SEARCH · Search NASA
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.
Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.
The aircraft exhaust engine nozzle system includes a fan nozzle to receive a fan flow from a fan disposed adjacent to an engine disposed above an airframe surface of the aircraft, a core nozzle disposed within the fan nozzle and receiving an engine core flow, and a pylon structure connected to the core nozzle and structurally attached with the airframe surface to secure the engine to the aircraft.
In 2010, the Robonaut Project embarked on a multi‐phase mission to perform technology demonstrations on‐board the International Space Station (ISS), showcasing state of the art robotics technologies through the use of Robonaut 2 (R2). This phased approach implements a strategy that allows for the use of ISS as a test bed during early development to both demonstrate capability and test technology while still making advancements in the earth based laboratories for future testing and operations in space. While R2 was performing experimental trials onboard the ISS during the first phase, engineers were actively designing for Phase 2, Intra‐Vehicular Activity (IVA) Mobility, that utilizes a set of zero‐g climbing legs outfitted with grippers to grasp handrails and seat tracks. In addition to affixing the new climbing legs to the existing R2 torso, it became clear that upgrades to the torso to both physically accommodate the climbing legs and to expand processing power and capabilities of the robot were required. In addition to these upgrades, a new safety architecture was also implemented in order to account for the expanded capabilities of the robot. The IVA climbing legs not only needed to attach structurally to the R2 torso on ISS, but also required power and data connections that did not exist in the upper body. The climbing legs were outfitted with a blind mate adapter and coarse alignment guides for easy installation, but the upper body required extensive rewiring to accommodate the power and data connections. This was achieved by mounting a custom adapter plate to the torso and routing the additional wiring through the waist joint to connect to the new set of processors. In addition to the power and data channels, the integrated unit also required updated electronics boards, additional sensors and updated processors to accommodate a new operating system, software platform, and custom control system. In order to perform the unprecedented task of building a robot in space, extensive practice sessions and meticulous procedures were required. Since crew training time is at a premium, the R2 team took a skills‐based training approach to ensure the astronauts were proficient with a basic skill set while refining the detailed procedures over several practice sessions and simulations. In addition to the crew activities, meticulous ground procedures were required in order to upgrade firmware on the upper body motor drivers. The new firmware for the IVA mobility unit needed to be deployed using the old software system. This also provided an opportunity to upgrade the upper body joints with new software and allowed for limited insight into the success of the updates. Complete verification that the updated firmware was successfully loaded was not confirmed until the rewiring of the upper body torso was complete.
A system, method, and apparatus for a vented launch vehicle adaptor (LVA) for a manned spacecraft with a "pusher" launch abort system are disclosed. The disclosed LVA provides a structural interface between a commercial crew vehicle (CCV) crew module/service module (CM/SM) spacecraft and an expendable launch vehicle. The LVA provides structural attachment of the module to the launch vehicle. It also provides a means to control the exhaust plume from a pusher-type launch abort system that is integrated into the module. In case of an on-pad or ascent abort, which requires the module to jettison away from the launch vehicle, the launch abort system exhaust plume must be safely directed away from critical and dangerous portions of the launch vehicle in order to achieve a safe and successful jettison.
The NASA Docking System (NDS) Block 1 is a key component of NASA's vision for space exploration. It is designed to provide capability for visiting vehicles to dock to the International Space Station's recently-installed International Docking Adapter ports. It is the first docking system to be developed by NASA since the Apollo-Soyuz Test Project of the 1970's. The NDS Block 1 includes provisions for capture, structural attachment, power/data transfer, and undocking. It uses a direct-drive electromechanical Stewart Platform capture system architecture, along with an innovative automated control scheme, to achieve an unprecedented level of performance and simplicity. Its design implements the new International Docking System Standard, which will be a key enabler of diverse and flexible exploration missions in future iterations. NDS qualification was completed in 2017 to support a planned first flight in 2018 on the Boeing CST-100 Starliner.
An integrated thermal model of the Structure Heat Intercept, Insulation, and Vibration Evaluation Rig (SHIIVER) was built in Thermal Desktop/Sinda-Fluint and used to predict thermal performance of the system with the technologies being tested from the eCryo project. The technologies tested include the vapor cooling of the structure using the boil-off and multilayer insulation (MLI) on the hydrogen tank. The thermal model is used to 1) understand system transient behaviors; 2) perform trade studies of different vapor-cooling designs; 3) define the test operations matrix, test procedures, and conditions; and 4) address thermal-related questions at different phases of SHIIVER project. Details of the thermal model are herein described, and results corresponding to the test operations are presented for four different configurations at different fill levels. The thermal performance of each configuration is predicted with the thermal model regarding the heat reduction and boil-off reduction by using vapor cooling and/or MLI on the structures attached to the hydrogen tank.
