ALBERT

Description: To reduce weight and increase the mobility, comfort, and performance of future spacesuits, flexible, thermally conductive fabrics and plastic tubes are needed for the Liquid Cooling and Ventilation Garment. Such improvements would allow astronauts to operate more efficiently and safely for extended extravehicular activities. As an approach to raise the thermal conductivity (TC) of an ethylene vinyl acetate copolymer (Elvax 260), it was compounded with three types of carbon based nanofillers: multi-walled carbon nanotubes (MWCNTs), vapor grown carbon nanofibers (CNFs), and expanded graphite (EG). In addition, other nanofillers including metallized CNFs, nickel nanostrands, boron nitride, and powdered aluminum were also compounded with Elvax 260 in the melt at various loading levels. In an attempt to improve compatibility between Elvax 260 and the nanofillers, MWCNTs and EG were modified by surface coating and through noncovalent and covalent attachment of organic molecules containing alkyl groups. Ribbons of the nanocomposites were extruded to form samples in which the nanofillers were aligned in the direction of flow. Samples were also fabricated by compression molding to yield nanocomposites in which the nanofillers were randomly oriented. Mechanical properties of the aligned samples were determined by tensile testing while the degree of dispersion and alignment of nanoparticles were investigated using high-resolution scanning electron microscopy. TC measurements were performed using a laser flash (Nanoflash ) technique. TC of the samples was measured in the direction of, and perpendicular to, the alignment direction. Additionally, tubing was also extruded from select nanocomposite compositions and the TC and mechanical flexibility measured.

Description: Most ongoing US activities related to space nuclear power and propulsion are sponsored by NASA. NASA-spons0red space nuclear work is currently focused on evaluating potential fission surface power (FSP) systems and on radioisotope power systems (RPS). In addition, significant efforts related to nuclear thermal propulsion (NTP) systems have been completed and will provide a starting point for potential future NTP work.

Description: This paper describes the Mars transportation vehicle design concepts developed by the Marshall Space Flight Center (MSFC) Advanced Concepts Office. These vehicle design concepts provide an indication of the most demanding and least demanding potential requirements for nuclear thermal propulsion systems for human Mars exploration missions from years 2025 to 2035. Vehicle concept options vary from large "all-up" vehicle configurations that would transport all of the elements for a Mars mission on one vehicle. to "split" mission vehicle configurations that would consist of separate smaller vehicles that would transport cargo elements and human crew elements to Mars separately. Parametric trades and sensitivity studies show NTP stage and engine design options that provide the best balanced set of metrics based on safety, reliability, performance, cost and mission objectives. Trade studies include the sensitivity of vehicle performance to nuclear engine characteristics such as thrust, specific impulse and nuclear reactor type. Tbe associated system requirements are aligned with the NASA Exploration Systems Mission Directorate (ESMD) Reference Mars mission as described in the Explorations Systems Architecture Study (ESAS) report. The focused trade studies include a detailed analysis of nuclear engine radiation shield requirements for human missions and analysis of nuclear thermal engine design options for the ESAS reference mission.

Description: With the SMART-1, Department of Defense, and commercial industry successes in Hall thruster technologies, NASA has started considering Hall thrusters for science missions. The recent Discovery proposals included a Hall thruster science mission and the In-Space Propulsion Project is investing in Hall thruster technologies. As the confidence in Hall thrusters improve, ambitious multi-thruster missions are being considered. Science missions often require large throttling ranges due to the 1/r(sup 2) power drop-off from the sun. Deep throttling of Hall thrusters will impact the overall system performance. Also, Hall thrusters can be throttled with both current and voltage, impacting erosion rates and performance. Last, electric propulsion thruster lifetime qualification has previously been conducted with long duration full power tests. Full power tests may not be appropriate for NASA science missions, and a combination of lifetime testing at various power levels with sufficient analysis is recommended. Analyses of various science missions and throttling schemes using the Aerojet BPT-4000 and NASA 103M HiVHAC thruster are presented.

