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At least 181 records · Page 10
All sourcesNASA NTRSDOE OSTI
NASA NTRS2014

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for an MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV), an Earth Return Vehicle (ERV), and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Super Sonic Retro Propulsion (SSRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars. The MAV uses a storable liquid bi-propellant propulsion system to deliver the ERV to a Mars phasing orbit. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Upon arrival at Earth, the ERV performs Earth and lunar swing-bys and is placed into a lunar trailing circular orbit - an Earth orbit, at lunar distance. A later mission, using Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft makes a controlled Earth re-entry preventing any unintended release of pristine martian materials into the Earth's biosphere. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX, published analyses from other sources, as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the effect of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, shows no significant stressors. A useful payload mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report options for the MAV and ERV, including propulsion systems, crewed versus robotic retrieval mission, as well as direct Earth entry. International planetary protection policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. We also describe preliminary compliance measures that will be the subject of future work. This work shows that emerging commercial capabilities as well as new methodologies can be used to efficiently support an important planetary science objective. The work also has applications for human exploration missions that use propulsive EDL techniques

Red Dragon↗
NASA NTRS2013

Mars Sample Return Using Commercial Capabilities: Mission Architecture Overview

Mars Sample Return (MSR) is the highest priority science mission for the next decade as recommended by the recent Decadal Survey of Planetary Science. This paper presents an overview of a feasibility study for a MSR mission. The objective of the study was to determine whether emerging commercial capabilities can be used to reduce the number of mission systems and launches required to return the samples, with the goal of reducing mission cost. The major element required for the MSR mission are described and include an integration of the emerging commercial capabilities with small spacecraft design techniques; new utilizations of traditional aerospace technologies; and recent technological developments. We report the feasibility of a complete and closed MSR mission design using the following scenario that covers three synodic launch opportunities, beginning with the 2022 opportunity: A Falcon Heavy injects a SpaceX Red Dragon capsule and trunk onto a Trans Mars Injection (TMI) trajectory. The capsule is modified to carry all the hardware needed to return samples collected on Mars including a Mars Ascent Vehicle (MAV); an Earth Return Vehicle (ERV); and hardware to transfer a sample collected in a previously landed rover mission to the ERV. The Red Dragon descends to land on the surface of Mars using Supersonic Retro Propulsion (SRP). After previously collected samples are transferred to the ERV, the single-stage MAV launches the ERV from the surface of Mars to a Mars phasing orbit. The MAV uses a storable liquid, pump fed bi-propellant propulsion system. After a brief phasing period, the ERV, which also uses a storable bi-propellant system, performs a Trans Earth Injection (TEI) burn. Once near Earth the ERV performs Earth and lunar swing-bys and is placed into a Lunar Trailing Orbit (LTO0 - an Earth orbit, at lunar distance. A later mission, using a Dragon and launched by a Falcon Heavy, performs a rendezvous with the ERV in the lunar trailing orbit, retrieves the sample container and breaks the chain of contact with Mars by transferring the sample into a sterile and secure container. With the sample contained, the retrieving spacecraft, makes a controlled Earth re-entry preventing any unintended release of pristine Martian materials into the Earth's biosphere. Other capsule type vehicles and associated launchers may be applicable. The analysis methods employed standard and specialized aerospace engineering tools. Mission system elements were analyzed with either direct techniques or by using parametric mass estimating relationships (MERs). The architecture was iterated until overall mission convergence was achieved on at least one path. Subsystems analyzed in this study include support structures, power system, nose fairing, thermal insulation, actuation devices, MAV exhaust venting, and GN&C. Best practice application of loads, mass growth contingencies, and resource margins were used. For Falcon Heavy capabilities and Dragon subsystems we utilized publically available data from SpaceX; published analyses from other sources; as well as our own engineering and aerodynamic estimates. Earth Launch mass is under 11 mt, which is within the estimated capability of a Falcon Heavy, with margin. Total entry masses between 7 and 10 mt were considered with closure occurring between 9 and 10 mt. Propellant mass fractions for each major phase of the EDL - Entry, Terminal Descent, and Hazard Avoidance - have been derived. An assessment of the entry conditions on the thermal protection system (TPS), currently in use for Dragon missions, has been made. And shows no significant stressors. A useful mass of 2.0 mt is provided and includes mass growth allowances for the MAV, the ERV, and mission unique equipment. We also report on alternate propellant options for the MAV and options for the ERV, including propulsion systems; crewed versus robotic retrieval mission; as well as direct Earth entry. International Planetary Protection Policies as well as verifiable means of compliance will have a large impact on any MSR mission design. We identify areas within our architecture where such impacts occur. This work shows that emerging commercial capabilities can be used to effectively integrated into a mission to achieve an important planetary science objective.

