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All sourcesNASA NTRSDOE OSTI
NASA NTRS2004

A Re-programmable Platform for Dynamic Burn-in Test of Xilinx Virtexll 3000 FPGA for Military and Aerospace Applications

Field Programmable Gate Arrays (FPGA) have played increasingly important roles in military and aerospace applications. Xilinx SRAM-based FPGAs have been extensively used in commercial applications. They have been used less frequently in space flight applications due to their susceptibility to single-event upsets. Reliability of these devices in space applications is a concern that has not been addressed. The objective of this project is to design a fully programmable hardware/software platform that allows (but is not limited to) comprehensive static/dynamic burn-in test of Virtex-II 3000 FPGAs, at speed test and SEU test. Conventional methods test very few discrete AC parameters (primarily switching) of a given integrated circuit. This approach will test any possible configuration of the FPGA and any associated performance parameters. It allows complete or partial re-programming of the FPGA and verification of the program by using read back followed by dynamic test. Designers have full control over which functional elements of the FPGA to stress. They can completely simulate all possible types of configurations/functions. Another benefit of this platform is that it allows collecting information on elevation of the junction temperature as a function of gate utilization, operating frequency and functionality. A software tool has been implemented to demonstrate the various features of the system. The software consists of three major parts: the parallel interface driver, main system procedure and a graphical user interface (GUI).

Field Programmable Gate Arrays (FPGA)↗
NASA NTRSDate not supplied

Emulated Spacecraft Communication Testbed for Evaluating Cognitive Networking Technology

The ability to emulate the full space protocol stack is an essential aspect required to evaluate and mature cognitive communication capabilities. The interaction between the physical layer and network layers is key to developing network optimizations for a dynamic and complex environment. We present a laboratory testbed for the evaluation of cognitive radio and networking techniques applied to space communications. The testbed is a high fidelity, flight-like hardware testbed consisting of software-defined radios, channel emulators, modems, and orbital analysis and scheduling software. The testbed uses RF links with signal quality, propagation delay, and Doppler effects driven by orbital mechanics simulations of emulated spacecraft. Our framework enables control of link bidirectionality, data rates, and interference sources. In addition to hardware radio nodes, the testbed can incorporate virtualized emulated nodes for larger and more challenging network scenarios. Our approach to a cognitive communication system uses delay tolerant networking (DTN) to mitigate the challenges of the space environment. While many DTN networks use only preplanned schedules, our system uses User-Initiated Service (UIS) to dynamically schedule service providers. Software-defined radio allows the system to adapt to a variety of service providers. Integration of DTN, UIS, and software-defined radio technologies provides a framework for the implementation of a cognitive communication system. This paper describes the testbed capabilities, network emulation approach, component integration, and initial end-to-end testing results.

cognitive radio↗
NASA NTRS2011

Determination of Barometric Altimeter Errors for the Orion Exploration Flight Test-1 Entry

The EFT-1 mission is the unmanned flight test for the upcoming Multi-Purpose Crew Vehicle (MPCV). During entry, the EFT-1 vehicle will trigger several Landing and Recovery System (LRS) events, such as parachute deployment, based on onboard altitude information. The primary altitude source is the filtered navigation solution updated with GPS measurement data. The vehicle also has three barometric altimeters that will be used to measure atmospheric pressure during entry. In the event that GPS data is not available during entry, the altitude derived from the barometric altimeter pressure will be used to trigger chute deployment for the drogues and main parachutes. Therefore it is important to understand the impact of error sources on the pressure measured by the barometric altimeters and on the altitude derived from that pressure. There are four primary error sources impacting the sensed pressure: sensor errors, Analog to Digital conversion errors, aerodynamic errors, and atmosphere modeling errors. This last error source is induced by the conversion from pressure to altitude in the vehicle flight software, which requires an atmosphere model such as the US Standard 1976 Atmosphere model. There are several secondary error sources as well, such as waves, tides, and latencies in data transmission. Typically, for error budget calculations it is assumed that all error sources are independent, normally distributed variables. Thus, the initial approach to developing the EFT-1 barometric altimeter altitude error budget was to create an itemized error budget under these assumptions. This budget was to be verified by simulation using high fidelity models of the vehicle hardware and software. The simulation barometric altimeter model includes hardware error sources and a data-driven model of the aerodynamic errors expected to impact the pressure in the midbay compartment in which the sensors are located. The aerodynamic model includes the pressure difference between the midbay compartment and the free stream pressure as a function of altitude, oscillations in sensed pressure due to wake effects, and an acoustics model capturing fluctuations in pressure due to motion of the passive vents separating the barometric altimeters from the outside of the vehicle.

