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

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NASA NTRS1999

Communications and Tracking of Visiting Vehicles near ISS: The Design of the Reusable Launch Vehicle Communications

The International Space Station (ISS) will provide for the visitation of various vehicles such as the Shuttle, Automated Transfer Vehicle (ATV), H-II Transfer Vehicle (HTV), Crew Return Vehicle (CRV), Reusable Launch Vehicle (RLV) and Soyuz. These vehicles will provide for crew replacement, consumables resupply, and equipment delivery. In order for these vehicles to approach and eventually dock with the ISS, there must be near continuous communications coverage between the visiting vehicle and the ISS, as well as communications between the vehicle and a Mission Control Center (MCC). Since the ISS communications systems are already designed and scheduled for ISS activation, the vehicles must either utilize these communications systems or provide their own. There are two means of two-way communications with the ISS. These are (1) S-Band communications using TDRSS, and (2) UHF communications using some form of the Space to Space Station Radio (SSSR) link. The RLV utilizes ISS compatible communications systems to communicate with both the ISS and a Mission Control Center. Since all vehicles must adhere to the Visiting Vehicle Interface requirements given in reference 1, the RLV communications system design must meet these requirements during entry into the ISS Approach Ellipsoid (AE) and during Proximity Operations. Included in this paper are descriptions of these communications approaches as well as their potential utilization in the ISS communications system.

Stillwagen, Frederic H.↗
NASA NTRS2002

Monitoring of Structural Integrity of Composite Structures by Embedded Optical Fiber Sensors

Real time monitoring of the mechanical integrity and stresses on key aerospace composite structures like aircraft wings, walls of pressure vessels and fuel tanks or any other structurally extended components and panels as in space telescopes is very important to NASA. Future military and commercial aircraft as well as NASA space systems such as Space Based Radar and International Space Station will incorporate a monitoring system to sense any degradation to the structure. In the extreme flight conditions of an aerospace vehicle it might be desirable to measure the strain every ten centimeters and thus fully map out the strain field of a composite component. A series of missions and vehicle health management requirements call for these measurements. At the moment thousands of people support a few vehicle launches per year. This number can be significantly reduced by implementing intelligent vehicles with integral nervous systems (smart structures). This would require maintenance to be performed only as needed. Military and commercial aircrafts have an equally compelling case. Maintenance yearly costs are currently reaching astronomical heights. Monitoring techniques are therefore required that allow for maintenance to be performed only when needed. This would allow improved safety by insuring that necessary tasks are performed while reducing costs by eliminating procedures that are costly and not needed. The advantages fiber optical sensors have over conventional electro-mechanical systems like strain gauges have been widely extolled in the research literature. These advantages include their small size, low weight, immunity to electrical resistance, corrosion resistance, compatibility with composite materials and process conditions, and multiplexing capabilities. One fiber optic device which is suitable for distributed sensing is the fiber Bragg grating (FBG). Researchers at NASA MSFC are currently developing techniques for using FBGs for monitoring the integrity of advanced structural materials expected to become the mainstay of the current and future generation space structures. Since carbon-epoxy composites are the materials of choice for the current space structures, the initial study is concentrated on this type of composite. The goals of this activity are to use embedded FBG sensors for measuring strain and temperature of composite structures, and to investigate the effects of various parameters such as composite fiber orientation with respect to the optical sensor, unidirectional fiber composite, fabrication process etc., on the optical performance of the sensor. This paper describes an experiment to demonstrate the use of an embedded FBG for measuring strain in a composite material. The performance of the fiber optic sensor is determined by direct comparison with results from more conventional instrumentation.

