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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 217 records · Page 12
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NASA NTRS2011

Performance Characterization of a Microchannel Liquid/Liquid Heat Exchanger Throughout an Extended Duration Life Test

Liquid/Liquid Heat Exchangers (L/L HX) are an integral portion of any spacecraft active thermal control system. For this study the X-38 L/L HX was used as a baseline. As detailed in a previous ICES manuscript, NASA paired with Pacific Northwest National Laboratory to develop a Microchannel L/L HX (MHX). This microchannel HX was designed to meet the same performance characteristics as the aforementioned X-38 HX. The as designed Microchannel HX has a 26% and 60% reduction in mass and volume, respectively. Due to the inherently smaller flow passages the design team was concerned about fouling affecting performance during extended missions. To address this concern, NASA has developed a test stand and is currently performing an 18 month life test on the MHX. This report will detail the up-to-date performance of the MHX during life testing.

Sheth, Rubik B.↗
NASA NTRS2011

Venting of a Water/Inhibited Propylene Glycol Mixture in a Vacuum Environment-Characterization and Representative Test Results

A planned use of the Orion space vehicle involves its residence at the International Space Station for six months at a time. One concept of operations involves temporarily venting portions of the idle Orion active thermal control system (ATCS) during the docked phase, preventing freezing. The venting would have to be reasonably complete with few, if any, completely filled pockets of frozen liquid. Even if pockets of frozen liquid did not damage the hardware during the freezing process, they could prevent the system from filling completely prior to its reactivation. The venting of single component systems in a space environment has been performed numerous times and is well understood. Local nucleation occurs at warm, relatively massive parts of the system, which creates vapor and forces the bulk liquid out of the system. The remnants of the liquid will freeze, then evaporate over time through local heating. Because the Orion ATCS working fluid is a 50/50 mixture of water and inhibited propylene glycol, its boiling behavior was expected to differ from that of a pure fluid. It was thought that the relatively high vapor pressure water might evaporate preferentially, leaving behind a mixture enriched with the low vapor pressure propylene glycol, which would be vaporization ]resistant. Owing to this concern, a test was developed to compare the evaporation behavior of pure water, a 50/50 mixture of water and inhibited propylene glycol, and inhibited propylene glycol. The test was performed using room temperature fluids in an instrumented thin walled stainless steel vertical tube. The 1 in x 0.035 in wall tube was instrumented with surface thermocouples and encased in closed cell polyurethane foam. Reticulated polyurethane foam was placed inside the tube to reduce the convection currents. A vacuum system connected to the top of the tube set the pressure boundary condition. Tests were run for the three fluids at back pressures ranging from 1 to 18 torr. During each test, the mass of the test article was measured as it changed over time, as was its temperature and backpressure. The tests were successful. Somewhat surprisingly, the results showed that the evaporation behavior of the three fluids had more similarities than differences. The 50/50 mixture evaporated similarly to the pure water - albeit at a slower rate. The test results indicate that our extensive space - based experience with venting of single component fluids can be applied to the problem of Orion ATCS venting as long as the appropriate puts, takes, and caveats are applied.

Ungar, Eugene K.↗
NASA NTRS2011

Spacecraft Radiator Freeze Protection Using a Regenerative Heat Exchanger

An active thermal control system architecture has been modified to include a regenerative heat exchanger (regenerator) inboard of the radiator. Rather than using a radiator bypass valve a regenerative heat exchanger is placed inboard of the radiators. A regenerator cold side bypass valve is used to set the return temperature. During operation, the regenerator bypass flow is varied, mixing cold radiator return fluid and warm regenerator outlet fluid to maintain the system setpoint. At the lowest heat load for stable operation, the bypass flow is closed off, sending all of the flow through the regenerator. This lowers the radiator inlet temperature well below the system set-point while maintaining full flow through the radiators. By using a regenerator bypass flow control to maintain system setpoint, the required minimum heat load to avoid radiator freezing can be reduced by more than half compared to a radiator bypass system.