Brown, Denise L.↗
NASA NTRS2015

Internship Abstract - Aerosciences and Flight Mechanics Intern

Mars is a hard place to land on, but my internship with NASA's Aerosciences & Flight Mechanics branch has shown me the ways in which men and women will one day land safely. I work on Mars Aerocapture, an aeroassist maneuver that reduces the fuel necessary to "capture" into Martian orbit before a descent. The spacecraft flies through the Martian atmosphere to lose energy through heating before it exits back into space, this time at a slower velocity and in orbit around Mars. Spacecraft will need to maneuver through the Martian atmosphere to accurately hit their orbit, and they will need to survive the generated heat. Engineering teams need simulation data to continue their designs, and the guidance algorithm that ensures a proper orbit insertion needs to be refined - two jobs that fell to me at the summer's start. Engineers within my branch have developed two concept aerocapture vehicles, and I run simulations on their behavior during the maneuver. I also test and refine the guidance algorithm. I spent the first few weeks familiarizing myself with the simulation software, troubleshooting various guidance bugs and writing code. Everything runs smoothly now, and I recently sent my first set of trajectory data to a Thermal Protection System group so they can incorporate it into their heat-bearing material designs. I hope to generate plenty of data in the next few weeks for various engineering groups before my internship ends mid-August. My major accomplishment so far is improving the guidance algorithm. It is a relatively new algorithm that promises higher accuracy and fuel efficiency, but it hasn't undergone extensive testing yet. I've had the opportunity to work with the principal developer - a professor at Iowa State University - to find and fix several issues. I was also assigned the task of expanding the branch's aerodynamic heating simulation software. I am excited to do this because engineers in the future will use my work to generate meaningful data and make design decisions. My internship has taught me to how to teach myself. There are no tutors, study sessions or professor office hours. When I am given an assignment I am expected to figure out how to accomplish it, and I have grown in my problem solving abilities. My summer experience has reinforced my drive to work at NASA, and I can definitely see myself working full time on the aerocapture project, or something similar.

Rangel, John↗
NASA NTRS2016

A Structured, Model-Based Systems Engineering Methodology for Operations System Design

Two widely accepted techniques for lowering the cost and risk of developing systems are (1) the use of a defined systems engineering (SE) process or methodology and (2) the reuse of existing (previously built) system components. The first technique is represented, for example, in materials published by NASA (e.g., NASA Systems Engineering Handbook) or by professional societies such as INCOSE (International Council on Systems Engineering). Well-formed SE techniques provide value by establishing the proper scope of the system (e.g., requirements), and by identifying and resolving problems relatively early in project lifecycles, when fixes are less expensive. The second technique (reuse) is applied most commonly to hardware and software; it seeks to avoid replicating design and implementation costs while also reducing risk by placing proven capabilities into operational use. In this paper, we outline a methodology combining these two techniques and extending reuse beyond hardware and software to foundational aspects of a Mission Operation System’s (MOS) design. We describe the system design artifacts that result (e.g., requirements, design documentation), as well as the reusable patterns and elements of the design, and their interrelationships. This approach is enabled by model-based systems engineering (MBSE) techniques and tools and is currently available in SysML form as a plug-in to MagicDraw. Additionally, usage of a rigorous MBSE approach allows for training materials and tutorials to be packaged within the overall model itself. The results of such an approach include decreased cost and risk during the design phase, improved ability of the MOS development team to investigate trade spaces and identify impacts to important flight-ground trade studies. Such results extend into decreased costs and risk in later phases due to improved design, decreased need for late fixes or development of "glue-ware" or scripts to fill unanticipated gaps in functionality, and improved ability to identify and plan testing and other validation activities. Finally, lower operational costs can be expected, both due to improved quality of the MOS, increased ease of maintaining updated knowledge of system configuration, and the fact that training and procedural materials are also updated at the same time as accepted system changes.