Mars Sample Return↗
NASA NTRS2017

Trajectory Design Considerations for Exploration Mission 1

Exploration Mission 1 (EM-1) will be the first mission to send an uncrewed Orion vehicle to cislunar space in 2018, targeted to a Distant Retrograde Orbit (DRO). Analysis of EM-1 DRO mission opportunities in 2018 help characterize mission parameters that are of interest to other subsystems (e.g., power, thermal, communications, flight operations, etc). Subsystems request mission design trades which include: landing lighting, addition of an Orion main engine checkout burn, and use of auxiliary thruster only cases. This paper examines the evolving trade studies that incorporate subsystem feedback and demonstrate the feasibility of these constrained mission trajectory designs and contingencies.

Dawn, Timothy F.↗
NASA NTRS2013

Toxicological Assessment of ISS Air Quality: Contingency Sampling - February 2013

Two grab sample containers (GSCs) were collected by crew members onboard ISS in response to a vinegar-like odor in the US Lab. On February 5, the first sample was collected approximately 1 hour after the odor was noted by the crew in the forward portion of the Lab. The second sample was collected on February 22 when a similar odor was noted and localized to the end ports of the microgravity science glovebox (MSG). The crewmember removed a glove from the MSG and collected the GSC inside the glovebox volume. Both samples were returned on SpaceX-2 for ground analysis.

Meyers, Valerie↗
NASA NTRS1986

Titan Probe navigation analysis

In the proposed Cassini mission, a combined Saturn Orbiter/Titan Probe spacecraft will be launched from the Space Shuttle to arrive at Saturn around 2002, by means of a delta-VEGA trajectory. After Saturn-orbit insertion and a pericrone raise maneuver, the probe will be released to enter the Titan atmosphere and impact onto its surface. During its descent phase and impact onto Titan, the probe will maintain radio contact with the orbiter. Since the Titan-probe experimental phase lasts for only about four hours, probe-orbiter geometry and probe-delivery accuracy are critical to successful completion of this part of the mission. From a preliminary navigation analysis for probe delivery accuracy, it seems feasible to deliver the probe within 50 km (1-sigma value) of the desired aim-point in the Titan B-plane. The covariance study, however, clearly indicates the need for optical data, in addition to radio metric data. A Monte Carlo study indicates that a Delta-V capability of 98 m/sec for trajectory correction maneuvers will be sufficient to cover 99 percent of all contingencies during the segment from Saturn-orbit insertion to Titan-probe release.

Vijayaraghavan, A.↗
NASA NTRS2015

Mission Design Analysis for the Martian Moon Phobos: Close Flybys, Missed Thrusts, and Other In-Flight Entertainment

A robotic mission to the Martian moons Phobos and Deimos would offer a wealth of scientific information and serve as a useful precursor to potential human missions. In this paper, we investigate a prospective mission enabled by solar electric propulsion that would explore Phobos via a series of flybys followed by capture into orbit around the moon. Of particular interest are low-cost options for capture and walkdown to the target science orbits aided by multi-body effects due to the mutual gravitational interaction of Phobos and Mars. We also consider contingency operations in the event of missed thrust or maneuver execution errors.