Albert J. Osei↗
NASA NTRS2003

Investigation of Structural Properties of Carbon-Epoxy Composites Using Embedded Fiber-Optic Bragg Gratings

Real time monitoring of the mechanical integrity and stresses on key aerospace composite structures like aircraft wings, walls of pressure vessels and fuel tanks or any other structurally extended components and panels as in space telescopes is very important to NASA. Future military and commercial aircraft as well as NASA space systems such as Space Based Radar and International Space Station will incorporate a monitoring system to sense any degradation to the structure. In the extreme flight conditions of an aerospace vehicle it might be desirable to measure the strain every ten centimeters and thus fully map out the strain field of a composite component. A series of missions and vehicle health management requirements call for these measurements. At the moment thousands of people support a few vehicle launches per year. This number can be significantly reduced by implementing intelligent vehicles with integral nervous systems (smart structures). This would require maintenance to be performed only as needed. Military and commercial aircrafts have an equally compelling case. Annual maintenance costs are currently reaching astronomical heights. Monitoring techniques are therefore required that allow for maintenance to be performed only when needed. This would allow improved safety by insuring that necessary tasks are performed while reducing costs by eliminating procedures that are costly and not needed. The advantages fiber optical sensors have over conventional electro-mechanical systems like strain gauges have been widely extolled in the research literature. These advantages include their small size, low weight, immunity to electrical resistance, corrosion resistance, compatibility with composite materials and process conditions, and multiplexing capabilities. One fiber optic device which is suitable for distributed sensing is the fiber Bragg grating (FBG). This is a periodic perturbation in the refractive index of the fiber core. When a broadband light is coupled into the optical fiber sensor, a reflection peak will be obtained centered around a wavelength called Bragg-wavelength. The Bragg-wavelength depends on the refractive index and the period of the grating, which both change due to mechanical and thermal strain applied to the sensor. The shift in the Bragg-wavelength is directly proportional to the strain. Researchers at NASA MSFC are currently developing techniques for using FBGs for monitoring the integrity of advanced structural materials expected to become the mainstay of the current and future generation space structures. Since carbon-epoxy composites are the materials of choice for the current space structures, the initial study is concentrated on this type of composite. The goals of this activity are to use embedded FBG sensors for measuring strain and temperature of composite structures, and to investigate the effects of various parameters such as composite fiber orientation with respect to the optical sensor, unidirectional fiber composite, fabrication process etc., on the optical performance of the sensor. This paper describes an experiment to demonstrate the use of an embedded FBG for measuring strain in a composite material. The performance of the fiber optic sensor is determined by direct comparison with results from more conventional instrumentation.

Osei, Albert J.↗
NASA NTRS2005

Multiple Payload Ejector for Education, Science and Technology Experiments

The education research community no longer has a means of being manifested on Space Shuttle flights, and small orbital payload carriers must be flown as secondary payloads on ELV flights, as their launch schedule, secondary payload volume and mass permits. This has resulted in a backlog of small payloads, schedule and cost problems, and an inability for the small payloads community to achieve routine, low-cost access to orbit. This paper will discuss Goddard's Wallops Flight Facility funded effort to leverage its core competencies in small payloads, sounding rockets, balloons and range services to develop a low cost, multiple payload ejector (MPE) carrier for orbital experiments. The goal of the MPE is to provide a low-cost carrier intended primarily for educational flight research experiments. MPE can also be used by academia and industry for science, technology development and Exploration experiments. The MPE carrier will take advantage of the DARPAI NASA partnership to perform flight testing of DARPA s Falcon small, demonstration launch vehicle. The Falcon is similar to MPE fiom the standpoint of focusing on a low-cost, responsive system. Therefore, MPE and Falcon complement each other for the desired long-term goal of providing the small payloads community with a low-cost ride to orbit. The readiness dates of Falcon and MPE are complementary, also. MPE is being developed and readied for flight within 18 months by a small design team. Currently, MPE is preparing for Critical Design Review in fall 2005, payloads are being manifested on the first mission, and the carrier will be ready for flight on the first Falcon demonstration flight in summer, 2006. The MPE and attached experiments can weigh up to 900 lb. to be compatible with Falcon demonstration vehicle lift capabilities fiom Wallops, and will be delivered to the Falcon demonstration orbit - 100 nautical mile circular altitude.