Bindschadler, Duane L.↗
NASA NTRSDate not supplied

Micrometeoroid and Orbital Debris (MMOD) Testing, Ballistic Limit Equation Definition and Risk Assessment of the Exploration Extravehicular Mobility Unit (xEMU)

A well-known hazard associated with exposure to the space environment is the risk of failure due to an impact from a micrometeoroid and orbital debris (MMOD) particle. As NASA prepares to return astronauts to the moon with the Artemis program, the next generation of spacesuit is in development to support future extravehicular activities (EVAs.) An MMOD impact to the spacesuit is of great concern as a large leak could prevent an astronaut from safely reaching an airlock in time resulting in a loss of life. The exploration extravehicular mobility unit (xEMU) must meet MMOD requirements for multiple environments including those in low earth orbit (LEO) as well as the meteoroid and secondary lunar regolith ejecta environments found on the lunar surface. The subject of this paper is an internal xEMU configuration design developed by NASA Johnson Space Center (JSC) personnel. This paper will expand on the hypervelocity impact (HVI) testing and ballistic limit equation (BLE) definition work that was partially presented at the 2nd International Orbital De-bris (IOC-II) Conference held in Sugar Land, TX in December 2023. The xEMU shares similarities with the legacy Extravehicular Mobility Unit (EMU) spacesuit that is currently used for ISS EVAs, however differences in the layup (e.g., materials, thicknesses, and layers) of the fabric environmental protection garment (EPG), portable life support system (xPLSS) and helmet required an extensive test program to determine ballistic performance. Over 100 hypervelocity impact (HVI) tests were performed by the NASA/JSC HVIT and White Sands Test Facility (WSTF) teams on the xEMU EPG, xPLSS and helmet to generate ballistic limit equations (BLEs) for MMOD impacts. Additionally, over 50 low speed tests (< 1km/s) were performed by the NASA/JSC HVIT and Southwest Research Institute (SwRI) teams on the xEMU EPG, xPLSS and helmet to generate BLEs for lunar ejecta impacts. Post testing, ballistic limit equations used to define the performance of the various regions on the xEMU spacesuit were developed from a generic set of BLEs. The HVI and low speed testing was performed to establish a physical basis for the equations with the co-efficients and exponents of the generic BLEs adjusted to fit the test data. The xEMU BLEs were added to the NASA/JSC software application used for space-craft MMOD risk assessments (BUMPER-3). A finite element model (FEM) of the xEMU spacesuit, which defines the size and shape of the spacesuit as well as the locations of the various shielding configurations, was created based on a solid model provided by the xEMU program office. Using the FEM file and added xEMU BLEs, BUMPER-3 assessments of the xEMU spacesuit for probability of no penetration (PNP) were performed. For the LEO assessment of a typical ISS EVA, the orbital debris and meteoroids environments were defined using the latest engineering models, ORDEM 3.2 and MEM-3 respectively. The lunar sur-face assessment again used the MEM-3 engineering model to define the meteoroid environ-ment along with the current released lunar surface ejecta model, NASA SP-8013 (developed during the Apollo Program). The Space Team in the Natural Environments Branch at Mar-shall Space Flight Center (MSFC) will soon release the new Lunar Meteoroid Ejecta Engineering Model (LMEEM), at which time the xEMU lunar surface EVA will be reassessed. Assessment of the MMOD risk for an 8-hour, 2-person EVA in both LEO and on the lunar surface showed that the xEMU spacesuit meets the program technical requirement of 1 in 2500 failure odds. Similar to the legacy EMU spacesuit, the majority of the MMOD risk (96% of the LEO EVA risk and 99% of the lunar surface EVA risk) is concentrated in regions of xEMU that are comprised primarily of softgoods (arms, legs, and gloves) rather than the hardgoods (xPLSS, hard upper torso and helmet).