SEP↗
NASA NTRS2007

Integrating Human Factors into Crew Exploration Vehicle Design

With NASA's new Vision for Exploration to send humans beyond Earth orbit, it is critical to consider the human as a system that demands early and continuous user involvement, and an iterative prototype/test/redesign process. Addressing human-system interface issues early on can be very cost effective even cost reducing when performed early in the design and development cycle. To achieve this goal within Crew Exploration Vehicle (CEV) Project Office, human engineering (HE) team is formed. Key tasks are to apply HE requirements and guidelines to hardware/software, and provide HE design, analysis and evaluation of crew interfaces. Initial activities included many practice-orientated evaluations using low-fidelity CEV mock-ups. What follows is a description of such evaluations that focused on a HE requirement regarding Net Habitable Volume (NHV). NHV is defined as the total remaining pressurized volume available to on-orbit crew after accounting for the loss of volume due to deployed hardware and structural inefficiencies which decrease functional volume. The goal of the NHV evaluations was to develop requirements providing sufficient CEV NHV for crewmembers to live and perform tasks in support of mission goals. Efforts included development of a standard NHV calculation method using computer models and physical mockups, and crew/ stakeholder evaluations. Nine stakeholders and ten crewmembers participated in the unsuited evaluations. Six crewmembers also participated in a suited evaluation. The mock-up was outfitted with volumetric representation of sub-systems such as seats, and stowage bags. Thirteen scenarios were developed to represent mission/crew tasks and considered to be primary volume drivers (e.g., suit donning) for the CEV. Unsuited evaluations included a structured walkthrough of these tasks. Suited evaluations included timed donning of the existing launch and entry suit to simulate a contingency scenario followed by doffing/ stowing of the suits. All mockup evaluations were videotaped. Structured questionnaires were used to document user interface issues and volume impacts of layout configuration. Computer model and physical measures of the NHV agreed within 1 percent. This included measurement of the gross habitable volume, subtraction of intrusive volumes, and other non-habitable spaces. Calculation method developed was validated as a standard means of measuring NHV, and was recommended as a verification method for the NHV requirements. Evaluations confirmed that there was adequate volume for unsuited scenarios and suit donning/ doffing activity. Seats, suit design stowage and waste hygiene system noted to be critical volume drivers. The low-fidelity mock-up evaluations along with human modeling analysis generated discussions that will lead to high-level systems requirements and human-centered design decisions. This approach allowed HE requirements and operational concepts to evolve in parallel with engineering system concepts and design requirements. As the CEV design matures, these evaluations will continue and help with design decisions, and assessment, verification and validation of HE requirements.

Whitmore, Mihriban↗
NASA NTRS1993

An interactive tool for design and support of lunar, gravity assist, and libration point trajectories

The 'Swingby' interactive tool for design, analysis, and support of maneuver planning for missions that involve transfers to the moon, the lunar orbits, multiple-gravity assists, and the libration-point orbits, runs on a PC and uses pull-down and pop-up menus to allow users to change physical constants and select perturbations for inclusion. Swingby can also target by ascertain parameters, or allow the computer to generate goals on the basis of mission requirements. Rapid mission analyses, operational maneuvers, and contingency plans are thereby obtainable.

Carrico, J.↗
NASA NTRS2009

Orion Crew Injury Prediction: Landing Conditions and Vehicle Interface Elements

This presentation reviews the work being done to research the possibility of injuries from landing the Orion Crew Exploration Vehicle in either water or land contingencies This was done using the Total Human Model for Safety (THUMS) finite element model. The purpose of the simulations was to compare variations of a Vehicle Interface Element (VIE) and to evaluate the potential for injury. The presentation is accompanied by several animations from the model runs, showing the projected motion, with and without the various suggested types of VIEs. There are charts showing the injury metrics with the various types of VIEs. The ultimate use of this analysis was to assist in a selection of possible VIE designs