Lechworth, Gary↗
NASA NTRS2016

Conventional and Bimodal Nuclear Thermal Rocket (NTR) Artificial Gravity Mars Transfer Vehicle Concepts

A variety of countermeasures have been developed to address the debilitating physiological effects of zero-gravity (0-g) experienced by cosmonauts and astronauts during their approximately 0.5 to 1.2 year long stays in low Earth orbit (LEO). Longer interplanetary flights, combined with possible prolonged stays in Mars orbit, could subject crewmembers to up to approximately 2.5 years of weightlessness. In view of known and recently diagnosed problems associated with 0-g, an artificial gravity (AG) spacecraft offers many advantages and may indeed be an enabling technology for human flights to Mars. A number of important human factors must be taken into account in selecting the rotation radius, rotation rate, and orientation of the habitation module or modules. These factors include the gravity gradient effect, radial and tangential Coriolis forces, along with cross-coupled acceleration effects. Artificial gravity Mars transfer vehicle (MTV) concepts are presented that utilize both conventional NTR, as well as, enhanced bimodal nuclear thermal rocket (BNTR) propulsion. The NTR is a proven technology that generates high thrust and has a specific impulse (Isp) capability of approximately 900 s-twice that of today's best chemical rockets. The AG/MTV concepts using conventional Nuclear Thermal Propulsion (NTP) carry twin cylindrical International Space Station (ISS)- type habitation modules with their long axes oriented either perpendicular or parallel to the longitudinal spin axis of the MTV and utilize photovoltaic arrays (PVAs) for spacecraft power. The twin habitat modules are connected to a central operations hub located at the front of the MTV via two pressurized tunnels that provide the rotation radius for the habitat modules. For the BNTR AG/MTV option, each engine has its own closed secondary helium(He)-xenon (Xe) gas loop and Brayton Rotating Unit (BRU) that can generate 10s of kilowatts (kWe) of spacecraft electrical power during the mission coast phase eliminating the need for large PVAs. A single inflatable TransHab-type habitation module is also used with multiple vertical floors oriented radial to the MTV spin axis. The BNTR MTV's geometry-long and linear-is naturally compatible with AG operation. By rotating the vehicle about its center-of-mass (CM) and perpendicular to its flight vector at approximately 3.0 to 5.2 rpm, a centrifugal force and AG environment corresponding to approximately 0.38 to 1.0 g can be established to help maintain crew fitness out to Mars and back. Vehicles using NTP/ Bimodal Nuclear Thermal Propulsion (BNTP) can more readily accommodate the heavier payload mass and increased RCS propellant loading associated with AG operation, and can travel faster to and from Mars thereby reducing the crew's exposure to galactic cosmic radiation and solar flares. Mission scenario descriptions, key vehicle features and operational characteristics for each propulsion option are presented using the lift capability and payload volumes estimated for the Space Launch System (SLS)-1B and followon Heavy Lift Vehicle (HLV).

Nuclear Propulsion↗
NASA NTRS2011

Spacesuit and Space Vehicle Comparative Ergonomic Evaluation

With the advent of the latest manned spaceflight objectives, a series of prototype launch and reentry spacesuit architectures were evaluated for eventual down selection by NASA based on the performance of a set of designated tasks. A consolidated approach was taken to testing, concurrently collecting suit mobility data, seat-suit-vehicle interface clearances and movement strategies within the volume of a Multi-Purpose Crew Vehicle mockup. To achieve the objectives of the test, a requirement was set forth to maintain high mockup fidelity while using advanced motion capture technologies. These seemingly mutually exclusive goals were accommodated with the construction of an optically transparent and fully adjustable frame mockup. The mockup was constructed such that it could be dimensionally validated rapidly with the motion capture system. This paper will describe the method used to create a motion capture compatible space vehicle mockup, the consolidated approach for evaluating spacesuits in action, as well as the various methods for generating hardware requirements for an entire population from the resulting complex data set using a limited number of test subjects. Kinematics, hardware clearance, suited anthropometry, and subjective feedback data were recorded on fifteen unsuited and five suited subjects. Unsuited subjects were selected chiefly by anthropometry, in an attempt to find subjects who fell within predefined criteria for medium male, large male and small female subjects. The suited subjects were selected as a subset of the unsuited subjects and tested in both unpressurized and pressurized conditions. Since the prototype spacesuits were fabricated in a single size to accommodate an approximately average sized male, the findings from the suit testing were systematically extrapolated to the extremes of the population to anticipate likely problem areas. This extrapolation was achieved by first performing population analysis through a comparison of suited subjects performance to their unsuited performance and then applying the results to the entire range of population. The use of a transparent space vehicle mockup enabled the collection of large amounts of data during human-in-the-loop testing. Mobility data revealed that most of the tested spacesuits had sufficient ranges of motion for tasks to be performed successfully. A failed tasked by a suited subject most often stemmed from a combination of poor field of view while seated and poor dexterity of the gloves when pressurized or from suit/vehicle interface issues. Seat ingress/egress testing showed that problems with anthropometric accommodation does not exclusively occur with the largest or smallest subjects, but rather specific combinations of measurements that lead to narrower seat ingress/egress clearance.

