS2024

Preliminary Medical Risk Estimates and Clinical Capability Needs for Late Artemis Missions

D. Hilmers↗

NASA NTRS2011

Micrometeoroid and Orbital Debris Threat Assessment: Mars Sample Return Earth Entry Vehicle

This report provides results of a Micrometeoroid and Orbital Debris (MMOD) risk assessment of the Mars Sample Return Earth Entry Vehicle (MSR EEV). The assessment was performed using standard risk assessment methodology illustrated in Figure 1-1. Central to the process is the Bumper risk assessment code (Figure 1-2), which calculates the critical penetration risk based on geometry, shielding configurations and flight parameters. The assessment process begins by building a finite element model (FEM) of the spacecraft, which defines the size and shape of the spacecraft as well as the locations of the various shielding configurations. This model is built using the NX I-deas software package from Siemens PLM Software. The FEM is constructed using triangular and quadrilateral elements that define the outer shell of the spacecraft. Bumper-II uses the model file to determine the geometry of the spacecraft for the analysis. The next step of the process is to identify the ballistic limit characteristics for the various shield types. These ballistic limits define the critical size particle that will penetrate a shield at a given impact angle and impact velocity. When the finite element model is built, each individual element is assigned a property identifier (PID) to act as an index for its shielding properties. Using the ballistic limit equations (BLEs) built into the Bumper-II code, the shield characteristics are defined for each and every PID in the model. The final stage of the analysis is to determine the probability of no penetration (PNP) on the spacecraft. This is done using the micrometeoroid and orbital debris environment definitions that are built into the Bumper-II code. These engineering models take into account orbit inclination, altitude, attitude and analysis date in order to predict an impacting particle flux on the spacecraft. Using the geometry and shielding characteristics previously defined for the spacecraft and combining that information with the environment model calculations, the Bumper-II code calculates a probability of no penetration for the spacecraft.

Christiansen, Eric L.↗

NASA NTRS2024

Discrete Event Simulation-Based Timeline Validation Using R2U2

The Gateway Vehicle Systems Manager (VSM), the top-level software control system in a distributed, hierarchical Autonomous System Management Architecture is, like most modern spacecraft software control systems, heavily data-driven. For example, schedules (timelines) will be developed on the ground and, due to the high degree of autonomy, contain complex procedures involving conditional branching, variable timing, and resource contention resolution. In order to verify that an uploaded timeline will function correctly, it is necessary to explore the feasible set of possible executions. While it is possible to test a timeline using a mission simulation, the complexity of the system and duration of a timeline limits the number of trials and therefore the test coverage. To address this problem, the VSM team is using a discrete event system model that can rapidly generate from a timeline sets of event sequences using Monte Carlo techniques. To achieve rapid and trustworthy checking of the event sequences, we use an offline version of the runtime model checking tool R2U2. This presentation describes the approach the VSM team is using to implement the discrete event simulation and evaluate event sequences using R2U2. The presentation will discuss: 1. Description of the timelines by VSM in the context of VSM operations 2. Expansion of a timeline into a sequence of atomic events 3. Adjustment, in the Monte Carlo environment, of an event sequence to account for uncertainty, external events, and failures 4. Definition of R2U2 input and mission-time linear temporal logic files 5. Generation and use of R2U2 verdict sequences 6. Lessons learned and future work

Verification↗

NASA NTRS1991

Contact dynamics testing of automated three-point docking mechanism

An evaluation of an OMV docking mechanism, based on an adaptation of the Shuttle Flight Support System Pallet Berthing Mechanism was completed. The mechanism uses automatically actuated motorized latches to engage towel bars on the target satellite. LED sensors establish the towel bar position within the capture envelope and the latch capture commands are issued. Then, locking pawls engage the bar, locking and pre-loading the mechanism. Two series of tests were conducted to test nominal and failure mode captures and to evaluate design parameters such as LED sensor locations, automatic closure algorithms, latch closure velocity, position/velocity entry envelopes, and closure method. The first test series involved single latch testing on the Flat Floor Facility, the 6 DOF Facility, and an analytic simulation model. The intent was to compare results in order to validate the various facilities. Reasonably good agreement was achieved. The second test series repeated the single latch testing on the refurbished 6 DOF Facility to validate the facility modifications. The individual latches were tested under free-drift conditions for functionality and performance. Next, the three-latch configuration underwent parametric testing. Test results validated the improved fidelity of the 6 DOF Facility and verified successful docking at the required entry velocity. The tests determined the 'best' design parameter definitions and concluded that the locking pawls should not lock until all three latches completely close.