Danelson, Kerry↗
NASA NTRSDate not supplied

Flight Demonstration of Unmanned Aircraft System (UAS) Traffic Management (UTM) at Technical Capability Level 4

The NASA Unmanned Aircraft Systems (UAS) Traffic Management (UTM) Project executed the fourth and final UTM Technical Capability Level demonstration between May and August 2019. Two Federal Aviation Administration (FAA)-designated UAS test sites managed the range, partners, and operations to meet the requirements set forth by the UTM Project. All stakeholders supported the execution of the flight testing through close collaboration. Results of the demonstration indicate the viability of the UTM concept to manage large scale operations and contingencies in an urban environment. The demonstration also provided insight into key technological gaps that must be addressed before such operations are routine, safe, and efficient. Standardization efforts related to UTM and the industry participants of those efforts can leverage the results and experiences of this flight activity to accelerate and more firmly ground forthcoming standards. The FAA and other regulators will be able to leverage results to inform future rule-making and identify additional gaps that require further analysis.

Unmanned Aircraft System Traffic Management↗
NASA NTRS1993

Operational radiological support for the US manned space program

Radiological support for the manned space program is provided by the Space Radiation Analysis Group at NASA/JSC. This support ensures crew safety through mission design analysis, real-time space environment monitoring, and crew exposure measurements. Preflight crew exposure calculations using mission design information are used to ensure that crew exposures will remain within established limits. During missions, space environment conditions are continuously monitored from within the Mission Control Center. In the event of a radiation environment enhancement, the impact to crew exposure is assessed and recommendations are provided to flight management. Radiation dosimeters are placed throughout the spacecraft and provided to each crewmember. During a radiation contingency, the crew could be requested to provide dosimeter readings. This information would be used for projecting crew dose enhancements. New instrumentation and computer technology are being developed to improve the support. Improved instruments include tissue equivalent proportional counter (TEPC)-based dosimeters and charged particle telescopes. Data from these instruments will be telemetered and will provide flight controllers with unprecedented information regarding the radiation environment in and around the spacecraft. New software is being acquired and developed to provide 'smart' space environmental data displays for use by flight controllers.

Golightly, Michael J.↗
NASA NTRS1999

Climate Change and a Global City: An Assessment of the Metropolitan East Coast Region

The objective of the research is to derive an assessment of the potential climate change impacts on a global city - in this case the 31 county region that comprises the New York City metropolitan area. This study comprises one of the regional components that contribute to the ongoing U.S. National Assessment: The Potential Consequences of Climate Variability and Change and is an application of state-of-the-art climate change science to a set of linked sectoral assessment analyses for the Metro East Coast (MEC) region. We illustrate how three interacting elements of global cities react and respond to climate variability and change with a broad conceptual model. These elements include: people (e.g., socio- demographic conditions), place (e.g., physical systems), and pulse (e.g., decision-making and economic activities). The model assumes that a comprehensive assessment of potential climate change can be derived from examining the impacts within each of these elements and at their intersections. Thus, the assessment attempts to determine the within-element and the inter-element effects. Five interacting sector studies representing the three intersecting elements are evaluated. They include the Coastal Zone, Infrastructure, Water Supply, Public Health, and Institutional Decision-making. Each study assesses potential climate change impacts on the sector and on the intersecting elements, through the analysis of the following parts: 1. Current conditions of sector in the region; 2. Lessons and evidence derived from past climate variability; 3. Scenario predictions affecting sector; potential impacts of scenario predictions; 4. Knowledge/information gaps and critical issues including identification of additional research questions, effectiveness of modeling efforts, equity of impacts, potential non-local interactions, and policy recommendations; and 5. Identification of coping strategies - i.e., resilience building, mitigation strategies, new technologies, education that affects decision-making, and better preparedness for contingencies.