England, Scott↗
NASA NTRSDate not supplied

Artemis IV Docking in Radiation Belt Charging Environment

NASA’s Artemis IV mission is planned to deliver the International Habitation Module (I-Hab) to the Gateway space station in lunar orbit. The I-Hab will be launched aboard the Space Launch System (SLS) vehicle as a co-manifested payload with the Orion spacecraft. After translunar injection begins, the Orion spacecraft will separate from the SLS Exploration Upper Stage (EUS) and then dock with the I-Hab to extract it from the EUS. Because of the altitude, orientation, and time for this to occur, the vehicle-to-vehicle potential between the I-Hab/EUS vehicle and Orion spacecraft could exceed several thousand volts. The docking of these two spacecrafts with such large differential potentials presents a challenge for the vehicles. This presentation describes the space environments at the docking altitude, the calculated vehicle-to-vehicle potential, and the possible impacts of the resulting voltage and current transients occurring at first contact. Additionally, possible risk mitigation tests to demonstrate compatibility with the transient current and voltages will be presented.

Matthew Mccollum↗
NASA NTRS1991

Reusable Reentry Satellite (RRS): Telemetry, Tracking, and Command (TT/C) Coverage Tradeoff Study

The Reusable Reentry Satellite (RRS) Telemetry, Tracking & Command (TT&C) Coverage Tradeoff Study described herein was performed during Part 1 of the RRS Phase B contract. This report is one of several that describes the results of various trade studies performed to arrive at a recommended design for the RRS satellite system. The overall RRS Phase B Study objective is to design a relatively inexpensive satellite to access space for extended periods of time, with eventual recovery of experiments on Earth. The RRS will be capable of: (1) being launched by a variety of expendable launch vehicles; (2) operating in low earth orbit as a free flying unmanned laboratory; and (3) executing an independent atmospheric reentry and soft landing. The RRS will be designed to be refurbished and reused up to three times a year for a period of 10 years. The expected principal use for such a system is research on the effects of variable gravity (0 to 1.5 g) and radiation on small animals, plants, lower life forms, tissue samples, and materials processes. This Summary Report provides a description of the RRS Telemetry, Tracking & Command Analysis performed to study the ground station coverage available to meet RRS requirements. Concepts were considered that used off-the-shelf technology, had generous margins of capability, and high reliability, were easily maintained, cost effective, and were compatible with existing support networks. From this study, a recommended RRS TT&C configuration will be developed.

Source record↗
NASA NTRS1999

Composites for Cryotank Structures. Present and Future: MSFC Perspective

The development of reusable launch vehicle systems for a single stage to orbit vehicle requires vehicles at liftoff with 85% to 94% of its mass consisting exclusively of propellants. These dry mass requirements drive designs to utilize stronger, lighter weight materials for structures. This technology development focus has allowed the introduction of composite materials in lieu of conventional metallic materials due to their higher specific strengths. Composite materials were successfully used for the liquid hydrogen tanks for the DC-XA, and a multi-lobed liquid hydrogen tank will be employed for the X-33. Another potential non-traditional application for composite materials is for liquid oxygen tanks, which is still being investigated. Traditionally, organic materials have been avoided wherever possible, due the potential fire hazard and the fact that composites fail conventional oxygen compatibility requirements. However, the potential weight savings warrant the investigation of the use of polymeric composite materials in oxygen environments. Since composites fail the conventional, time-proven test methods because they are considered flammable by test, we have embarked on an innovative approach to oxygen compatibility testing and evaluation focused on the use environments and attempts to eliminate or "design away" all potential ignition sources. Oxygen compatibility is defined as the ability of a material to coexist with oxygen and potential ignition sources with an acceptable, manageable degree of risk.