Rourke, Kenneth H.↗

NASA NTRS2002

Second Generation RLV Space Vehicle Concept

NASA has a long history of conducting development programs and projects in a consistent fashion. Systems Engineering within those programs and projects has also followed a given method outlined by such documents as the NASA Systems Engineering Handbook. The relatively new NASA Space Launch Initiative (SLI) is taking a new approach to developing a space vehicle, with innovative management methods as well as new Systems Engineering processes. With the program less than a year into its life cycle, the efficacy of these new processes has yet to be proven or disproven. At $776M for phase 1, SLI represents a major portion of the NASA focus; however, the new processes being incorporated are not reflected in the training provided by NASA to its engineers. The NASA Academy of Program and Project Leadership (APPL) offers core classes in program and project management and systems engineering to NASA employees with the purpose of creating a "knowledge community where ideas, skills, and experiences are exchanged to increase each other's capacity for strong leadership". The SLI program is, in one sense, a combination of a conceptual design program and a technology program. The program as a whole doesn't map into the generic systems engineering project cycle as currently, and for some time, taught. For example, the NASA APPL Systems Engineering training course teaches that the "first step in developing an architecture is to define the external boundaries of the system", which will require definition of the interfaces with other systems and the next step will be to "define all the components that make up the next lower level of the system hierarchy" where fundamental requirements are allocated to each component. Whereas, the SLI technology risk reduction approach develops architecture subsystem technologies prior to developing architectures. The higher level architecture requirements are not allowed to fully develop and undergo decomposition and allocation down to the subsystems before the subsystems must develop allocated requirements based on the highest level of requirements. In the vernacular of the project cycles prior to the mid 1990's, the architecture definition portion of the program appears to be at a generic Phase A stage, while the subsystems are operating at Phase B. Even the management structure of the SLI program is innovative in its approach to Systems Engineering and is not reflected in the APPL training modules. The SLI program has established a Systems Engineering office as an office separate from the architecture development or the subsystem technology development, while that office does have representatives within these other offices. The distributed resources of the Systems Engineering Office are co-located with the respective Project Offices. This template is intended to provide systems engineering as an integrated function at the Program Level. the program management of SLI and the MAT agree that "program/project managers and the systems engineering team must work closely together towards the single objective of delivering quality products that meet the customer needs". This paper will explore the differences between the methods being taught by NASA, which represent decades of ideas, and those currently in practice in SLI. Time will tell if the innovation employed by SLI will prove to be the model of the future. For now, it is suggested that the training of the present exercise the flexibility of recognizing the new processes employed by a major new NASA program.

Bailey, Michelle↗

NASA NTRS2005

The Mega Mesospheric Parachute

The current understanding and modeling of the upper reaches of the atmosphere is incomplete. Upper atmospheric interactions with the lower atmosphere, effects of ionizing radiation, high altitude cloud phenomena, and the dynamical interaction with the magnetosphere require greater definition. The scientific objective of obtaining a greater understanding of the upper atmosphere can be achieved by designing, implementing, testing, and utilizing a facility that provides long period in-situ measurements of the mesosphere. Current direct sub-sonic measurements of the upper atmosphere are hampered by the approximately one minute sub-sonic observation window of a ballistic sounding rocket regardless of the launch angle. In-situ measurements at greater than transonic speeds impart energy into the molecular atmospheric system and distort the true atmospheric chemistry. A long duration, sub-sonic capability will significantly enhance our ability to observe and measure: (1) mesospheric lightning phenomena (sprites and blue jets) (2) composition, structure and stratification of noctilucent clouds (3) physics of seasonal radar echoes, gravity wave phenomena (4) chemistry of mesospheric gaseous ratio mixing (5) mesospheric interaction of ionizing radiation (6) dynamic electric and magnetic fields This new facility will also provide local field measurements which complement those that can be obtained through external measurements from satellite and ground-based platforms. The 400 foot (approximately 130 meter) diameter lightweight mega-mesospheric parachute system, deployed with a sounding rocket, is proposed herein as a method to increase sub-sonic mesospheric measurement time periods by more than an order of magnitude. The report outlines a multi-year evolving science instrumentation suite in parallel with the development of the mega meso-chute facility. The developmental issues surrounding the meso-chute are chiefly materials selection (thermal and structural) and deployment mechanism physics. Three mission cases were conceived and developed to include cost and schedules estimates. Each scenario has increasing scientific utility with paralleling launch weight, parachute hang-time, deployment altitude, and parachute size: (1) Case #1: $8.4M@24 months, 6kg payload, 20 min., 50km alt., 80 m. dia. (2) Case #2: $10.4M@24 months, 6kg payload, 20 min., 60km alt, 130m. dia. (3) Case #3: $13.6M@36 months, 30kg payload, 30 min., 90km alt., 200m. dia. The initial breakout cost for the parachute system is approximately $2M@24 months. This report identifies that although the challenges of the mega-meso-chute may be difficult, they can be surmounted and valuable results can be achieved.

