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Artemis is a Common Lunar Lander (CLL) design for the Space Exploration Initiative (SEI). Structure factors for the CLL's primary and secondary structures are listed in tabular form. Additionally, engineering drawings of various systems, including the propulsion and landing systems, are presented.
International Space Station (ISS) crew and flight controller training documentation is used to aid in training operations. The Generic Simulations References SharePoint (Gen Sim) site is a database used as an aid during flight simulations. The Gen Sim site is used to make individual mission segment timelines, data, and flight information easily accessible to instructors. The Waste and Hygiene Compartment (WHC) training schematic includes simple and complex fluid schematics, as well as overall hardware locations. It is used as a teaching aid during WHC lessons for both ISS crew and flight controllers. ISS flight control documentation is used to support all aspects of ISS mission operations. The Quick Look Database and Consolidated Tool Page are imagery-based references used in real-time to help the Operations Support Officer (OSO) find data faster and improve discussions with the Flight Director and Capsule Communicator (CAPCOM). A Quick Look page was created for the Permanent Multipurpose Module (PMM) by locating photos of the module interior, labeling specific hardware, and organizing them in schematic form to match the layout of the PMM interior. A Tool Page was created for the Maintenance Work Area (MWA) by gathering images, detailed drawings, safety information, procedures, certifications, demonstration videos, and general facts of each MWA component and displaying them in an easily accessible and consistent format. Participation in ISS mechanisms and maintenance lessons, mission simulation On-the-Job Training (OJT), and real-time flight OJT was used as an opportunity to train for day-to-day operations as an OSO, as well as learn how to effectively respond to failures and emergencies during mission simulations and real-time flight operations.
The performances of three existing high pressure oxygen mechanical impact test systems were tested at two different test sites. The systems from one test site were fabricated from the same design drawing, whereas the system tested at the other site was of different design. Energy delivered to the test sample for each test system was evaluated and compared. Results were compared to the reaction rates obtained.
A mobile remote manipulator system is disclosed for assembly, repair and logistics transport on, around and about a space station square bay truss structure. The vehicle is supported by a square track arrangement supported by guide pins integral with the space station truss structure and located at each truss node. Propulsion is provided by a central push-pull drive mechanism that extends out from the vehicle one full structural bay over the truss and locks drive rods into the guide pins. The draw bar is now retracted and the mobile remote manipulator system is pulled onto the next adjacent structural bay. Thus, translation of the vehicle is inchworm style. The drive bar can be locked onto two guide pins while the extendable draw bar is within the vehicle and then push the vehicle away one bay providing bidirectional push-pull drive. The track switches allow the vehicle to travel in two orthogonal directions over the truss structure which coupled with the bidirectional drive, allow movement in four directions on one plane. The top layer of this trilayered vehicle is a logistics platform. This platform is capable of 369 degees of rotation and will have two astronaut foot restraint platforms and a space crane integral.
Non-optimum factors are used during aerospace conceptual and preliminary design to account for the increased weights of as-built structures due to future manufacturing and design details. Use of higher-fidelity non-optimum factors in these early stages of vehicle design can result in more accurate predictions of a concept s actual weights and performance. To help achieve this objective, non-optimum factors are calculated for the aluminum-alloy gores that compose the ogive and ellipsoidal bulkheads of the Space Shuttle Super-Lightweight Tank propellant tanks. Minimum values for actual gore skin thicknesses and weld land dimensions are extracted from selected production drawings, and are used to predict reference gore weights. These actual skin thicknesses are also compared to skin thicknesses predicted using classical structural mechanics and tank proof-test pressures. Both coarse and refined weights models are developed for the gores. The coarse model is based on the proof pressure-sized skin thicknesses, and the refined model uses the actual gore skin thicknesses and design detail dimensions. To determine the gore non-optimum factors, these reference weights are then compared to flight hardware weights reported in a mass properties database. When manufacturing tolerance weight estimates are taken into account, the gore non-optimum factors computed using the coarse weights model range from 1.28 to 2.76, with an average non-optimum factor of 1.90. Application of the refined weights model yields non-optimum factors between 1.00 and 1.50, with an average non-optimum factor of 1.14. To demonstrate their use, these calculated non-optimum factors are used to predict heavier, more realistic gore weights for a proposed heavy-lift launch vehicle s propellant tank bulkheads. These results indicate that relatively simple models can be developed to better estimate the actual weights of large structures for future launch vehicles.