Rosenzweig, Cynthia↗
NASA NTRS2014

Application of Fault Management Theory to the Quantitive Selection of a Launch Vehicle Abort Trigger Suite

SHM/FM theory has been successfully applied to the selection of the baseline set Abort Triggers for the NASA SLS center dot Quantitative assessment played a useful role in the decision process  M&FM, which is new within NASA MSFC, required the most "new" work, as this quantitative analysis had never been done before center dot Required development of the methodology and tool to mechanize the process center dot Established new relationships to the other groups  The process is now an accepted part of the SLS design process, and will likely be applied to similar programs in the future at NASA MSFC  Future improvements center dot Improve technical accuracy ‒Differentiate crew survivability due to an abort, vs. survivability even no immediate abort occurs (small explosion with little debris) ‒Account for contingent dependence of secondary triggers on primary triggers ‒Allocate "Δ LOC Benefit" of each trigger when added to the previously selected triggers. center dot Reduce future costs through the development of a specialized tool  Methodology can be applied to any manned/unmanned vehicle, in space or terrestrial

Lo, Yunnhon↗
NASA NTRS1980

Galileo probe forebody entry thermal protection - Aerothermal environments and heat shielding requirements

Solutions are presented for the aerothermal heating environments and the material thermal response for the forebody heatshield on the candidate 242 kg Galileo probe entering the modeled nominal and cold-dense Jovian atmospheres. In the flowfield analysis, a finite difference procedure was employed to obtain benchmark predictions of pressure, radiation and convective heating rates (both laminar and turbulent) and the corresponding wall blowing obtained under the steady state approximation. The fluxes over the probe flank were found to be in a range where spallation is an important mass loss mechanism. The predicted heating rates were also used as boundary conditions for a charring materials ablation which was used to predict thermochemical based surface recession, mass loss and bondline temperatures. The contingency factor of 30% currently employed by NASA was found to be insufficient for entry into the cold-dense atmosphere.

Nicolet, W. E.↗
NASA NTRSDate not supplied

A Concept of Operations (ConOps) of an In-time Aviation Safety Management System (IASMS) for Advanced Air Mobility (AAM)

The growth of new emerging operations involving Advanced Air Mobility (AAM) necessitates developing a perspective for an In-time Aviation Safety Management System (IASMS). This perspective advances from the National Academies report on IASMS and its recommendation for developing a Concept of Operations (ConOps) for IASMS. A ConOps has been developed for In-time System-Wide Safety Assurance (ISSA) from which the IASMS ConOps pivots to provide a robust scope commensurate with the broad vision defined by the National Academies. The IASMS ConOps focuses on emerging operations and spans innovations in Unmanned Aircraft System (UAS) and an increasingly complex ecosystem comprised of a widening mix of vehicles and technologies, Urban Air Mobility (UAM) with industry-federated services, traditional operations, as well as new supersonic aircraft and space launch systems.The challenge for the IASMS ConOps is to be broad to encompass innovations in the coming years and decades while agile to ensure levels of safety compatible with operational and certification requirements of the National Airspace System (NAS). The IASMS ConOps interweaves increasing complexity of operational safety capabilities and unlocking UAS Maturity Levels (UMLs). The relationships between increased complexity of automation and automated systems, fewer operators who are not as traditionally higher skilled, more complex operational environments, and aviation operations management with mixed aircraft and equipage pose a multi-dimensional space for IASMS capabilities essential for safety assurance and risk management. Instantiating IASMS capabilities and how they would be integral to AAM operations and increasing maturity of UAM could be accomplished through a series of Safety Demonstrators. These Safety Demonstrators could provide increased understanding and insight into use of controls for risk mitigation, means of compliance for certification, and operational experience with safety services such as in relation to contingency management. The IASMS capabilities can be viewed as initially residing with the vehicle, airspace, and Supplemental Data Service Provider (SDSP). For example, vehicle capabilities include communications including the command and control link, Remote Identification (ID), conflict advisory/alerting, and UAS system monitoring. These capabilities monitor and assess data such as battery health, aircraft state, and human performance. Complexity of ISAMS capabilities depends on a number of factors. These factors are intendedonly as a notional categorization with the purpose being to reflect the complexity of the AAM ecosystem that would drive up the complexity of ISSA capabilities including systems, sensors, models, standards, and controls. Factors could include the Vehicle Flight Management, Environment, Airspace, and Contingency Management. Each of these factors can be comprised of multiple sub-factors that contribute to increasing complexity. For example, Airspace at a lower level of complexity could be dedicated to UTM operations that are unmonitored, and at a higher level of complexity could involve mixed UTM and ATM operations. The IASMS concept includes safety services that provide data and information to different participants in AAM. The roles and responsibilities of participants can be defined using the Responsible-Accountable-Consulted-Informed (RACI) analysis. For example, for the safety service involving the Remote ID, the Operator would be accountable for providing the data, the Vehicle would be responsible for transmitting it, and the USS, SDSP, Vertiports, FIMS (FAA), and Public Entities such as safety services would be informed by receiving the data. The IASMS ConOps identifies the capabilities needed for risk mitigation and safety assurance in the increasingly complex national airspace. The ConOps serves as a pathway for engaging with industry to gain operational experience including through the Safety Demonstrator series, the RACI analysis, and operational complexity factors. The ConOps serves to integrate these different perspectives to build a cohesive and cogent approach to an AAM safety management system.