Kloesel, Kurt J.↗

NASA NTRS1990

On-orbit assembly/servicing task definition study

The OEXP vehicles being envisioned to carry out the Presidential space goals of a lunar outpost and human exploration of Mars will require on-orbit assembly, refurbishment, checkout, and launch. The On-orbit Assembly/Servicing Task Definition Study applies the space vehicle processing experience and procedures archives resident at NASA's Kennedy Space Center (KSC) to determine the task flows, and resources/facilities necessary to process the OEXP vehicles at Space Station Freedom (SSF). This data base is examined to find the closest analogies to OEXP vehicle components and assembly/refurbishment tasks. Transition tables are generated to provide traceability from KSC hardware processing experience to analogous on-orbit processing of the OEXP vehicles. Iterations in which the task flows are broken down into realistic extravehicular activity (EVA) primitive subtasks and times, and to apply automation and robotic technology to reduce crew risks and minimize EVA time, will enhance the value and accuracy of the predicted flows. These processing scenarios and the resulting resource/facility requirements are used to determine impacts of SSF, resulting in change requests to SSF requirements for provision of 'hooks and scars' to evolve the assembly complete Space Station into a transportation node. Study results to date include assembly analysis of the Martin Marietta Phobos Gateway Vehicle, refurbishment analysis of the Martin Marietta Lunar Evolution Piloted and Cargo Vehicles, and assembly analysis of the Boeing Mars Evolution Vehicle. The results of this study will be accumulated into the vehicle processing operations data base for subsequent modeling, life cycle cost, vehicle growth, and SSF impact analysis.

Vargo, Rick↗

NASA NTRS1995

The local metallicity-surface brightness relationship in galactic disks

We present the results of a first attempt to employ multiaperture masks to obtain spectrophotometry of H II regions in nearby galaxies. A total of 97 H II regions in six southern spiral galaxies were observed using a combination of multiaperture masks and conventional long-slit spectrophotometry. The oxygen abundances derived from the multiaperture mask observations using the empirical abundance diagnostic R(sub 23) are shown to be consistent with those from long-slit spectra and generally show better reproducibility and object definition. Although the number of objects that can be observed simultaneously with this particular system is still quite limited compared with either imaging spectrophotometry or fiber-fed spectrographs, the spectral resolution offered and high throughput in the blue help make multiaperture spectrophotometry a competitive technique for increasing the sampling of H II regions in both radial distance and luminosity. There is still no clear trend of abundance gradient with either the galaxy's luminosity or its Hubble type, although the extrapolated central abundance does appear to correlate with galaxy luminosity/mass. In order to avoid difficulty in choosing an appropriate normalizing radius, we instead plot the oxygen abundance against the underlying I-band surface brightness at the radial distance of the H II region and confirm the existence of a local metallicity-surface brightness reltaionship within the disks of spiral galaxies. Although the simple closed-boc model of galaxy evolution predicts almost the right form of this relationship, a more realistic multizone model employing expnentially decreasing gas infall provides a more satisfactory fit to the observational data, provided the expected enriched gas return from dying low-mass stars shedding their envelopes at late epochs is properly taken into account. This same model, with a star formation law based upon self-regulating star formation in a three-dimensional disk (Dopita & Ryder 1994), is equally capable of accounting for the observed relationship between recent massive star formation and stellar surface brightness (Ryder & Dopita 1994).