Description: Rocket propulsion determines the primary characteristics of any space vehicle; how fast and far it can go, its lifetime, and its capabilities. It is the primary factor in safety and reliability and the biggest cost driver. The extremes of heat and pressure produced by propulsion systems push the limits of materials used for manufacturing. Space travel is very unforgiving with little room for errors, and so many things can go wrong with these very complex systems. So we have to plan for failure and that makes it costly. But what is more exciting than the roar of a rocket blasting into space? By its nature the propulsion world is conservative. The stakes are so high at every launch, in terms of payload value or in human life, that to introduce new components to a working, qualified system is extremely difficult and costly. Every launch counts and no risks are tolerated, which leads to the space world's version of Catch-22:"You can't fly till you flown." The last big 'game changer' in propulsion was the use of liquid hydrogen as a fuel. No new breakthrough, low cost access to space system will be developed without new efficient propulsion systems. Because there is no large commercial market driving investment in propulsion, what propulsion research is done is sponsored by government funding agencies. A further difficulty in propulsion technology development is that there are so few new systems flying. There is little opportunity to evolve propulsion technologies and to update existing systems with results coming out of research as there is in, for example, the auto industry. The biggest hurdle to space exploration is getting off the ground. The launch phase will consume most of the energy required for any foreseeable space exploration mission. The fundamental physical energy requirements of escaping earth's gravity make it difficult. It takes 60,000 kJ to put a kilogram into an escape orbit. The vast majority (-97%) of the energy produced by a launch vehicle is used to get propellants off the ground to be burned later. A modem launch vehicle is usually able to put no more than 1.5%-3% of its total liftoff weight into low earth orbit.

Description: Nuclear and radioisotope powered electric thrusters are being developed as primary in-space propulsion systems for potential future robotic and piloted space missions. Possible applications for high power nuclear electric propulsion include orbit raising and maneuvering of large space platforms, lunar and Mars cargo transport, asteroid rendezvous and sample return, and robotic and piloted planetary missions, while lower power radioisotope electric propulsion could significantly enhance or enable some future robotic deep space science missions. This paper provides an overview of recent U.S. high power electric thruster research programs, describing the operating principles, challenges, and status of each technology. Mission analysis is presented that compares the benefits and performance of each thruster type for high priority NASA missions. The status of space nuclear power systems for high power electric propulsion is presented. The paper concludes with a discussion of power and thruster development strategies for future radioisotope electric propulsion systems,

Description: The oxidation of SiC-protected carbon/carbon through machined slots and naturally occurring craze cracks in the SiC was studied. The slot and crack geometries were characterized, and the subsurface oxidation of the carbon/carbon substrate at temperatures of 1000 to 1300 C in air was assessed using weight change, x-ray computed tomography, and optical microscopy of sections. Rate constants were derived from these measurements and compared with a two-step diffusion control model of carbon oxidation. Oxidation kinetic measurements on both the specimens with machined slots and with naturally occurring craze cracks showed good agreement with the model.

Description: An analysis of the heat transfer in a tool for producing neat resin disks was conducted to determine how to bring about a better agreement between the tool temperature and the applied temperature profile. Using the commercial code FLUENT to investigate the relative effects of heat conduction into the tool and heat loss from the tool by convection, it was shown that convective heat transfer appears more important than conduction in controlling the tool performance. Decreasing the height of the tool was predicted to decrease the heat losses by convection. Redesign of the tool based on this analysis resulted in the tool experiencing the applied temperature profile.

Description: This paper will present an overview of efforts to investigate the application of Raman spectroscopy for the characterization of Kevlar materials. Raman spectroscopy is a laser technique that is sensitive to molecular interactions in materials such as Kevlar, graphite and carbon used in composite materials. The overall goal of this research reported here is to evaluate Raman spectroscopy as a potential nondestructive evaluation (NDE) tool for the detection of stress rupture in Kevlar composite over-wrapped pressure vessels (COPVs). Characterization of the Raman spectra of Kevlar yarn and strands will be presented and compared with analytical models provided in the literature. Results of testing to investigate the effects of creep and high-temperature aging on the Raman spectra will be presented.