During aerospace vehicle conceptual and preliminary design, empirical non-optimum factors are typically applied to predicted structural component weights to account for undefined manufacturing and design details. Non-optimum factors are developed here for 32 aluminum-lithium 2195 orthogrid panels comprising the liquid hydrogen tank barrel of the Space Shuttle External Tank using measured panel weights and manufacturing drawings. Minimum values for skin thickness, axial and circumferential blade stiffener thickness and spacing, and overall panel thickness are used to estimate individual panel weights. Panel non-optimum factors computed using a coarse weights model range from 1.21 to 1.77, and a refined weights model (including weld lands and skin and stiffener transition details) yields non-optimum factors of between 1.02 and 1.54. Acreage panels have an average 1.24 non-optimum factor using the coarse model, and 1.03 with the refined version. Non-optimum factors are also calculated for the aluminum-alloy gores that compose the ogive and ellipsoidal bulkheads of the Space Shuttle Super-Lightweight Tank propellant tanks. Minimum values for actual gore skin thicknesses and weld land dimensions are extracted from component production drawings, and are used to predict reference gore weights. These actual skin thicknesses are also compared to skin thicknesses predicted using classical structural mechanics and tank proof-test pressures. Both coarse and refined weights models are developed for the gores. The coarse model is based on the proof-test skin thicknesses, and the refined model uses the actual skin thicknesses and weld land details. Both models include manufacturing tolerance weight estimates. To determine the gore non-optimum factors, these reference weights are then compared to flight hardware weights reported in a mass properties database. The gore non-optimum factors computed using the coarse weights model range from 1.18 to 2.36, with an average non-optimum factor of 1.68. Application of the refined weights model yields non-optimum factors between 1.00 and 1.45, with an average non-optimum factor of 1.13. To demonstrate their utility, these calculated non-optimum factors are then used to predict more realistic (i.e., heavier) gore weights for a proposed heavy-lift launch vehicle's propellant tank bulkheads.
The mechanics of the oblique impingement of an axisymmetric jet on a plane surface are examined in detail. The stagnation point is discussed. A schematic drawing of the problem and coordinate system used to describe the flow field are given. The kinematic features of the flow above the plate are examined in the context of the conservation of mass, the vorticity of the jet, and the vorticity introduced by the jetplate interaction. The dynamic features of the flow are examined in terms of the surface pressure distribution and the cause-effect relationships which exist between the pressure and velocity/vorticity distributions. Flow calculations performed are given. The investigation is relevant to the flow resulting from the interaction of the propulsion jet with the main airfoil (STOL aircraft), and is appropriate to an over- or under- wing configuration.
The teleoperator being developed to augment the Space Transportation System (STS) for satellite placement, retrieval, or servicing at altitudes or orbital planes where it would be impractical to use the shuttle is primarily a general purpose propulsion stage that can be fitted with manipulator arms, automated servicers and satellite retrieval devices for particular missions. Design concepts for a general purpose retrieval device for docking with a satellite to which a grappling fixture has been attached, and for a retrieval device for docking with the Solar Maximum Mission (SMM) spacecraft were defined. The mechanical aspects of these two devices are discussed as well as the crew operations involved and problems created by the requirement for remote control. Drawings for the two retrieval device concepts are included.
A Kalman filter was designed to yield optimal estimates of geophysical parameters from Very Long Baseline Interferometry (VLBI) group delay data. The geophysical parameters are the polar motion components, adjustments to nutation in obliquity and longitude, and a change in the length of day parameter. The VLBI clock (and clock rate) parameters and atmospheric zenith delay parameters are estimated simultaneously. Filter background is explained. The IRIS (International Radio Interferometric Surveying) VLBI data are Kalman filtered. The resulting polar motion estimates are examined. There are polar motion signatures at the times of three large earthquakes occurring in 1984 to 1986: Mexico, 19 September, 1985 (Magnitude M sub s = 8.1); Chile, 3 March, 1985 (M sub s = 7.8); and Taiwan, 14 November, 1986 (M sub s = 7.8). Breaks in polar motion occurring about 20 days after the earthquakes appear to correlate well with the onset of increased regional seismic activity and a return to more normal seismicity (respectively). While the contribution of these three earthquakes to polar motion excitations is small, the cumulative excitation due to earthquakes, or seismic phenomena over a Chandler wobble damping period may be significant. Mechanisms for polar motion excitation due to solid earth phenomena are examined. Excitation functions are computed, but the data spans are too short to draw conclusions based on these data.