Ryder, Stuart D.↗

NASA NTRS1986

Computer models of the spacecraft wake

Until recently, computations of space plasma flow over a spacecraft have been unstable for ratios of spacecraft dimension to Debye length typical of the low Earth orbit environment. Calculations are presented of the spacecraft/environment interaction based on two computer codes, MACH and POLAR. MACH, an inside-out particle tracking code, was developed for the purpose of validating the physics of POLAR in regimes where these are no comprehensive theoretical or experimental results. While the spacecraft which can be treated by MACH are restricted to simple geometries, the methodology is more fundamental than POLAR. MACH generates self-consistent solutions within the context of quasisteady Vlasov plasma flow and achieves Debye ratios previously unobtainable. POLAR uses a three-dimensional finite-element representation of the vehicle in a staggered mesh. The plasma sheath is modeled by outside-in particle tracking. Solutions for the plasma flow, wake and vehicle charging are obtained by Vlasov-Poisson iteration; charge stabilization techniques make the results virtually insensitive to the Debye ratio. POLAR reproduces the Laframboise static plasma solutions for sperical probes and fits the Makita-Kuriki probe data for spheres in a flowing plasma in regions where comparisons are valid. POLAR and MACH solutions for the particle and electrostatic potential structure of the wake of a charged disk in a low-altitude flow are shown for Mach numbers 4, 5, and 8. New features of the solutions include ion focussing in the wake and a definitive determination of the sheath edge in the wake which shows that the sheath is not an equipotential.

Rubin, A. G.↗

NASA NTRS1999

Improving Scientific Capabilities in Space in the 21st Century: The NASA New Millennium Program

NASA's New Millennium Program (NMP) has been chartered to identify and validate in space emerging, revolutionary technologies that will enable less costly, more capable future science missions. The program utilizes a unique blend of science guidance and industry partnering to ferret out technology solutions to enable science capabilities in space which are presently technically infeasible, or unaffordable. Those technologies which present an unacceptably high risk to future science missions (whether small PI-led or operational) are bundled into technology validation missions. These missions seek to validate the technologies in a manner consistent with their future uses, thus reducing the associated risk to the first user, and obtaining meaningful science data as well. The Space Readiness Coherent Lidar Experiment (SPARCLE) was approved as the second NMP Earth Observing mission (EO2) in October 1997, and assigned to Marshall Space Flight Center for implementation. Leading up to mission confirmation, NMP sponsored a community workshop in March 1996 to draft Level-1 requirements for a doppler wind lidar mission, as well as other space-based lidar missions (such as DIAL). Subsequently, a study group was formed and met twice to make recommendations on how to perform a comparison of coherent and direct detection wind lidars in space. These recommendations have guided the science validation plan for the SPARCLE mission, and will ensure that future users will be able to confidently assess the risk profile of future doppler wind missions utilizing EO2 technologies. The primary risks to be retired are: (1) Maintenance of optical alignments through launch and operations on orbit, and (2) Successful velocity estimation compensation for the Doppler shift due to the platform motion, and due to the earth's rotation. This includes the need to account for all sources of error associated with pointing control and knowledge. The validation objectives are: (1) Demonstrate measurement of tropospheric winds from space using a scanning coherent Doppler lidar technique that scales to meet future research (e.g. ESSP) and operational (e.g. NPOESS) mission requirements. Specifically, produce and validate LOS wind data with single shot accuracy of 1-2 m/s in regions of high signal-to-noise ratio (SNR), and low atmospheric wind turbulence and wind shear, (2) Collect the atmospheric and instrument performance data in various scanning modes necessary to validate and improve instrument performance models that will enable the definition of future missions with greater confidence. Such data include aerosol backscatter data over much of the globe, and high SNR data such as that from surface returns, and (3) Produce a set of raw instrument data with which advanced signal processing techniques can be developed. This objective will permit future missions to better understand how to extract wind information from low backscatter regions of the atmosphere.

Raymond, Carol A.↗

NASA NTRS2016

Carbon Solubility in Silicon-Iron-Bearing Metals during Core Formation on Mercury

Recent results obtained from the MErcury Surface, Space ENvironment, GEochemistry, and Ranging (MESSENGER) spacecraft showed the surface of Mercury has high S abundances (approximately 4 wt%) and low Iron(II) Oxide abundances (less than 2 wt%). Based on these extreme values, the oxygen fugacity of Mercury's surface materials was estimated to be approximately 3 to 7 log(sub 10) units below the IW buffer (Delta IW-3 to Delta IW-7). This highly reducing nature of the planet has resulted in a large core and relatively thin mantle, extending to only approximately 420 km depth (corresponding to a core-mantle boundary pressure of approximately 4-7 GPa) within the planet. Furthermore, MESSENGER results have suggested the presence of carbon on the surface of the planet. Previous experimental results from have also suggested the possibility of a primary floatation crust on Mercury composed of graphite, produced after a global magma ocean event. With these exotic conditions of this compositional end-member planet, it begs the question, what is the core composition of Mercury? Although no definitive conclusion has been reached, previous studies have made advances towards answering this question. Riner et al. and Chen et al. looked at iron sulfide systems and implemented various crystallization and layered core scenarios to try and determine the composition and structure of Mercury's core. Malavergne et al. examined core crystallization scenarios in the presence of sulfur and silicon. Hauck et al. used the most recent geophysical constraints from the MESSENGER spacecraft to model the internal structure of Mercury, including the core, in a iron-sulfur-silicon system. More recently, Chabot et al. conducted a series of metal-silicate partitioning experiments in a iron-sulfur-silicon system. These results showed the core of Mercury has the potential to contain more than 15 wt% silicon. However, with the newest results from MESSENGER's low altitude campaign, carbon is another potential light element that could be incorporated into Mercury's core. The goal of this study is to determine the carbon concentration at graphite saturation in various silicon-iron bearing metals relevant to possible mercurian core compositions. Future experiments will include the addition of sulfur into these metals.