Fiber Optic Strain Measurements: a) Successfully attached silica fiber optic sensors to both metallics and composites; b) Accomplished valid EFPI strain measurements to 1850 F; c) Successfully attached EFPI sensors to large scale hot-structures; and d) Attached and thermally validated FBG bond and epsilon(sub app). Future Development a) Improve characterization of sensors on C-C and C-SiC substrates; b) Apply application to other composites such as SiC-SiC; c) Assist development of interferometer based Sapphire sensor currently being conducted under a Phase II SBIR; and d) Complete combined thermal/mechanical testing of FBG on composite substrates in controlled laboratory environment.
Oriented attachment (OA) of nanoparticles is an important and common crystallization pathway. Efforts to isolate and quantify the interparticle forces that underlie OA have only recently been initiated and the role played by the interfacial solvent structure in the confined region between nanocrystal surfaces in enabling and controlling OA remains an open question. In this work, classical molecular dynamics (MD) simulations were performed to compute the free energy landscape of nanoparticles approaching mineral substrates in both vacuum and pure water. Periclase (MgO), muscovite (KAl 3 Si 3 O 10 (OH) 2 ), and pyrophyllite (Al 2 Si 4 O 10 (OH) 2 ) nanoparticles exposing the (100), (001), and (001) surfaces, respectively, were brought to contact with substrates exposing the same surfaces. These mineral surfaces were selected to investigate a range of hydrophilicities (periclase > muscovite > pyrophyllite). The MD simulations revealed that the interfacial water structure created free energy minima for nanoparticle approach that corresponded to integer numbers of intervening water layers. The depth of the free energy minima and the free energy barriers between minima correlated positively with the hydrophilicity of the mineral surface. While the range of solvent-mediated interactions was predicted to be short (a few intervening water layers at most for MgO), the free energy barriers were calculated to be large (fraction of an eV per nm 2 ) and the effect of misalignment diminished rapidly with the number of intervening water layers. The results help explain the formation of mesocrystals and the observed ability of nanoparticles to rotate before attachment while in a long-lived solvent-separated state.
An External Tank (ET) attach ring is used in the Space Shuttle System to transfer lateral loads between the ET and the Solid Rocket Booster (SRB). Following the Challenger (51-L) accident, the flight performance of the ET attach ring was reviewed, and negative margins of safety and failed bolts in the attach ring were subsequently identified. The analyses described in this report were performed in order to understand the existing ET attach ring structural response to motor case internal pressurization as well as to aid in an ET attach ring redesign effort undertaken by NASA LaRC. The finite element model as well as the results from linear and nonlinear static structural analyses are described.
The James Webb Space Telescope (JWST) is a 6.6m diameter, segmented, deployable telescope for cryogenic IR space astronomy (approx.40K). The JWST Observatory architecture includes the Optical Telescope Element and the Integrated Science Instrument Module (ISIM) element that contains four science instruments (SI) including a Guider. The ISIM optical metering structure is a roughly 2.2x1.7x2.2mY, asymmetric frame that is composed of carbon fiber and resin tubes bonded to invar end fittings and composite gussets and clips. The structure supports the SIs, isolates the SIs from the OTE, and supports thermal and electrical subsystems. The structure is attached to the OTE structure via strut-like kinematic mounts. The ISM structure must meet its requirements at the approx.40K cryogenic operating temperature. The SIs are aligned to the structure s coordinate system under ambient, clean room conditions using laser tracker and theodolite metrology. The ISM structure is thermally cycled for stress relief and in order to measure temperature-induced mechanical, structural changes. These ambient-to-cryogenic changes in the alignment of SI and OTE-related interfaces are an important component in the JWST Observatory alignment plan and must be verified.