Micrometeoroid↗
NASA NTRS1998

The Role of Computational Fluid Dynamics at Marshall Space Flight Center

The Marshall Space Flight Center (MSFC) is the Center of Excellence for Space Propulsion Technology within NASA. One of MSFC's goals is to lower the cost of access to space. MSFC's Fluid Dynamics Division, a part of the Structures and Dynamics Lab, assists in accomplishing this goal. This division is divided into two branches: experimental and analytical, whose work is complimentary and integrated. In the analysis branch, the emphasis is to apply, rather than develop computational fluid dynamics (CFD), to reduce the cost of design and development of: (1) pump and turbine feedlines and feed manifolds; (2) all primary and secondary flow paths in pumps; (3) all primary and secondary flow paths in turbines; (4) combustion chambers; (5) rocket engine nozzles and plumes; and (6) launch vehicles ascent. The broad range of applications imposes unique constraints on the software. The use of CFD in evaluating rocket engine hardware, specifically, is of value due to several reasons: (1) Rocket engine hardware is extremely expensive to build and test; (2) high engineering and development cost is due to limited production; (3) the cost of one engine test is greater than the average yearly hardware and software cost; (4) CFD represents a relatively small investment relative to the impact it has on the engine development; and (5) computing costs are continuously decreasing. In terms of accuracy the predictable error is more useful than the inconsistent accuracy. The achievable accuracy varies from application to application. The lack of accuracy is often caused by improper analyses: such as improperly applied or located boundary conditions, or attempting steady analysis of unsteady flows. This report is in viewgraph format.

Garcia, Roberto↗
NASA NTRS1997

The International Space Station Assembly on Schedule

As engineers continue to prepare the International Space Station (ISS) for in-orbit assembly in the year 2002, ANSYS software has proven instrumental in resolving a structural problem in the project's two primary station modules -- Nodes 1 and 2. Proof pressure tests performed in May revealed "low temperature, post-yield creep" in some of the Nodes' gussets, which were designed to reinforce ports for loads from station keeping and reboost motion of the entire space station. An extensive effort was undertaken to characterize the creep behavior of the 2219-T851 aluminum forging material from which the gussets were made. Engineers at Sverdrup Technology, Inc. (Huntsville, AL) were responsible for conducting a combined elastic-plastic-creep analysis of the gussets to determine the amount of residual compressive stress which existed in the gussets following the proof pressure tests, and to determine the stress-strain history in the gussets while on-orbit. Boeing, NASA's Space Station prime contractor, supplied the Finite Element Analysis (FEA) model geometry and developed the creep equations from the experimental data taken by NASA's Marshall Space Flight Center and Langley Research Center. The goal of this effort was to implement the uniaxial creep equations into a three dimensional finite element program, and to determine analytically whether or not the creep was something that the space station program could live with. The objective was to show analytically that either the creep rate was at an acceptable level, or that the node module had to be modified to lower the stress levels to where creep did not occur. The elastic-plastic-creep analysis was performed using the ANSYS finite element program of ANSYS, Inc. (Houston, PA). The analysis revealed that the gussets encountered a compressive stress of approximately 30,000 pounds per square inch (psi) when unloaded. This compressive residual stress significantly lowered the maximum tension stress in the gussets which decreased the creep strain rate. The analysis also showed that the gussets would not experience a great deal of creep from future pressure tests if braces or struts proposed by Boeing were installed to redistribute stress away from them. Subsequent analysis of on-orbit station keeping and reboost loads convinced Boeing that the gussets should be removed altogether.

Source record↗
NASA NTRS2004

Development of a Crosslink Channel Simulator

Distributed Spacecraft missions are an integral part of current and future plans for NASA and other space agencies. Many of these multi-vehicle missions involve utilizing the array of spacecraft as a single, instrument requiring communication via crosslinks to achieve mission goals. NASA s Goddard Space Flight Center (GSFC) is developing the Formation Flying Test Bed (FFTB) to provide a hardware-in-the-loop simulation environment to support mission concept development and system trades with a primary focus on Guidance, Navigation, and Control (GN&C) challenges associated with spacecraft flying. The goal of the FFTB is to reduce mission risk by assisting in mission planning and analysis, provide a technology development platform that allows algorithms to be developed for mission functions such as precision formation navigation and control and time synchronization. The FFTB will provide a medium in which the various crosslink transponders being used in multi-vehicle missions can be integrated for development and test; an integral part of the FFTB is the Crosslink Channel Simulator (CCS). The CCS is placed into the communications channel between the crosslinks under test, and is used to simulate on-mission effects to the communications channel such as vehicle maneuvers, relative vehicle motion, or antenna misalignment. The CCS is based on the Starlight software programmable platform developed at General Dynamics Decision Systems and provides the CCS with the ability to be modified on the fly to adapt to new crosslink formats or mission parameters. This paper briefly describes the Formation Flying Test Bed and its potential uses. It then provides details on the current and future development of the Crosslink Channel Simulator and its capabilities.