Ungar, Eugene K.↗
NASA NTRS2011

Impacts of an Ammonia Leak on the Cabin Atmosphere of the International Space Station

Toxic chemical release into the cabin atmosphere is one of the three major emergency scenarios identified on the International Space Station (ISS). The release of anhydrous ammonia, the coolant used in the U.S. On-orbit Segment (USOS) External Active Thermal Control Subsystem (EATCS), into the ISS cabin atmosphere is one of the most serious toxic chemical release cases identified on board ISS. The USOS Thermal Control System (TCS) includes an Internal Thermal Control Subsystem (ITCS) water loop and an EATCS ammonia loop that transfer heat at the interface heat exchanger (IFHX). Failure modes exist that could cause a breach within the IFHX. This breach would result in high pressure ammonia from the EATCS flowing into the lower pressure ITCS water loop. As the pressure builds in the ITCS loop, it is likely that the gas trap, which has the lowest maximum design pressure within the ITCS, would burst and cause ammonia to enter the ISS atmosphere. It is crucial to first characterize the release of ammonia into the ISS atmosphere in order to develop methods to properly mitigate the environmental risk. This paper will document the methods used to characterize an ammonia leak into the ISS cabin atmosphere. A mathematical model of the leak was first developed in order to define the flow of ammonia into the ISS cabin atmosphere based on a series of IFHX rupture cases. Computational Fluid Dynamics (CFD) methods were then used to model the dispersion of the ammonia throughout the ISS cabin and determine localized effects and ventilation effects on the dispersion of ammonia. Lastly, the capabilities of the current on-orbit systems to remove ammonia were reviewed and scrubbing rates of the ISS systems were defined based on the ammonia release models. With this full characterization of the release of ammonia from the USOS TCS, an appropriate mitigation strategy that includes crew and system emergency response procedures, personal protection equipment use, and atmosphere monitoring and scrubbing hardware can be established.

Duchesne, Stephanie M.↗
NASA NTRS2012

Impacts of an Ammonia Leak on the Cabin Atmosphere of the International Space Station

Toxic chemical release into the cabin atmosphere is one of the three major emergency scenarios identified on the International Space Station (ISS). The release of anhydrous ammonia, the coolant used in the U.S. On-orbit Segment (USOS) External Active Thermal Control Subsystem (EATCS), into the ISS cabin atmosphere is one of the most serious toxic chemical release cases identified on board ISS. The USOS Thermal Control System (TCS) includes an Internal Thermal Control Subsystem (ITCS) water loop and an EATCS ammonia loop that transfer heat at the interface heat exchanger (IFHX). Failure modes exist that could cause a breach within the IFHX. This breach would result in high pressure ammonia from the EATCS flowing into the lower pressure ITCS water loop. As the pressure builds in the ITCS loop, it is likely that the gas trap, which has the lowest maximum design pressure within the ITCS, would burst and cause ammonia to enter the ISS atmosphere. It is crucial to first characterize the release of ammonia into the ISS atmosphere in order to develop methods to properly mitigate the environmental risk. This paper will document the methods used to characterize an ammonia leak into the ISS cabin atmosphere. A mathematical model of the leak was first developed in order to define the flow of ammonia into the ISS cabin atmosphere based on a series of IFHX rupture cases. Computational Fluid Dynamics (CFD) methods were then used to model the dispersion of the ammonia throughout the ISS cabin and determine localized effects and ventilation effects on the dispersion of ammonia. Lastly, the capabilities of the current on-orbit systems to remove ammonia were reviewed and scrubbing rates of the ISS systems were defined based on the ammonia release models. With this full characterization of the release of ammonia from the USOS TCS, an appropriate mitigation strategy that includes crew and system emergency response procedures, personal protection equipment use, and atmosphere monitoring and scrubbing hardware can be established.

Duchesne, Stephanie M.↗
NASA NTRS2012

Numerical Study of Ammonia Leak and Dispersion in the International Space Station

Release of ammonia into the International Space Station (ISS) cabin atmosphere can occur if the water/ammonia barrier breach of the active thermal control system (ATCS) interface heat exchanger (IFHX) happens. After IFHX breach liquid ammonia is introduced into the water-filled internal thermal control system (ITCS) and then to the cabin environment through a ruptured gas trap. Once the liquid water/ammonia mixture exits ITCS, it instantly vaporizes and mixes with the U.S. Laboratory cabin air that results in rapid deterioration of the cabin conditions. The goal of the study is to assess ammonia propagation in the Station after IFHX breach to plan the operation procedure. A Computational Fluid Dynamics (CFD) model for accurate prediction of airflow and ammonia transport within each of the modules in the ISS cabin was developed. CFD data on ammonia content in the cabin aisle way of the ISS and, in particular, in the Russian On- Orbit Segment during the period of 15 minutes after gas trap rupture are presented for four scenarios of rupture response. Localized effects of ammonia dispersion and risk mitigation are discussed.