The James Webb Space Telescope (JWST) is a 6.6m diameter, segmented, deployable telescope for cryogenic IR space astronomy (approximately 40K). The JWST Observatory architecture includes the Optical Telescope Element and the Integrated Science Instrument Module (ISIM) element that contains four science instruments (SI) including a Guider. The ISM optical metering structure is a roughly 2.2x1.7x2.2m, asymmetric frame that is composed of carbon fiber and resin tubes bonded to invar end fittings and composite gussets and clips. The structure supports the SIs, isolates the SIs from the OTE, and supports thermal and electrical subsystems. The structure is attached to the OTE structure via strut-like kinematic mounts. The ISIM structure must meet its requirements at the approximately 40K cryogenic operating temperature. The SIs are aligned to the structure's coordinate system under ambient, clean room conditions using laser tracker and theodolite metrology. The ISIM structure is thermally cycled for stress relief and in order to measure temperature-induced mechanical, structural changes. These ambient-to-cryogenic changes in the alignment of SI and OTE-related interfaces are an important component in the JWST Observatory alignment plan and must be verified. We report on the planning for and preliminary testing of a cryogenic metrology system for ISIM based on photogrammetry. Photogrammetry is the measurement of the location of custom targets via triangulation using images obtained at a suite of digital camera locations and orientations. We describe metrology system requirements, plans, and ambient photogrammetric measurements of a mock-up of the ISIM structure to design targeting and obtain resolution estimates. We compare these measurements with those taken from a well known ambient metrology system, namely, the Leica laser tracker system. We also describe the data reduction algorithm planned to interpret cryogenic data from the Flight structure. Photogrammetry was selected from an informal trade study of cryogenic metrology systems because its resolution meets sub-allocations to ISIM alignment requirements and it is a non-contact method that can in principle measure six degrees of freedom changes in target location. In addition, photogrammetry targets can be readily related to targets used for ambient surveys of the structure. By thermally isolating the photogrammetry camera during testing, metrology can be performed in situ during thermal cycling. Photogrammetry also has a small but significant cryogenic heritage in astronomical instrumentation metrology. It was used to validate the displacement/deformation predictions of the reflectors and the feed horns during thermal/vacuum testing (90K) for the Microwave Anisotropy Probe (MAP). It also was used during thermal vacuum testing (100K) to verify shape and component alignment at operational temperature of the High Gain Antenna for New Horizons. With tighter alignment requirements and lower operating temperatures than the aforementioned observatories, ISIM presents new challenges in the development of this metrology system.
Direct numerical simulations with C3-chemistry and radiative heat-loss models have been performed to reveal the internal structure of propagating and attached flames in an axisymmetric fuel jet of methane, ethane, ethylene, acetylene, or propane in air under normal and zero gravity. Observations of the flames were also made at the NASA Glenn 2.2-Second Drop Tower. In computations, the fuel issued into quasi-quiescent air for a fixed mixing time before it was ignited along the centerline at stoichiometry. The edge of the flame propagated through a flammable layer at the laminar flame speed of the stoichiometric fuel-air mixture independent of gravity. For all cases, a peak reactivity spot, i.e., reaction kernel, was formed in the flame base, thereby holding a trailing diffusion flame. The location of the reaction kernel in the attached flames depended inversely on the reactivity. The reaction-kernel correlations between the reactivity and the velocity were developed further using variables related to local Damkahler and Peclet numbers.
Current options for Lunar habitat architecture include inflatable habitats and airlocks. Inflatable structures can have mass and volume advantages over conventional structures. Inflatable structures are perceived to carry additional risk because they are at a lower Technical Readiness Level (TRL) than conventional metallic structures. One of the risks associated with inflatable structures is understanding the tolerance to component damage and the resulting behavior of the system after the damage is introduced. The Damage Tolerance Test (DTT) is designed to study the structural integrity of an expandable structure during and subsequent to induced damage. The TransHab Project developed an experimental inflatable module developed at Johnson Space Center in the 1990's. The TransHab design was originally envisioned for use in Mars Transits but was also studied as a potential habitat for the International Space Station (ISS). The design of the TransHab module was based on a woven design using an Aramid fabric. Testing of this design demonstrated a high level of predictability and repeatability and good correlation with analytical predictions of stresses and deflections. Based on JSC's experience with the design and analysis of woven inflatable structures, the Damage Tolerance Test article was designed and fabricated using a woven design. The Damage Tolerance Test Article consists of a load bearing restraint layer, a bladder or gas barrier, and a structural metallic core. The test article restraint layer is fabricated from one inch wide Kevlar webbing that is woven in a basket weave pattern. Underneath the structural restraint layer is the bladder or gas barrier. For this test the bladder was required to maintain pressure for testing only and was not representative of a flight design. The bladder and structural restraint layer attach to the structural core of the module at steel bulkheads at each end. The two bulkheads are separated by a 10 foot center tube which provides the structural support for the module when in a non-inflated state as well as resists a portion of the axial load when pressurized. The longitudinal members of the structural restraint layer are attached to the bulkheads using a series of clevises that are bolted to the bulkheads. Strain gages are placed on the clevises that can measure change in load when the structural restraint is inflated. The test module is 88 inches in diameter and 120 inches in height. The objectives of the DTT are to (1) verify the structural integrity of the assembled and pressurized structure when a section of the structural restraint layer is cut by a foreign object, and (2) verify the load distribution of the structural restraint layer during pressurization, before and after the structural restraint layer is severed. For this test, a longitudinal structural restraint strap will be severed using a linear shape charge. The linear shape charge was designed specifically for this application to cut only a single longitudinal strap, while not damaging the bladder. An array of strain gages were located at the bulkhead mounted clevises where the longitudinal restraint layer straps are attached. The DTT article was inflated to 45 psig, 25% of the ultimate design pressure, and one of the one-inch wide longitudinal structural members was severed. Strain gage measurements of loading in an array of longitudinal straps were taken throughout pressurization of the module to 45 psig, before firing of the linear shape charge, and after firing of the shape charge and separation of the strap. During testing not only were the original objectives met but better than expected results occurred. This paper will discuss space inflatable structures, damage tolerance analysis, test results, and applicability to the Lunar architecture.