Vaughn, Timothy P.↗
NASA NTRS2017

NASA's In-Space Manufacturing Project: A Roadmap for a Multimaterial Fabrication Laboratory in Space

Human space exploration to date has been limited to low Earth orbit and the moon. The International Space Station (ISS) provides a unique opportunity for NASA to partner with private industry for development and demonstration of the technologies needed to support exploration initiatives. One challenge that is critical to sustainable and safer exploration is the ability to manufacture and recycle materials in space. This paper provides an overview of NASA's in-space manufacturing (ISM) project, its past and current activities (2014-2017), and how technologies under development will ultimately culminate in a multimaterial fabrication laboratory ("ISM FabLab") to be deployed on the International Space Station in the early 2020s. ISM is a critical capability for the long endurance missions NASA seeks to undertake in the coming decades. An unanticipated failure that can be adapted for in low earth orbit, through a resupply launch or a return to earth, may instead result in a loss of mission while in transit to Mars. To have a suite of functional ISM capabilities that are compatible with NASA's exploration timeline, ISM must be equipped with the resources necessary to develop these technologies and deploy them for testing prior to the scheduled de-orbit of ISS in 2024. The presentation provides a broad overview of ISM projects activities culminating with the Fabrication Laboratory for ISS. In 2017, the in-space manufacturing project issued a broad agency announcement for this capability. Requirements of the Fabrication Laboratory as stated in the solicitation will be discussed. The FabLab will move NASA and private industry significantly closer to changing historical paradigms for human spaceflight where all materials used in space are launched from earth. While the current ISM FabLab will be tested on ISS, future systems are eventually intended for use in a deep space habitat or transit vehicle. The work of commercial companies funded under NASA's Small Business Innovative Research Program (SBIR) is also discussed, as these activities, from development of recyclable packaging for ISS to additive manufacturing capabilities for metals and electronics, could also potentially be infused into future exploration capabilities. Key data from ISM projects to date will also be summarized.

Prater, Tracie↗
NASA NTRS2013

A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions

The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.

Mercer, Carolyn R.↗
NASA NTRS2014

A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions

The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.

Mercer, Carolyn R.↗
NASA NTRS2009

A Systems Approach to Lower Cost Missions: Following the Rideshare Paradigm

Small-satellite rideshare capabilities and opportunities for low-cost access to space have been evolving over the past 10 years. Small space launch vehicle technology is rapidly being developed and demonstrated, including the Minotaur series and the Space X Falcon, among others, along with the lower cost launch facilities at Alaska's Kodiak Launch Complex, NASA's Wallops Flight Facility, and the Reagan Test Site in the Pacific. Demonstrated capabilities for the launch of multiple payloads have increased (and continue to increase) significantly. This will allow more efficient and cost-effective use of the various launch opportunities, including utilizing the excess capacity of the emerging Evolved Expendable Launch Vehicle (EELV)-based missions. The definition of standardized interfaces and processes, along with various user guides and payload implementation plans, has been developed and continues to be refined. Top-level agency policies for the support of low-cost access to space for small experimental payloads, such as the DoD policy structure on auxiliary payloads, have been defined and provide the basis for the continued refinement and implementation of these evolving technologies. Most importantly, the coordination and cooperative interfaces between the various stakeholders continues to evolve. The degree of this coordination and technical interchange is demonstrated by the wide stakeholder participation at the recent 2008 Small Payload Rideshare Workshop, held at NASA's Wallops Flight Facility. This annual workshop has been the major platform for coordination and technical interchange within the rideshare community and with the various sponsoring agencies. These developments have provided the foundation for a robust low-cost small payload rideshare capability. However, the continued evolution, sustainment, and utilization of these capabilities will require continued stakeholder recognition, support, and nourishing. Ongoing, coordinated effort, partnering, and support between stakeholders is essential to acquire the improved organizational processes and efficiencies required to meet the needs of the growing small payload community for low-cost access to space. Further, a mix of capabilities developed within the space community for Operationally Responsive Space, an international committee investigating space systems cross-compatibility, and an industry-based organization seeking small satellite "standardization" all work toward a new paradigm: sharing or leveraging resources amongst multiple users. The challenge: where are those users, and what is the best way to leverage them? What is leveraged-mass, power, cost-sharing? And how does one sort through these options? What policies may prevent the use of some options? Who are the "other users" that might share or leverage capabilities? This paper presents a systematic look at both the users and the launch options, and suggests a way forward.