Vander Kaaden, Kathleen E.↗

NASA NTRS2000

GCSS Cirrus Parcel Model Comparison Project

The Cirrus Parcel Model Comparison Project, a project of GCSS Working Group on Cirrus Cloud Systems (WG2), involves the systematic comparison of current models of ice crystal nucleation and growth for specified, typical, cirrus cloud environments. The goal of this project is to document and understand the factors resulting in significant inter-model differences. The intent is to foment research leading to model improvement and validation. In Phase 1 of the project reported here, simulated cirrus cloud microphysical properties are compared for situations of "warm" (-40 C) and "cold" (-60 C) cirrus subject to updrafts of 4, 20 and 100 cm/s, respectively. Five models participated. These models employ explicit microphysical schemes wherein the size distribution of each class of particles (aerosols and ice crystals) is resolved into bins. Simulations are made including both homogeneous and heterogeneous ice nucleation mechanisms. A single initial aerosol population of sulfuric acid particles is prescribed for all simulations. To isolate the treatment of the homogeneous freezing (of haze drops) nucleation process, the heterogeneous nucleation mechanism is disabled for a second parallel set of simulations. Qualitative agreement is found for the homogeneous-nucleation-only simulations, e.g., the number density of nucleated ice crystals increases with the strength of the prescribed updraft. However, non-negligible quantitative differences are found. Detailed analysis reveals that the homogeneous nucleation formulation, aerosol size, ice crystal growth rate (particularly the deposition coefficient), and water vapor uptake rate are critical components that lead to differences in predicted microphysics. Systematic bias exists between results based on a modified classical theory approach and models using an effective freezing temperature approach to the treatment of nucleation. Each approach is constrained by critical freezing data from laboratory studies, but each includes assumptions that can only be justified by further laboratory data. Consequently, it is not yet clear if the two approaches can be made consistent. Large haze particles may deviate considerably from equilibrium size in moderate to strong updrafts (20-100 cm/s) at -60 C when the commonly invoked equilibrium assumption is lifted. The resulting difference in particle-size-dependent solution concentration of haze particles may significantly affect the ice nucleation rate during the initial nucleation interval. The uptake rate for water vapor excess by ice crystals is another key component regulating the total number of nucleated ice crystals. This rate, the product of ice number concentration and ice crystal diffusional growth rate, which is sensitive to the deposition coefficient when ice particles are small, partially controls the peak nucleation rate achieved in an air parcel and the duration of the active nucleation time period. The effects of heterogeneous nucleation are most pronounced in weak updraft situations. Vapor competition by the nucleated (heterogeneous) ice crystals limits the achieved ice supersaturation and thus suppresses the contribution of homogeneous nucleation. Correspondingly, ice crystal number density is markedly reduced. Definitive laboratory and atmospheric benchmark data are needed for the heterogeneous nucleation process. Inter-model differences are correspondingly greater than in the case of the homogeneous nucleation process acting alone.