Three methods of attaching surface insulation tiles to shuttle primary structure have been proposed: direct bond, mechanical attachment, and subpanels with standoffs. The direct bond approach is lightweight but is difficult to refurbish and inspect. The subpanel approach is heavier but allows for easy refurbishment since subpanels are easily removed and replaced. The mechanical attachment approach allows easy refurbishment and inspection and is lightweight when an efficient insulator is used between surface insulation tiles and primary structure.
The invention is directed to a microfabricated device. The device includes a substrate that is etched to define mechanical structures at least some of which are anchored laterally to the remainder of the substrate. Electrical isolation at points where mechanical structures are attached to the substrate is provided by filled isolation trenches. Filled trenches may also be used to electrically isolate structure elements from each other at points where mechanical attachment of structure elements is desired. The performance of microelectromechanical devices is improved by 1) having a high-aspect-ratio between vertical and lateral dimensions of the mechanical elements, 2) integrating electronics on the same substrate as the mechanical elements, 3) good electrical isolation among mechanical elements and circuits except where electrical interconnection is desired.
Information about vibrations of structure and/or of nonvibration sensor attached to structure extracted from output of sensor. Sensor operated in usual way except, output fed to power-spectral-density analyzer. Vibrational components easily distinguishable in analyzer output because they have frequencies much higher than those of more-slowly-varying temperature, pressure, or other normally desired components. Spectral-analysis technique applied successfully to high-frequency resistance changes in output of platinum-wire resistance thermometer: vibrational peaks in resistance frequency spectrum confirmed by spectrum from accelerometer. Technique also showed predicted 17-kHz vibrational resonance in strain-guage-supporting beam in pressure sensor.
The analytical and experimental results of an all-composite wing stub box are presented in this report. The wing stub box, which is representative of an inboard portion of a commercial transport high-aspect-ratio wing, was fabricated from stitched graphite-epoxy material with a Resin Film Infusion manufacturing process. The wing stub box was designed and constructed by the McDonnell Douglas Aerospace Company as part of the NASA Advanced Composites Technology program. The test article contained metallic load-introduction structures on the inboard and outboard ends of the graphite-epoxy wing stub box. The root end of the inboard load introduction structure was attached to a vertical reaction structure, and an upward load was applied to the outermost tip of the outboard load introduction structure to induce bending of the wing stub box. A finite element model was created in which the center portion of the wing-stub-box upper cover panel was modeled with a refined mesh. The refined mesh was required to represent properly the geometrically nonlinear structural behavior of the upper cover panel and to predict accurately the strains in the stringer webs of the stiffened upper cover panel. The analytical and experimental results for deflections and strains are in good agreement.
The manned core platform of the Space Station will repeatedly provide observations of the tropical and subtropical parts of earth from about 500 km altitude in a 28.5 degree orbit. This orbit covers most of the landmasses in the Southern Hemisphere, half of the world ocean, and more than 120 nations. The platform can provide remote-sensing activities by means of: (1) sensors attached to structural members of the platform; (2) sensors deployed on one or more pointable platforms attached to structural members, or man-tended in separate orbits; (3) sensors operated from the resource nodes; (4) sensors operated through earth-viewing, optical-quality windows; and (5) hand-held cameras operated by crew members throughout the facility. Major applications of observations from the core platform in the exploration for and production of resources are discussed.
Structural and aerodynamic developments in attached inflatable decelerators for deployment of payload at supersonic speeds