Description: In 2005-06, the Prometheus program funded a number of tasks at the NASA-Marshall Space Flight Center (MSFC) to support development of a Nuclear Thermal Propulsion (NTP) system for future manned exploration missions. These tasks include the following: 1. NTP Design Develop Test & Evaluate (DDT&E) Planning 2. NTP Mission & Systems Analysis / Stage Concepts & Engine Requirements 3. NTP Engine System Trade Space Analysis and Studies 4. NTP Engine Ground Test Facility Assessment 5. Non-Nuclear Environmental Simulator (NTREES) 6. Non-Nuclear Materials Fabrication & Evaluation 7. Multi-Physics TCA Modeling. This presentation is a overview of these tasks and their accomplishments

Description: Multi-layer insulation, MLI, is a material used on rocket bodies and satellites mainly for thermal insulation. MLI can be comprised of a variety of materials, layer numbers, and dimensions based on its purpose. A common composition of MLI consists of outer facing copper-colored Kapton with an aluminized backing for the top and bottom layers and the middle consisting of alternating layers of DARCON or Nomex netting with aluminized Mylar. If this material became separated from the spacecraft or rocket body its orbit would vary greatly in eccentricity due to its high area to mass (A/m) and susceptibility to solar radiation pressure perturbations. Recently a debris population was found with high A/m, which could be MLI. Laboratory photometric measurements of one intact piece and three different layers of MLI is presented in an effort to predict the characteristics of a MLI light curve and aid in identifying the source of the new population. For this paper, the layers used will be consistent with the common MLI mentioned in the above paragraph. Using a robotic arm, the piece was rotated from 0-360 degrees in one degree increments along the object s longest axis. Laboratory photometric data was recorded with a CCD camera using various filters (Johnson B, Johnson V and Bessell R). The measurements were taken at an 18 degree (light-object-camera) phase angle. As expected, the MLI pieces showed characteristics similar to a bimodal magnitude plot of a flat plate, but with more photometric features, dependant upon the layer of MLI. Time exposures varied from piece to piece such that the amount of pixels saturated would be minimal. In addition to photometric laboratory measurements, laboratory spectral measurements are shown for the same MLI samples. Spectral data will be combined to match the wavelength region of photometric data so a measure of truth can be established for the photometric measurements. Spectral data shows a strong absorption feature near 4800 angstroms, which is due to the copper color of Kapton. If the debris is MLI and the outer layer of copper coloring of Kapton is present, evidence would be seen spectrally by the specific absorption feature as well as using R-B (red-blue) light curves. Using laboratory photometric measurements and the results from spectral laboratory measurements, an optical property database is provided for an object with a high A/m. The benefits of this database for remote optical measurements of orbital debris are shown by illustrating the optical properties expected for a high A/m object, specifically common satellite and rocket body MLI.

Description: New requirements for thermal insulation include robust Multilayer insulation (MU) systems that work for a range of environments from high vacuum to no vacuum. Improved MLI systems must be simple to install and maintain while meeting the life-cycle cost and thermal performance objectives. Performance of actual MLI systems has been previously shown to be much worse than ideal MLI. Spacecraft that must contain cryogens for both lunar service (high vacuum) and ground launch operations (no vacuum) are planned. Future cryogenic spacecraft for the soft vacuum environment of Mars are also envisioned. Industry products using robust MLI can benefit from improved cost-efficiency and system safety. Novel materials have been developed to operate as excellent thermal insulators at vacuum levels that are much less stringent than the absolute high vacuum requirement of current MLI systems. One such robust system, Layered Composite Insulation (LCI), has been developed by the Cryogenics Test Laboratory at NASA Kennedy Space Center. The experimental testing and development of LCI is the focus of this paper. LCI thermal performance under cryogenic conditions is shown to be six times better than MLI at soft vacuum and similar to MLI at high vacuum. The experimental apparent thermal conductivity (k-value) and heat flux data for LCI systems are compared with other MLI systems.

Description: The research testing and demonstration of new bulk-fill materials for cryogenic thermal insulation systems was performed by the Cryogenics Test Laboratory at NASA Kennedy Space Center. Thermal conductivity testing under actual-use cryogenic conditions is a key to understanding the total system performance encompassing engineering, economics, and materials factors. A number of bulk fill insulation materials, including aerogel beads, glass bubbles, and perlite powder, were tested using a new cylindrical cryostat. Boundary temperatures for the liquid nitrogen boil-off method were 293 K and 78 K. Tests were performed as a function of cold vacuum pressure from high vacuum to no vacuum conditions. Results are compared with other complementary test methods in the range of 300 K to 20 K. Various testing techniques are shown to be required to obtain a complete understanding of the operating performance of a material and to provide data for answers to design engineering questions.