Son, Chang H.↗
NASA NTRS2012

Creating the Deep Space Environment for Testing the James Webb Space Telescope at the Johnson Space Center's Chamber A

Chamber A is the largest thermal vacuum chamber at the Johnson Space Center and is one of the largest space environment chambers in the world. The chamber is 19.8 m (65 ft) in diameter and 36.6 m (120 ft) tall and is equipped with cryogenic liquid nitrogen panels (shrouds) and gaseous helium shrouds to create a simulated space environment. It was originally designed and built in the mid 1960's to test the Apollo Command and Service Module and several manned tests were conducted on that spacecraft, contributing to the success of the program. The chamber has been used since that time to test spacecraft active thermal control systems, Shuttle DTO, DOD, and ESA hardware in simulated Low Earth Orbit (LEO) conditions. NASA is now moving from LEO towards exploration of locations with environments approaching those of deep space. Therefore, Chamber A has undergone major modifications to enable it to simulate these deeper space environments. Environmental requirements were driven, and the modifications were funded, by the James Webb Space Telescope program, and this telescope which will orbit Solar/Earth L2, will be the first test article to benefit from the chamber s new capabilities. To accommodate JWST, the Chamber A high vacuum system has been modernized, additional LN2 shrouds have been installed, the liquid nitrogen system has been modified to remove dependency on electrical power and increase its reliability, a new helium shroud/refrigeration system has been installed to create a colder more stable and uniform heat sink and, the controls have been updated to increase the level of automation and improve operator interfaces. Testing of these major modifications was conducted in August 2012 and this initial test was very successful, with all major systems exceeding their performance requirements. This paper will outline the changes in the overall environmental requirements, discuss the technical design data that was used in the decisions leading to the extensive modifications, and describe the new capabilities of the chamber.

Homan, Jonathan L.↗
NASA NTRS2012

Evaluating Three Active Thermal Architectures for Exploration Missions

Mass and cost are typically the two biggest challenges facing space craft designers. Active thermal control systems for crewed space-craft are typically among the more massive and costly systems on the vehicle. A study was completed evaluating three different thermal control system architectures to evaluate overall performance, mass and cost for a typical exploration mission profile. The architectures that were evaluated were 1 - a two-loop system using an internal liquid loop interfacing with an external liquid loop and flow loop with flow through radiators; 2 - a-single loop architecture with flow through radiators utilizing a regenerative heat exchanger and heater; and 3 - a single-loop architecture with heat pipe radiators. Environmental conditions, calculated for a given lunar exploration mission, and mission heat load profiles, generated based on previous Orion time lines, were evalauated through the phases of the on orbit mission. Performance for each of the architectures was evaluated along with the resultant mass of each system. Recommendations include adding a thermal topping system to lunar missions due to the extreme hot environments encountered in near-lunar approaches.

Cross, Cynthia D.↗
NASA NTRS2013

Creating the Deep Space Environment for Testing the James Webb Space Telescope (JWST) at NASA Johnson Space Center's Chamber A

Chamber A is the largest thermal vacuum chamber at the Johnson Space Center and is one of the largest space environment chambers in the world. The chamber is 19.8 m (65 ft) in diameter and 36.6 m (120 ft) tall and is equipped with cryogenic liquid nitrogen panels (shrouds) and gaseous helium shrouds to create a simulated space environment. It was originally designed and built in the mid 1960 s to test the Apollo Command and Service Module and several manned tests were conducted on that spacecraft, contributing to the success of the program. The chamber has been used since that time to test spacecraft active thermal control systems, Shuttle DTO, DOD, and ESA hardware in simulated Low Earth Orbit (LEO) conditions. NASA is now moving from LEO towards exploration of locations with environments approaching those of deep space. Therefore, Chamber A has undergone major modifications to enable it to simulate these deeper space environments. Environmental requirements were driven, and modifications were funded by the James Webb Space Telescope program, and this telescope which will orbit Solar/Earth L2, will be the first test article to benefit from the chamber s new capabilities. To accommodate JWST, the Chamber A high vacuum system has been modernized, additional LN2 shrouds have been installed, the liquid nitrogen system has been modified to remove dependency on electrical power and increase its reliability, a new helium shroud/refrigeration system has been installed to create a colder more stable and uniform heat sink, and the controls have been updated to increase the level of automation and improve operator interfaces. Testing of these major modifications was conducted in August of 2012 and this initial test was very successful, with all major systems exceeding their performance requirements. This paper will outline the changes in overall environmental requirements, discuss the technical design data that was used in the decisions leading to the extensive modifications, and describe the new capabilities of the chamber.