Lin, Ruei-Fong↗

NASA NTRS1995

Numerical thermal analyses of heat exchangers for the stirling engine application

The Regenerator, Cooler and Heater for the NASA Space Power Research Engine (SPRE) have been analyzed in detail for laminar, incompressible and oscillatory flow conditions. Each component has been analyzed independently and in detail with the regenerator being modeled as two-parallel-plates channel with a solid wall. The ends of the channel are exposed to two reservoir maintained at different temperature thus providing an axial temperature gradient along the channel. The cooler and heater components have been modeled as circular pipes with isothermal walls. Two different types of thermal boundary conditions have been investigated for the cooler and heater, namely, symmetric and asymmetric temperature inflow. In symmetric temperature inflow the flow enters the channel with the same temperature in throughout the velocity cycle whereas, in asymmetric temperature inflow the flow enters with a different temperature in each half cycle. The study was conducted over a wide range of Maximum Reynolds number (RE(max) varying from 75 to 60000, Valensi number (Va) from 2.5 to 800, and relative amplitude of fluid displacement (A(sub r) from 0.357 to 1.34. A two dimensional Finite volume method based on the SIMPLE algorithm was used to solve the governing partial differential equations. Post processing programs were developed to effectively describe the heat transfer mechanism under oscillatory flows. The computer code was validated by comparing with existing analytical solutions for oscillating flows. The thermal field have been studied with the help of temperature contour and three dimensional plots. The instantaneous friction factor, wall heat flux and heat transfer coefficient have been examined. It has been concluded that in general, the frictional factor and heat transfer coefficient are higher under oscillatory flow conditions when the Valensi number is high. Also, the thermal efficiency decreases for lower A(r) values. Further, the usual steady state definition for the heat transfer coefficient does not seem to be valid.

Kannapareddy, Mohan Raj↗

NASA NTRS2015

The Use of UML for Software Requirements Expression and Management

It is common practice to write English-language "shall" statements to embody detailed software requirements in aerospace software applications. This paper explores the use of the UML language as a replacement for the English language for this purpose. Among the advantages offered by the Unified Modeling Language (UML) is a high degree of clarity and precision in the expression of domain concepts as well as architecture and design. Can this quality of UML be exploited for the definition of software requirements? While expressing logical behavior, interface characteristics, timeliness constraints, and other constraints on software using UML is commonly done and relatively straight-forward, achieving the additional aspects of the expression and management of software requirements that stakeholders expect, especially traceability, is far less so. These other characteristics, concerned with auditing and quality control, include the ability to trace a requirement to a parent requirement (which may well be an English "shall" statement), to trace a requirement to verification activities or scenarios which verify that requirement, and to trace a requirement to elements of the software design which implement that requirement. UML Use Cases, designed for capturing requirements, have not always been satisfactory. Some applications of them simply use the Use Case model element as a repository for English requirement statements. Other applications of Use Cases, in which Use Cases are incorporated into behavioral diagrams that successfully communicate the behaviors and constraints required of the software, do indeed take advantage of UML's clarity, but not in ways that support the traceability features mentioned above. Our approach uses the Stereotype construct of UML to precisely identify elements of UML constructs, especially behaviors such as State Machines and Activities, as requirements, and also to achieve the necessary mapping capabilities. We describe this approach in the context of a space-based software application currently under development at the Jet Propulsion Laboratory.

model-based engineering↗

NASA NTRS2008

Radiative Susceptibility of Cloudy Atmospheres to Droplet Number Perturbations: 2. Global analysis from MODIS

Global distributions of albedo susceptibility for areas covered by liquid clouds are presented for 4 months in 2005. The susceptibility estimates are based on expanded definitions presented in a companion paper and include relative cloud droplet number concentration (CDNC) changes, perturbations in cloud droplet asymmetry parameter and single-scattering albedo, atmospheric/surface effects, and incorporation of the full solar spectrum. The cloud properties (optical thickness and effective radius) used as input in the susceptibility calculations come from MODIS Terra and Aqua Collection 5 gridded data. Geographical distributions of susceptibility corresponding to absolute ( absolute cloud susceptibility ) and relative ( relative cloud susceptibility ) CDNC changes are markedly different indicating that the detailed nature of the cloud microphysical perturbation is important for determining the radiative forcing associated with the first indirect aerosol effect. However, both types of susceptibility exhibit common characteristics such as significant reductions when perturbations in single-scattering properties are omitted, significant increases when atmospheric absorption and surface albedo effects are ignored, and the tendency to decrease with latitude, to be higher over ocean than over land, and to be statistically similar between the morning and afternoon MODIS overpasses. The satellite-based susceptibility analysis helps elucidate the role of present-day cloud and land surface properties in indirect aerosol forcing responses. Our realistic yet moderate CDNC perturbations yield forcings on the order of 1-2 W/sq m for cloud optical property distributions and land surface spectral albedos observed by MODIS. Since susceptibilities can potentially be computed from model fields, these results have practical application in assessing the reasonableness of model-generated estimates of the aerosol indirect radiative forcing.

Oreopoulos, Lazaros↗