Modifications to the ground plane, to insure a good electrical return path during the pulse discharge, were made using a ring of beryllium copper finger stock attached to the underside of the aluminum ground plate. Experiments on annealing of wafers with ion implantation damage continued. The entire surface of 100 mm diameter wafers were annealed by one pulse for the standard implant (10 keV, phosphorus, 2x10 to the 15th power ions/sq cm). While samples are being fabricated into solar cells for electrical characterization, work is continuing on improvement of the electron beam uniformity and the optimization of the diode parameters. The engineering design was completed and the manufacturing detail drawings were released for fabrication. Assembly of the subcomponents for the exit and entrance locks is almost complete. These components include the cassettes, the indexing mechanisms, main doors, and wafer carrier transfer modules. The 'Y' track and three phase transition track sections are under final assembly and test.
To support development of the Boeing-Rocketdyne RS84 rocket engine, a fill-flow, reaction turbine geometry was integrated into the NASA-MSFC turbine air-flow test facility. A mechanical design was generated which minimized the amount of new hardware while incorporating all test and instrUmentation requirements. This paper provides details of the mechanical design for this Turbine Air-Flow Task (TAFT) test rig. The mechanical design process utilized for this task included the following basic stages: Conceptual Design. Preliminary Design. Detailed Design. Baseline of Design (including Configuration Control and Drawing Revision). Fabrication. Assembly. During the design process, many lessons were learned that should benefit future test rig design projects. Of primary importance are well-defined requirements early in the design process, a thorough detailed design package, and effective communication with both the customer and the fabrication contractors. The test rig provided steady and unsteady pressure data necessary to validate the computational fluid dynamics (CFD) code. The rig also helped characterize the turbine blade loading conditions. Test and CFD analysis results are to be presented in another JANNAF paper.
To support development of the Boeing-Rocketdyne RS84 rocket engine, a full-flow, reaction turbine geometry was integrated into the NASA-MSFC turbine air-flow test facility. A mechanical design was generated which minimized the amount of new hardware while incorporating all test and instrumentation requirements. This paper provides details of the mechanical design for this Turbine Air-Flow Task (TAFT) test rig. The mechanical design process utilized for this task included the following basic stages: Conceptual Design. Preliminary Design. Detailed Design. Baseline of Design (including Configuration Control and Drawing Revision). Fabrication. Assembly. During the design process, many lessons were learned that should benefit future test rig design projects. Of primary importance are well-defined requirements early in the design process, a thorough detailed design package, and effective communication with both the customer and the fabrication contractors.
Successfully communicating the complex risks associated with radiation exposure is a difficult undertaking; communicating those risks within the high-risk context of space travel is uniquely challenging. Since the potential risks of space radiation exposure are not expected to be realized until much later in life, it is hard to draw comparisons between other spaceflight risks such as hypoxia and microgravity-induced bone loss. Additionally, unlike other spaceflight risks, there is currently no established mechanism to mitigate the risks of incurred radiation exposure such as carcinogenesis. Despite these challenges, it is the duty of the Space Radiation Analysis Group (SRAG) at NASA's Johnson Space Center to provide astronauts with the appropriate information to effectively convey the risks associated with exposure to the space radiation environment. To this end, astronauts and their flight surgeons are provided with an annual radiation risk report documenting the astronaut's individual radiation exposures from space travel, medical, and internal radiological procedures throughout the astronaut's career. In an effort to improve this communication and education tool, this paper critically reviews the current report style and explores alternative report styles to define best methods to appropriately communicate risk to astronauts, flight surgeons, and management.
The International Space Station (ISS) is currently equipped with a complex, heavy, and power consuming system that recovers approximately 50% of O2 from metabolic CO2. Future long duration missions will require a sustainable and highly efficient system capable of yielding a minimum of 75% O2 recovery. A Macrofluidic Electrochemical Reactor (MFECR) technology development effort is currently underway at NASA Marshall Space Flight Center (MSFC) to significantly increase current O2 recovery efficiency and reduce complexity of the system. This paper presents a comprehensive multi-physic 3D model developed at MSFC on CO2 conversion to O2 and C2H4 at standard conditions via MFECR. The 3D spatial domain of the model is a replica of the actual MFECR’s 3D drawing generated for the MFECR fabrication and operated to recover O2 from CO2 yielding C2H4 as byproduct. Electrochemical (EC) physics that includes EC multicomponent reaction mechanisms, mass transport, and current density distributions is coupled in the model with all the other physics phenomena involved in the process, such as free and porous fluid flow, multicomponent mass transfer, heat transfer, and DC electrical current generation along with Joule heating effect. The authors plan to use experimental results to validate this comprehensive and rigorous model and build a reliable simulator that will not only assist the authors on the MFECR design but also optimize its operation.