access to space↗
NASA NTRS1999

Upper-Stage Flight Experiment

For propulsion applications that require that the propellants are storable for long periods, have a high density impulse, and are environmentally clean and non-toxic, the best choice is a combination of high-concentration hydrogen peroxide (High Test Peroxide, or HTP) and a liquid hydrocarbon (LHC) fuel. The HTP/LHC combination is suitable for low-cost launch vehicles, space taxi and space maneuvering vehicles, and kick stages. Orbital Sciences Corporation is under contract with the NASA Marshall Space Flight Center in cooperation with the Air Force Research Lab to design, develop and demonstrate a new low-cost liquid upper stage based on HTP and JP-8. The Upper Stage Flight Experiment (USFE) focuses on key technologies necessary to demonstrate the operation of an inherently simple propulsion system with an innovative, state-of-the-art structure. Two key low-cost vehicle elements will be demonstrated - a 10,000 lbf thrust engine and an integrated composite tank structure. The suborbital flight test of the USFE is scheduled for 2001. Preceding the flight tests are two major series of ground tests at NASA Stennis Space Center and a subscale tank development program to identify compatible composite materials and to verify their compatibility over long periods of time. The ground tests include a thrust chamber development test series and an integrated stage test. This paper summarizes the results from the first phase of the thrust chamber development tests and the results to date from the tank material compatibility tests. Engine and tank configurations that meet the goals of the program are described.

Anderson, W. E.↗
NASA NTRS2012

Comparative Ergonomic Evaluation of Spacesuit and Space Vehicle Design

With the advent of the latest human spaceflight objectives, a series of prototype architectures for a new launch and reentry spacesuit that would be suited to the new mission goals. Four prototype suits were evaluated to compare their performance and enable the selection of the preferred suit components and designs. A consolidated approach to testing was taken: concurrently collecting suit mobility data, seat-suit-vehicle interface clearances, and qualitative assessments of suit performance within the volume of a Multi-Purpose Crew Vehicle mockup. It was necessary to maintain high fidelity in a mockup and use advanced motion-capture technologies in order to achieve the objectives of the study. These seemingly mutually exclusive goals were accommodated with the construction of an optically transparent and fully adjustable frame mockup. The construction of the mockup was such that it could be dimensionally validated rapidly with the motioncapture system. This paper describes the method used to create a space vehicle mockup compatible with use of an optical motion-capture system, the consolidated approach for evaluating spacesuits in action, and a way to use the complex data set resulting from a limited number of test subjects to generate hardware requirements for an entire population. Kinematics, hardware clearance, anthropometry (suited and unsuited), and subjective feedback data were recorded on 15 unsuited and 5 suited subjects. Unsuited subjects were selected chiefly based on their anthropometry in an attempt to find subjects who fell within predefined criteria for medium male, large male, and small female subjects. The suited subjects were selected as a subset of the unsuited medium male subjects and were tested in both unpressurized and pressurized conditions. The prototype spacesuits were each fabricated in a single size to accommodate an approximately average-sized male, so select findings from the suit testing were systematically extrapolated to the extremes of the population to anticipate likely problem areas. This extrapolation was achieved by first comparing suited subjects performance with their unsuited performance, and then applying the results to the entire range of the population. The use of a transparent space vehicle mockup enabled the collection of large amounts of data during human-in-the-loop testing. Mobility data revealed that most of the tested spacesuits had sufficient ranges of motion for the selected tasks to be performed successfully. A suited subject's inability to perform a task most often stemmed from a combination of poor field of view in a seated position, poor dexterity of the pressurized gloves, or from suit/vehicle interface issues. Seat ingress and egress testing showed that problems with anthropometric accommodation did not exclusively occur with the largest or smallest subjects, but also with specific combinations of measurements that led to narrower seat ingress/egress clearance.