Homan, Jonathan L.↗
NASA NTRS2013

Creating the Deep Space Environment for Testing the James Webb Space Telescope at NASA Johnson Space Center's Chamber A

Chamber A is the largest thermal vacuum chamber at the Johnson Space Center and is one of the largest space environment chambers in the world. The chamber is 19.8 m (65 ft.) in diameter and 36.6 m (120 ft.) tall and is equipped with cryogenic liquid nitrogen panels (shrouds) and gaseous helium shrouds to create a simulated space environment. It was originally designed and built in the mid 1960 s to test the Apollo Command and Service Module and several manned tests were conducted on that spacecraft, contributing to the success of the program. The chamber has been used since that time to test spacecraft active thermal control systems, Shuttle DTO, DOD, and ESA hardware in simulated Low Earth Orbit (LEO) conditions. NASA is now moving from LEO towards exploration of locations with environments approaching those of deep space. Therefore, Chamber A has undergone major modifications to enable it to simulate these deeper space environments. Environmental requirements were driven, and modifications were funded by the James Webb Space Telescope program, and this telescope, which will orbit Solar/Earth L2, will be the first test article to benefit from the chamber s new capabilities. To accommodate JWST, the Chamber A high vacuum system has been modernized, additional LN2 shrouds have been installed, the liquid nitrogen system has been modified to minimize dependency on electrical power and increase its reliability, a new helium shroud/refrigeration system has been installed to create a colder more stable and uniform heat sink, and the controls have been updated to increase the level of automation and improve operator interfaces. Testing of these major modifications was conducted in August of 2012 and this initial test was very successful, with all major systems exceeding their performance requirements. This paper will outline the changes in overall environmental requirements, discuss the technical design data that was used in the decisions leading to the extensive modifications, and describe the new capabilities of the chamber.

Homan, Jonathan L.↗
NASA NTRS2014

MMOD Impact Damage to ISS

Paper will describe micrometeoroid and orbital debris (MMOD) damage that has been observed on the International Space Station (ISS). Several hundred documented MMOD damage sites on ISS have been identified through imagery from the windows of ISS modules or docked vehicles. Sites that are observable from ISS or shuttle windows exhibiting distinct features of hypervelocity impact damage are usually greater than 5mm in diameter. Many smaller features are revealed in on-orbit imagery are typically less distinct and difficult to characterize but could be MMOD damage. Additional images of on-orbit damage features have been collected by astronauts during extra vehicular activities. Ground inspection of returned ISS hardware has also contributed to the database of ISS MMOD impact damage. A handful of orbital replacement units (ORU) from the ISS active thermal control and electrical power subsystems were swapped out and returned during the Space Shuttle program. In addition, a reusable logistics module was deployed on ISS for a total 59.4 days on 11 shuttle missions between 2001 and 2011 and then brought back in the shuttle payload bay. All of this returned hardware was subjected to detailed post-flight inspections for MMOD damage, and a database with nearly 1000 impact records has been collected. A description of the largest observed damages will be provided in the paper. In addition, a discussion of significant MMOD impact sites with operational or design aspects will be presented. Some of the ISS modules/subsystems damaged by MMOD to be included in the discussion are (1) Solar Arrays, (2) US and Russian windows, (3) EVA handrails, (4) Radiators, and (5) Russian FGB module.