In-Time Aviation Safety Management System↗
NASA NTRSDate not supplied

Artemis I Trajectory Design and Optimization

This paper presents the overall trajectory design and optimization process for NASA’s Artemis I mission to send an uncrewed Orion vehicle to a lunar Distant Retrograde Orbit (DRO). The on-orbit trajectory begins at the Space Launch System (SLS) core separation and ends at the Orion service module Earth Entry Interface (EI) point. The details of the trajectory optimization process are presented, including design of nominal and extended mission options, launch windows, and abort options. Novel design techniques are also discussed to account for contingencies, such as using auxiliary thrusters to protect against main engine failure and applying trajectory shaping to mitigate or reduce eclipse durations.

Artemis I↗
NASA NTRS2020

Robustness Computation of Dynamic Controllability in Probabilistic Temporal Networks with Ordinary Distributions

Most existing works in Probabilistic Simple Temporal Networks (PSTNs) base their frameworks on well-defined probability distributions. This paper addresses on PSTN Dynamic Controllability (DC) robustness measure, i.e. the execution success probability of a network under dynamic control.We consider PSTNs where the probability distributions of the contingent edges are ordinary distributed (e.g. non-parametric, non-symmetric). We introduce the concepts of dispatching protocol (DP) as well as DP-robustness, the probability of success under a predefined dynamic policy.We propose a fixed-parameter pseudo-polynomial time algorithm to compute the exact DP-robustness of any PSTN under \textit{NextFirst} protocol, and apply to various PSTN datasets, including the real case of planetary exploration in the context of the Mars 2020 rover, and propose an original structural analysis.

Saint-Guillain, Michael↗
NASA NTRS2008

International Space Station (ISS) S-Band Corona Discharge Anomaly Consultation

The Assembly and Contingency Radio Frequency Group (ACRFG) onboard the International Space Station (ISS) is used for command and control communications and transmits (45 dBm or 32 watts) and receives at S-band. The system is nominally pressurized with gaseous helium (He) and nitrogen (N2) at 8 pounds per square inch absolute (psia). MacDonald, Dettwiler and Associates Ltd. (MDA) was engaged to analyze the operational characteristics of this unit in an effort to determine if the anomalous behavior was a result of a corona event. Based on this analysis, MDA did not recommend continued use of this ACRFG. The NESC was requested to provide expert support in the area of high-voltage corona and multipactoring in an S-Band RF system and to assess the probability of corona occurring in the ACRFG during the planned EVA. The NESC recommended minimal continued use of S/N 002 ACRFG until a replacement unit can be installed. Following replacement, S/N 002 will be subjected to destructive failure analysis in an effort to determine the proximate and root cause(s) of the anomalous behavior.

Kichak, Robert A.↗