England, Scott↗
NASA NTRS2016

Green Propellant Landing Demonstration at U.S. Range

The Green Propellant Loading Demonstration (GPLD) was conducted December 2015 at Wallops Flight Facility (WFF), leveraging work performed over recent years to bring lower toxicity hydrazine replacement green propellants to flight missions. The objective of this collaboration between NASA Goddard Space Flight Center (GSFC), WFF, the Swedish National Space Board (SNSB), and Ecological Advanced Propulsion Systems (ECAPS) was to successfully accept LMP-103S propellant at a U.S. Range, store the propellant, and perform a simulated flight vehicle propellant loading. NASA GSFC Propulsion (Code 597) managed all aspects of the operation, handling logistics, preparing the procedures, and implementing the demonstration. In addition to the partnership described above, Moog Inc. developed an LMP-103S propellant-compatible titanium rolling diaphragm flight development tank and loaned it to GSFC to act as the GPLD flight vessel. The flight development tank offered the GPLD an additional level of flight-like propellant handling process and procedures. Moog Inc. also provided a compatible latching isolation valve for remote propellant expulsion. The GPLD operation, in concert with Moog Inc. executed a flight development tank expulsion efficiency performance test using LMP-103S propellant. As part of the demonstration work, GSFC and WFF documented Range safety analyses and practices including all elements of shipping, storage, handling, operations, decontamination, and disposal. LMP-103S has not been previously handled at a U.S. Launch Range. Requisite for this activity was an LMP-103S Risk Analysis Report and Ground Safety Plan. GSFC and WFF safety offices jointly developed safety documentation for application into the GPLD operation. The GPLD along with the GSFC Propulsion historical hydrazine loading experiences offer direct comparison between handling green propellant versus safety intensive, highly toxic hydrazine propellant. These described motives initiated the GPLD operation in order to investigate the handling and process safety variances in project resources between LMP-103S and typical in-space propellants. The GPLD risk reduction operation proved successful for many reasons including handling the green propellant at a U.S. Range, loading and pressurizing a flight-like tank, expelling the propellant, measuring the tank expulsion efficiency, and most significantly, GSFC propulsion personnel's new insight into the LMP-103S propellant handling details.

Propellant↗
NASA NTRS2016

Green Propellant Loading Demonstration at U.S. Range

The Green Propellant Loading Demonstration (GPLD) was conducted December 2015 at Wallops Flight Facility (WFF), leveraging work performed over recent years to bring lower toxicity hydrazine replacement green propellants to flight missions. The objective of this collaboration between NASA Goddard Space Flight Center (GSFC), WFF, the Swedish National Space Board (SNSB), and Ecological Advanced Propulsion Systems (ECAPS) was to successfully accept LMP-103S propellant at a U.S. Range, store the propellant, and perform a simulated flight vehicle propellant loading. NASA GSFC Propulsion (Code 597) managed all aspects of the operation, handling logistics, preparing the procedures, and implementing the demonstration. In addition to the partnership described above, Moog Inc. developed an LMP-103S propellant-compatible titanium rolling diaphragm flight development tank and loaned it to GSFC to act as the GPLD flight vessel. The flight development tank offered the GPLD an additional level of flight-like propellant handling process and procedures. Moog Inc. also provided a compatible latching isolation valve for remote propellant expulsion. The GPLD operation, in concert with Moog Inc. executed a flight development tank expulsion efficiency performance test using LMP-103S propellant. As part of the demonstration work, GSFC and WFF documented Range safety analyses and practices including all elements of shipping, storage, handling, operations, decontamination, and disposal. LMP-103S has not been previously handled at a U.S. Launch Range. Requisite for this activity was an LMP-103S Risk Analysis Report and Ground Safety Plan. GSFC and WFF safety offices jointly developed safety documentation for application into the GPLD operation. The GPLD along with the GSFC Propulsion historical hydrazine loading experiences offer direct comparison between handling green propellant versus safety intensive, highly toxic hydrazine propellant. These described motives initiated the GPLD operation in order to investigate the handling and process safety variances in project resources between LMP-103S and typical in-space propellants. The GPLD risk reduction operation proved successful for many reasons including handling the green propellant at a U.S. Range, loading and pressurizing a flight-like tank, expelling the propellant, measuring the tank expulsion efficiency, and most significantly, GSFC propulsion personnel's new insight into the LMP-103S propellant handling details.