Hyde, James L.↗
NASA NTRS2014

Technology Maturation in Preparation for the Cryogenic Propellant Storage and Transfer (CPST) Technology Demonstration Mission (TDM)

In support of its goal to find an innovative path for human space exploration, NASA embarked on the Cryogenic Propellant Storage and Transfer (CPST) Project, a Technology Demonstration Mission (TDM) to test and validate key cryogenic capabilities and technologies required for future exploration elements, opening up the architecture for large in-space cryogenic propulsion stages and propellant depots. Recognizing that key Cryogenic Fluid Management (CFM) technologies anticipated for on-orbit (flight) demonstration would benefit from additional maturation to a readiness level appropriate for infusion into the design of the flight demonstration, the NASA Headquarters Space Technology Mission Directorate (STMD) authorized funding for a one-year technology maturation phase of the CPST project. The strategy, proposed by the CPST Project Manager, focused on maturation through modeling, concept studies, and ground tests of the storage and fluid transfer of CFM technology sub-elements and components that were lower than a Technology Readiness Level (TRL) of 5. A technology maturation plan (TMP) was subsequently approved which described: the CFM technologies selected for maturation, the ground testing approach to be used, quantified success criteria of the technologies, hardware and data deliverables, and a deliverable to provide an assessment of the technology readiness after completion of the test, study or modeling activity. The specific technologies selected were grouped into five major categories: thick multilayer insulation, tank applied active thermal control, cryogenic fluid transfer, propellant gauging, and analytical tool development. Based on the success of the technology maturation efforts, the CPST project was approved to proceed to flight system development.

propellants↗
NASA NTRS2014

Space Evaporator Absorber Radiator (SEAR) for Thermal Storage on Manned Spacecraft

Future manned exploration spacecraft will need to operate in challenging thermal environments. State-of the- art technology for active thermal control relies on sublimating water ice and venting the vapor overboard in very hot environments. This approach can lead to large loss of water and a significant mass penalty for the spacecraft. This paper describes an innovative thermal control system that uses a Space Evaporator Absorber Radiator (SEAR) to control spacecraft temperatures in highly variable environments without venting water. SEAR uses heat pumping and energy storage by LiCl/water absorption to enable effective cooling during hot periods and regeneration during cool periods. The LiCl absorber technology has the potential to absorb over 800 kJ per kg of system mass, compared to phase change heat sink systems that typically achieve approx. 50 kJ/kg. The optimal system is based on a trade-off between the mass of water saved and extra power needed to regenerate the LiCl absorber. This paper describes analysis models and the predicted performance and optimize the size of the SEAR system, estimated size and mass of key components, and power requirements for regeneration. We also present a concept design for an ISS test package to demonstrate operation of a subscale system in zero gravity.

Izenson, Michael G.↗
NASA NTRS2014

Cryogenic Boil-Off Reduction System Testing

Cryogenic propellants such as liquid hydrogen (LH2) and liquid oxygen (LO2) are a part of NASA's future space exploration due to the high specific impulse that can be achieved using engines suitable for moving 10's to 100's of metric tons of payload mass to destinations outside of low earth orbit. However, the low storage temperatures of LH2 and LO2 cause substantial boil-off losses for missions with durations greater than several days. The losses can be greatly reduced by incorporating high performance cryocooler technology to intercept heat load to the propellant tanks and by the integration of self-supporting multi-layer insulation. The active thermal control technology under development is the integration of the reverse turbo- Brayton cycle cryocooler to the propellant tank through a distributed cooling network of tubes coupled to a shield in the tank insulation and to the tank wall itself. Also, the self-supporting insulation technology was utilized under the shield to obtain needed tank applied LH2 performance. These elements were recently tested at NASA Glenn Research Center in a series of three tests, two that reduced LH2 boil-off and one to eliminate LO2 boil-off. This test series was conducted in a vacuum chamber that replicated the vacuum of space and the temperatures of low Earth orbit. The test results show that LH2 boil-off was reduced 60% by the cryocooler system operating at 90K and that robust LO2 zero boil-off storage, including full tank pressure control was achieved.