Hunt, Chris↗
NASA NTRS2000

The Low Temperature Microgravity Physics Experiments Project

The Low Temperature Microgravity Physics Facility (LTMPF) is being developed by NASA to provide long duration low temperature and microgravity environment on the International Space Station (ISS) for performing fundamental physics investigations. Currently, six experiments have been selected for flight definition studies. More will be selected in a two-year cycle, through NASA Research Announcement. This program is managed under the Low Temperature Microgravity Physics Experiments Project Office at the Jet Propulsion Laboratory. The facility is being designed to launch and returned to earth on a variety of vehicles including the HII-A and the space shuttle. On orbit, the facility will be connected to the Exposed Facility on the Japanese Experiment Module, Kibo. Features of the facility include a cryostat capable of maintaining super-fluid helium at a temperature of 1.4 K for 5 months, resistance thermometer bridges, multi-stage thermal isolation system, thermometers capable of pico-Kelvin resolution, DC SQUID magnetometers, passive vibration isolation, and magnetic shields with a shielding factor of 80dB. The electronics and software architecture incorporates two VME buses run using the VxWorks operating system. Technically challenging areas in the design effort include the following: 1) A long cryogen life that survives several launch and test cycles without the need to replace support straps for the helium tank. 2) The minimization of heat generation in the sample stage caused by launch vibration 3) The design of compact and lightweight DC SQUID electronics. 4) The minimization of RF interference for the measurement of heat at pico-Watt level. 5) Light weighting of the magnetic shields. 6) Implementation of a modular and flexible electronics and software architecture. The first launch is scheduled for mid-2003, on an H-IIA Rocket Transfer Vehicle, out of the Tanegashima Space Center of Japan. Two identical facilities will be built. While one facility is onboard the ISS, the other is re-integrated on the ground with new experiments. When the cryogen of the facility in space are exhausted, it will be swapped with the other facility with the new experiment. A total of 20 science missions are envisioned over the next 20 years.

Holmes, Warren↗
NASA NTRS2014

Evolution of the JPSS Ground Project Calibration and Validation System

The Joint Polar Satellite System (JPSS) is the National Oceanic and Atmospheric Administrations (NOAA) next-generation operational Earth observation Program that acquires and distributes global environmental data from multiple polar-orbiting satellites. The JPSS Program plays a critical role to NOAAs mission to understand and predict changes in weather, climate, oceans, coasts, and space environments, which supports the Nation’s economy and protection of lives and property. The National Aerospace and Atmospheric Administration (NASA) is acquiring and implementing the JPSS, comprised of flight and ground systems on behalf of NOAA. The JPSS satellites are planned to fly in the afternoon orbit and will provide operational continuity of satellite-based observations and products for NOAA Polar-orbiting Operational Environmental Satellites (POES) and the Suomi National Polar-orbiting Partnership (SNPP) satellite. To support the JPSS Calibration and Validation (CalVal) node Government Resource for Algorithm Verification, Independent Test, and Evaluation (GRAVITE) services facilitate: Algorithm Integration and Checkout, Algorithm and Product Operational Tuning, Instrument Calibration, Product Validation, Algorithm Investigation, and Data Quality Support and Monitoring. GRAVITE is a mature, deployed system that currently supports the SNPP Mission and has been in operations since SNPP launch. This paper discusses the major re-architecture for Block 2.0 that incorporates SNPP lessons learned, architecture of the system, and demonstrates how GRAVITE has evolved as a system with increased performance. It is now a robust, stable, reliable, maintainable, scalable, and secure system that supports development, test, and production strings, replaces proprietary and custom software, uses open source software, and is compliant with NASA and NOAA standards.