Green Propellant↗
NASA NTRS2025

Evaluating the Viability of Compact and Portable X-Ray Systems for an Exploration Medical System in a Ground Demonstration

MOTIVATION FOR INCLUDING X-RAY CAPABILITIES For upcoming exploration missions, the need for enhanced medical care becomes critical due to extended mission durations, significant communication delays, and minimal evacuation opportunities. Previous evidence by our team has revealed that among the 119 medical conditions targeted for management during spaceflight within NASA Exploration Medical Capability’s IMPACT Condition List, at least 36 could benefit from radiography (XR). Utilizing XR for diagnosis and management is hypothesized to significantly improve management of crew health by enabling the immediate evaluation and confirmation of potential injuries or illnesses. Beyond clinical applications, XR also holds potential for non-destructive testing (NDT). This includes applications such as assessing the structural integrity of the spacecraft, analyzing surface and meteorite samples, and inspecting onboard electronics. THREE CANDIDATE X-RAY SYSTEMS CHOSEN FOR GROUND DEMONSTRATION The Exploration Medical Capability Element (ExMC) and the Exploration Medical Integrated Product Team (XMIPT) of the Mars Campaign Office initiated early background work for ground demonstrations. In FY21, ExMC published a Concept of Operations to guide requirements development. By FY23, XMIPT and yet2, a technology scouting and open innovation consulting firm, had completed a market survey and trade study to identify potential miniature XR systems. Selection criteria included commercial-off-the-shelf availability, low mass and volume, and regulatory compliance. The top three candidate devices—Remedi REMEX-KA6, MinXray Impact, and FujiFilm Xair—were acquired to characterize the requirements and capabilities of each device. To facilitate testing, phantoms, and radiographic personal protective equipment (PPE) were purchased, and a dedicated space was designated for XRS usage at Glenn Research Center. During this presentation, the mass, volume, and power requirements for each of the three piloted devices are revealed, as well as information regarding the detector, mA, and kV of the devices. GOAL AND OBJECTIVES OF A MINI XRS GROUND DEMONSTRATION The primary goal of ExMC/XMIPT technology demonstrations is to bridge the gap in available, flight-ready medical device technology by flight-testing diagnostic and treatment technologies essential for managing medical conditions during exploration missions. These technologies must adhere to vehicle constraints such as mass, volume, power, and data requirements, integrate seamlessly with medical decision-support tools, and support increasingly Earth-independent operations. There are three main objectives for the future ground demonstration of these three devices. First, we aim to determine the full capabilities of these three miniature XR systems within the context of the spaceflight environment. While medical applications are the primary focus for the miniature XR, a comprehensive exploration of non-medical uses has been initiated by an XMIPT-sponsored NASA SPARK campaign to identify collaborators. Second, we plan to establish criteria and to use insights gained from evaluating each miniature XR against those criteria to select the most suitable system among the three candidates. Third, we intend to evaluate their suitability for flight certification, which includes assessing its durability for launch, reentry, and exposure to high background radiation, as well as its compatibility with existing data architecture systems. Numerous subject matter experts from NASA and partner institutions will support these objectives.

C A Haddix↗