The International Space Station (ISS) is currently equipped with a complex, heavy, and power consuming system that recovers approximately 50% of O2 from metabolic CO2. Future long duration human missions to the Moon and Mars will necessitate a sustainable and highly efficient metabolic oxygen recovery system capable of yielding a minimum of 75% O2 recovery. A Macrofluidic Electrochemical Reactor (MFECR) technology development effort is currently underway at NASA Marshall Space Flight Center (MSFC) to significantly increase current metabolic O2 recovery efficiency, expand mission sustainability, and reduce complexity of the system. The novel design combines CO2 conversion to O2 along with C2H4 as byproduct and water electrolysis (currently conducted in two separate units) into a single compact unit that runs at standard conditions and is theoretically capable of generating O2 with a theoretical maximum metabolic CO2 conversion of 73% while consuming less than metabolic water. This paper presents a comprehensive multi-physic 3D model developed at MSFC on CO2 conversion to O2 and C2H4 at standard conditions via MFECR. The 3D spatial domain of the model is a replica of the actual MFECR’s 3D drawing generated for the MFECR fabrication and operated to recover O2 from CO2 yielding C2H4 as byproduct. Electrochemical (EC) physics that includes EC multicomponent reaction mechanisms, mass transport, and electrical current density distributions is coupled in the model with all the other physics phenomena involved in the MFECR’s process, such as two-phase flow, free and porous fluid regimes, multicomponent mass transfer, heat transfer, and DC electrical current generation along with Joule heating effect. The EC reaction sections of the MFECR consists of two porous gas diffusion electrodes (GDE) and an electrolyte serpentine channel sandwiched in the middle. The CO2 feeds the cathode serpentine channel and part of the O2 product is fed back to the anode serpentine chamber. An alkaline solution feeds the electrolyte serpentine chamber wetting the GDEs of both, the anode and cathode allowing the OH- ionic transport between them. The EC reactions in the cathode’s GDE yield C2H4 from CO2 and H2 from water while the EC reaction in the anode’s GDE yields O2. The authors will present in this paper the validation of the model using experimental data and the utilization of the validated model in building a reliable simulator that will not only assist the authors on the MFECR design but also the optimization of its operation in the ISS and future spatial human missions.
The design and testing of a heat pipe for spacecraft application is presented. The application in mind calls for heat loads up to 20 watts, a set-point temperature of 294K, and a sink that varies from -220K to nearly as high as the set-point. The overall heat pipe length is 137 cm. Two basically different mechanisms of achieving variable conductance in the pipe by vapor-flow throttling were studied. In one, the thermal resistance between the heat source and sink is due to a saturation-temperature drop corresponding to the vapor-pressure drop developed across the valve. In the other, the pressure difference across the valve induces capillary groove and wick dry out in an evaporation region, and thus results in an increased thermal resistance. This mechanism was selected for fabrication and testing. The pipe is a stainless-steel/methanol two-heat-pipe system. Results are presented and discussed. Engineering drawings and specifications of the pipe are shown.
The effect of various pre-treatments on the elastic strain coefficient (alpha) (defined as the reciprocal of the modulus of elasticity E) (Epsilon) and on the mechanical hysteresis of copper has been investigated. Variables comprising the pre-treatments were pre-straining by stretching in a tensile testing machine and by drawing through a die, aging at room and elevated temperatures and annealing. The variation of the elastic strain coefficient with test stress was also investigated.
This paper explores the novel Strayton engine concept. This engine combines the cycles of a Brayton engine with that of a Stirling engine to create a highly efficient recuperating gas turbine engine. In the explored case, both Brayton cycle and Stirling cycle engines are used to generate electrical power. Additionally, the Stirling engine is used to draw heat out of the Brayton turbine (acting to cool the turbine blades), while also pumping heat into Brayton cycle just before combustion occurs (acting as the mechanism for recuperation). The purpose of this paper is to detail the system level modeling techniques used to generate the simulation, perform a cycle analysis of the combined cycle engine, identify key technologies and challenges associated with the concept, and compare potential performance gains with existing gas turbine engines and internal combustion engines. Topics such as controls, blade cooling effects, engine weight, and heat transfer using heat pipe are also explored. Results from this work show potential architectures that could provide the required heat transfer rates, potential control strategies, and performance benefits, including efficiency gains between 10% and 3% on engines ranging from 200HP to 670HP with the combined cycle engine.