Data Processing↗
NASA NTRS2016

Evolution of the JPSS Ground Project Calibration and Validation System

The Joint Polar Satellite System (JPSS) is the National Oceanic and Atmospheric Administration's (NOAA) next-generation operational Earth observation Program that acquires and distributes global environmental data from multiple polar-orbiting satellites. The JPSS Program plays a critical role to NOAA's mission to understand and predict changes in weather, climate, oceans, coasts, and space environments, which supports the Nation's economy and protection of lives and property. The National Aeronautics and Space Administration (NASA) is acquiring and implementing the JPSS, comprised of flight and ground systems, on behalf of NOAA. The JPSS satellites are planned to fly in the afternoon orbit and will provide operational continuity of satellite-based observations and products for NOAA Polar-orbiting Operational Environmental Satellites (POES) and the Suomi National Polar-orbiting Partnership (SNPP) satellite. To support the JPSS Calibration and Validation (CalVal) node Government Resource for Algorithm Verification, Independent Test, and Evaluation (GRAVITE) services facilitate: Algorithm Integration and Checkout, Algorithm and Product Operational Tuning, Instrument Calibration, Product Validation, Algorithm Investigation, and Data Quality Support and Monitoring. GRAVITE is a mature, deployed system that currently supports the SNPP Mission and has been in operations since SNPP launch. This paper discusses the major re-architecture for Block 2.0 that incorporates SNPP lessons learned, architecture of the system, and demonstrates how GRAVITE has evolved as a system with increased performance. It is now a robust, stable, reliable, maintainable, scalable, and secure system that supports development, test, and production strings, replaces proprietary and custom software, uses open source software, and is compliant with NASA and NOAA standards.

Next-Generation↗
NASA NTRS2017

Physiological and Molecular Changes in Various Biological Organisms Cultured under Simulated Microgravity Conditions

Microgravity is one of the most import factors in space flight where its impact on living biological organisms is concerned. Many different ailments have been reported in astronauts such as spaceflight related osteopenia, cardiovascular concerns, and loss of eye sight. In order to understand why µg causes these issues we must understand what is happening at the most basic of biological structures, the cell. The work done in this report is a culmination of contributions made to a much larger project. The project seeks to understand how cellular physiology is changing in SMG conditions and use this knowledge to feed into a follow-up study on the genetic changes that are seen in SMG environments. Cells were imaged using confocal microscopy after 20hrs and 48hrs in a 3D clinostat called the Gravite. Lengths, widths, heights, and total cell areas were measured using an image analysis software package ImageJ. There were significant differences in lengths and widths of cell nuclei, and total area of cell coverage. The report then discusses some of the problems with the testing apparatus and how 3D printing technology may be used to create better sample holders for the 3D clinostat.

cell biology↗
NASA NTRS2005

Robotic and Human-Tended Collaborative Drilling Automation for Subsurface Exploration

Future in-situ lunar/martian resource utilization and characterization, as well as the scientific search for life on Mars, will require access to the subsurface and hence drilling. Drilling on Earth is hard - an art form more than an engineering discipline. Human operators listen and feel drill string vibrations coming from kilometers underground. Abundant mass and energy make it possible for terrestrial drilling to employ brute-force approaches to failure recovery and system performance issues. Space drilling will require intelligent and autonomous systems for robotic exploration and to support human exploration. Eventual in-situ resource utilization will require deep drilling with probable human-tended operation of large-bore drills, but initial lunar subsurface exploration and near-term ISRU will be accomplished with lightweight, rover-deployable or standalone drills capable of penetrating a few tens of meters in depth. These lightweight exploration drills have a direct counterpart in terrestrial prospecting and ore-body location, and will be designed to operate either human-tended or automated. NASA and industry now are acquiring experience in developing and building low-mass automated planetary prototype drills to design and build a pre-flight lunar prototype targeted for 2011-12 flight opportunities. A successful system will include development of drilling hardware, and automated control software to operate it safely and effectively. This includes control of the drilling hardware, state estimation of both the hardware and the lithography being drilled and state of the hole, and potentially planning and scheduling software suitable for uncertain situations such as drilling. Given that Humans on the Moon or Mars are unlikely to be able to spend protracted EVA periods at a drill site, both human-tended and robotic access to planetary subsurfaces will require some degree of standalone, autonomous drilling capability. Human-robotic coordination will be important, either between a robotic drill and humans on Earth, or a human-tended drill and its visiting crew. The Mars Analog Rio Tinto Experiment (MARTE) is a current project that studies and simulates the remote science operations between an automated drill in Spain and a distant, distributed human science team. The Drilling Automation for Mars Exploration (DAME) project, by contrast: is developing and testing standalone automation at a lunar/martian impact crater analog site in Arctic Canada. The drill hardware in both projects is a hardened, evolved version of the Advanced Deep Drill (ADD) developed by Honeybee Robotics for the Mars Subsurface Program. The current ADD is capable of 20m, and the DAME project is developing diagnostic and executive software for hands-off surface operations of the evolved version of this drill. The current drill automation architecture being developed by NASA and tested in 2004-06 at analog sites in the Arctic and Spain will add downhole diagnosis of different strata, bit wear detection, and dynamic replanning capabilities when unexpected failures or drilling conditions are discovered in conjunction with simulated mission operations and remote science planning. The most important determinant of future 1unar and martian drilling automation and staffing requirements will be the actual performance of automated prototype drilling hardware systems in field trials in simulated mission operations. It is difficult to accurately predict the level of automation and human interaction that will be needed for a lunar-deployed drill without first having extensive experience with the robotic control of prototype drill systems under realistic analog field conditions. Drill-specific failure modes and software design flaws will become most apparent at this stage. DAME will develop and test drill automation software and hardware under stressful operating conditions during several planned field campaigns. Initial results from summer 2004 tests show seven identifi distinct failure modes of the drill: cuttings-removal issues with low-power drilling into permafrost, and successful steps at executive control and initial automation.