Propellant Storage↗
NASA NTRS2015

Space Evaporator Absorber Radiator (SEAR) for Thermal Storage on Manned Spacecraft

Future manned exploration spacecraft will need to operate in challenging thermal environments. State-of-the-art technology for active thermal control relies on sublimating water ice and venting the vapor overboard in very hot environments, and or heavy phase change material heat exchangers for thermal storage. These approaches can lead to large loss of water and a significant mass penalties for the spacecraft. This paper describes an innovative thermal control system that uses a Space Evaporator Absorber Radiator (SEAR) to control spacecraft temperatures in highly variable environments without venting water. SEAR uses heat pumping and energy storage by LiCl/water absorption to enable effective cooling during hot periods and regeneration during cool periods. The LiCl absorber technology has the potential to absorb over 800 kJ per kg of system mass, compared to phase change heat sink systems that typically achieve approx. 50 kJ/kg. This paper describes analysis models to predict performance and optimize the size of the SEAR system, estimated size and mass of key components, and an assessment of potential mass savings compared with alternative thermal management approaches. We also describe a concept design for an ISS test package to demonstrate operation of a subscale system in zero gravity.

Izenson, Michael G.↗
NASA NTRS2014

ISS Ammonia Pump Failure, Recovery, and Lesson Learned A Hydrodynamic Bearing Perspective

The design, development, and operation of long duration spaceflight hardware has become an evolutionary process in which meticulous attention to details and lessons learned from previous experiences play a critical role. Invaluable to this process is the ability to retrieve and examine spaceflight hardware that has experienced a premature failure. While these situations are rare and unfortunate, the failure investigation and recovery from the event serve a valuable purpose in advancing future space mechanism development. Such a scenario began on July 31, 2010 with the premature failure of an ammonia pump on the external active thermal control system of the International Space Station. The ground-based inspections of the returned pump and ensuing failure investigation revealed five potential bearing forces that were un-accounted for in the design phase and qualification testing of the pump. These forces could combine in a number of random orientations to overload the pump bearings leading to solid-surface contact, wear, and premature failure. The recovery plan identified one of these five forces as being related to the square of the operating speed of the pump and this fact was used to recover design life through a change in flight rules for the operation of the pump module. Through the course of the failure investigation, recovery, and follow-on assessment of pump wear life, design guidance has been developed to improve the life of future mechanically pumped thermal control systems for both human and robotic exploration missions.

Bruckner, Robert J.↗
NASA NTRS2016

Integrated Refrigeration and Storage for Advanced Liquid Hydrogen Operations

NASA has used liquefied hydrogen (LH2) on a large scale since the beginning of the space program as fuel for the Centaur and Apollo upper stages, and more recently to feed the three space shuttle main engines. The LH2 systems currently in place at the Kennedy Space Center (KSC) launch pads are aging and inefficient compared to the state-of-the-art. Therefore, the need exists to explore advanced technologies and operations that can drive commodity costs down, and provide increased capabilities. The Ground Operations Demonstration Unit for Liquid Hydrogen (GODU-LH2) was developed at KSC to pursue these goals by demonstrating active thermal control of the propellant state by direct removal of heat using a cryocooler. The project has multiple objectives including zero loss storage and transfer, liquefaction of gaseous hydrogen, and densification of liquid hydrogen. The key technology challenge was efficiently integrating the cryogenic refrigerator into the LH2 storage tank. A Linde LR1620 Brayton cycle refrigerator is used to produce up to 900W cooling at 20K, circulating approximately 22 g/s gaseous helium through the hydrogen via approximately 300 m of heat exchanger tubing. The GODU-LH2 system is fully operational, and is currently under test. This paper will discuss the design features of the refrigerator and storage system, as well as the current test results.

Liquefaction↗
NASA NTRS2018

The Demonstration of a Robotic External Leak Locator for the International Space Station

The International Space Station (ISS) and all currently conceivable future manned spacecraft are susceptible to mission impacts due to fluid/gas leaks to the exterior environment. For example, there is a well-known risk of ammonia leaks from the ISS External Active Thermal Control System (EATCS) loops and as of 2016 there was no method to locate them. It was, therefore, critical to develop a method for detecting and locating leaks to preserve vehicle health. The Robotic External Leak Locator (RELL) was developed and deployed to the ISS to provide this capability. An on-orbit validation and demonstration was successfully completed in December 2016 and leak locating operations occurred in February 2017. This paper discusses the results of these exercises including measurements of the environment around ISS, detection of a small ammonia leak and implementation of leak locating methodologies. RELL is a collaboration between NASA's Goddard Space Flight Center (GSFC) and Johnson Space Center (JSC) and was launched to the ISS as a Technology Demonstration Payload in December 2015 on Orbital-ATK Commercial Resupply Flight 4.

Naids, Adam J.↗