Glass, Brian↗
NASA NTRS2016

A Multi-Center Space Data System Prototype Based on CCSDS Standards

Deep space missions beyond earth orbit will require new methods of data communications in order to compensate for increasing Radio Frequency (RF) propagation delay. The Consultative Committee for Space Data Systems (CCSDS) standard protocols Spacecraft Monitor & Control (SM&C), Asynchronous Message Service (AMS), and Delay/Disruption Tolerant Networking (DTN) provide such a method. However, the maturity level of this protocol stack is insufficient for mission inclusion at this time. This Space Data System prototype is intended to provide experience which will raise the Technical Readiness Level (TRL) of this protocol set. In order to reduce costs, future missions can take advantage of these standard protocols, which will result in increased interoperability between control centers. This prototype demonstrates these capabilities by implementing a realistic space data system in which telemetry is published to control center applications at the Jet Propulsion Lab (JPL), the Marshall Space Flight Center (MSFC), and the Johnson Space Center (JSC). Reverse publishing paths for commanding from each control center are also implemented. The target vehicle consists of realistic flight computer hardware running Core Flight Software (CFS) in the integrated Power, Avionics, and Power (iPAS) Pathfinder Lab at JSC. This prototype demonstrates a potential upgrade path for future Deep Space Network (DSN) modification, in which the automatic error recovery and communication gap compensation capabilities of DTN would be exploited. In addition, SM&C provides architectural flexibility by allowing new service providers and consumers to be added efficiently anywhere in the network using the common interface provided by SM&C's Message Abstraction Layer (MAL). In FY 2015, this space data system was enhanced by adding telerobotic operations capability provided by the Robot API Delegate (RAPID) family of protocols developed at NASA. RAPID is one of several candidates for consideration and inclusion in a new international standard being developed by the CCSDS Telerobotic Operations Working Group. Software gateways for the purpose of interfacing RAPID messages with the existing SM&C based infrastructure were developed. Telerobotic monitor, control, and bridge applications were written in the RAPID framework, which were then tailored to the NAO telerobotic test article hardware, a product of Aldebaran Robotics.

Rich, Thomas M.↗
NASA NTRS2019

Shape Sensing for Wings with Spars and Ribs Using Simulated Strain

Active trim shape control can be used to minimize error between target and actual aircraft trim shape during flight. Trim shape sensing for aircraft during flight is not only important for highly flexible aircraft, such as the National Aeronautics and Space Administration (NASA) Helios Prototype remotely piloted flying wing aircraft, but also for a delta-wing type aircraft, such as a supersonic commercial transport aircraft. A two-step theory utilizing distributed strain for a real-time shape sensing of a full three-dimensional structure has been introduced previously. This study focuses on the application of the two-step theory to finite element models of a wing with spars and ribs such as the X-59 QueSST aircraft (Lockheed Martin Corporation, Bethesda, Maryland), a tapered wing, a dihedral/anhedral wing, and a stiffened dihedral/anhedral wing. A finely meshed finite element structural model is desired to capture accurate curvature distributions along the neutral axes of the wing cross sections during pre-test analysis for shape sensing of a wing with ribs and spars. The two-step theory used in this study gives excellent deformation correlation with the MSC/NASTRAN (MSC Software, Newport Beach, California) results along the neutral axis for all test cases used in this study except the X-59 QueSST aircraft.

strain↗
NASA NTRS2000

Hardware Development Process for Human Research Facility Applications

The simple goal of the Human Research Facility (HRF) is to conduct human research experiments on the International Space Station (ISS) astronauts during long-duration missions. This is accomplished by providing integration and operation of the necessary hardware and software capabilities. A typical hardware development flow consists of five stages: functional inputs and requirements definition, market research, design life cycle through hardware delivery, crew training, and mission support. The purpose of this presentation is to guide the audience through the early hardware development process: requirement definition through selecting a development path. Specific HRF equipment is used to illustrate the hardware development paths. The source of hardware requirements is the science community and HRF program. The HRF Science Working Group, consisting of SCientists from various medical disciplines, defined a basic set of equipment with functional requirements. This established the performance requirements of the hardware. HRF program requirements focus on making the hardware safe and operational in a space environment. This includes structural, thermal, human factors, and material requirements. Science and HRF program requirements are defined in a hardware requirements document which includes verification methods. Once the hardware is fabricated, requirements are verified by inspection, test, analysis, or demonstration. All data is compiled and reviewed to certify the hardware for flight. Obviously, the basis for all hardware development activities is requirement definition. Full and complete requirement definition is ideal prior to initiating the hardware development. However, this is generally not the case, but the hardware team typically has functional inputs as a guide. The first step is for engineers to conduct market research based on the functional inputs provided by scientists. CommerCially available products are evaluated against the science requirements as well as modifications needed to meet program requirements. Options are consolidated and the hardware development team reaches a hardware development decision point. Within budget and schedule constraints, the team must decide whether or not to complete the hardware as an in-house, subcontract with vendor, or commercial-off-the-shelf (COTS) development. An in-house development indicates NASA personnel or a contractor builds the hardware at a NASA site. A subcontract development is completed off-site by a commercial company. A COTS item is a vendor product available by ordering a specific part number. The team evaluates the pros and cons of each development path. For example, in-bouse developments utilize existing corporate knowledge regarding bow to build equipment for use in space. However, technical expertise would be required to fully understand the medical equipment capabilities, such as for an ultrasound system. It may require additional time and funding to gain the expertise that commercially exists. The major benefit of subcontracting a hardware development is the product is delivered as an end-item and commercial expertise is utilized. On the other hand, NASA has limited control over schedule delays. The final option of COTS or modified COTS equipment is a compromise between in-house and subcontracts. A vendor product may exist that meets all functional requirements but req uires in-house modifications for successful operation in a space environment. The HRF utilizes equipment developed using all of the paths described: inhouse, subcontract, and modified COTS.

Bauer, Liz↗
DOE OSTI2021

Air Blast Mesh Sensitivity and Pressure Mapping Study

Nose cone structural and aerodynamic characteristics are essential for intelligent design of aircraft, spacecraft, and ballistic systems. Finite element analysis can be used to help understand the structural integrity and flight characteristics of different nose cones. A mesh sensitivity study was undertaken for a particular nose cone geometry that was used in tests at LANL facilities in order to confirm the integrity of the meshed geometry. A simple cone that best matched closed-form theoretical solutions was modeled, and received good correlation to the theory. Complexity was then added back to the nose cone. Parameters applied to the simple cone were then implemented in the nose cone geometry giving assurance of accuracy after the geometry was changed. Nose cone results averaged 6.3% error for radial displacement when compared with the theoretical. Hoop stress averaged 6.0% error and meridional stress averaged 5.7% error at the finest mesh level. Meshes showed signs of convergence when compared to all three theoretical solutions. Finally, pressure time-history data from LANL computational fluid dynamics simulations was applied to the surface of the final nose cone geometry. The pressure data was interpolated from pressure gauge locations onto nearby meshed elements, which allowed for FEA software to run simulations on the cone with the pressure data as a loading condition. The pressure mapping resulted in the ability to understand the nose cone’s rigid body motion that in turn can inform design of future nose cones.

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