Description: One of the toughest challenges facing Solid Rocket Booster (SRB) engineers is to ensure that any design changes made to the Shuttle-Derived Booster Separation Motors (BSM) for future space exploration vehicles is able to withstand the increasingly hostile motor firing environment without cracking its critical component - the graphite throat. This paper presents a critical analysis methodology and techniques for assessing effects of BSM design changes with great accuracy and precision. For current Space Shuttle operation, the motor firing occurs at SRB separation - approximately 125 seconds after Shuttle launch at an altitude of about 28 miles. The motor operation event lasts about two seconds, however, the surface temperature of the graphite throat increases approximately 3400 F in less than one second with a corresponding increase in surface pressure of approximately 2200 pounds per square inch (psi) in less than one-tenth of a second. To capture this process fully and accurately, a two-phase sequentially coupled thermal-mechanical finite element approach was developed. This method allows the time- and location-dependent pressure fields to interact with the spatial-temporal thermal fields throughout the operation. The material properties of graphite throat are orthotropic and temperature-dependent. The analysis involves preload and multiple body contacts.

Description: No abstract available

Description: The NASA New Millennium Program Space Technology 7 (ST7) project will validate technology for precision spacecraft control. The ST7 disturbance reduction system (DRS) will contain new micropropulsion technology to be flown as part of the European Space Agency's LISA (laser interferometer space antenna) Pathfinder project. After launch into a low Earth orbit in early 2010, the LISA Pathfinder spacecraft will be maneuvered to a halo orbit about the Earth-Sun LI Lagrange point for operations. The DRS will control the position of the spacecraft relative to a reference to an accuracy of one nanometer over time scales of several thousand seconds. To perform the control the spacecraft will use a new colloid thruster technology. The thrusters will operate over the range of 5 to 30 micro-Newtons with precision of 0.1 micro-Newton. The thrust will be generated by using a high electric field to extract charged droplets of a conducting colloid fluid and accelerating them with a precisely adjustable voltage. The control position reference will be provided by the European LISA Technology Package, which will include two nearly free-floating test masses. The test mass position and attitude will be sensed and adjusted using electrostatic capacitance bridges. The DRS will control the spacecraft position with respect to one test mass while minimizing disturbances on the second test mass. The dynamic control system will cover eighteen degrees of freedom, six for each of the test masses and six for the spacecraft. In the absence of other disturbances, the test masses will slowly gravitate toward local concentrations of spacecraft mass. The test mass acceleration must be minimized to maintain the acceleration of the enclosing drag-free spacecraft within the control authority of the micropropulsion system. Therefore, test mass acceleration must be predicted by accurate measurement of mass distribution, then offset by the placement of specially shaped balance masses near each test mass. The - acceleration is characterized by calculating the gravitational effect of over ten million modeled points of a nearly 500-kg spacecraft. This paper provides an overview of the mission technology and the process of precision mass modeling of the DRS equipment.

Description: With the introduction of the new manned space effort through the Constellation Program, there is an interest to have a basic comparison of the current Space Shuttle Main Engine (SSME) to the J-2X engine used for the second stage of both the Ares I and Ares V rockets. This paper seeks to compare size, weight and thrust capabilities while drawing simple conclusions on differences between the two engines.

Description: The detection of corrosion beneath Space Shuttle Orbiter thermal protective system is traditionally accomplished by removing the Reusable Surface Insulation tiles and performing a visual inspection of the aluminum substrate and corrosion protection system. This process is time consuming and has the potential to damage high cost tiles. To evaluate non-intrusive NDE methods, a Proof of Concept (PoC) experiment was designed and test panels were manufactured. The objective of the test plan was three-fold: establish the ability to detect corrosion hidden from view by tiles; determine the key factor affecting detectability; roughly quantify the detection threshold. The plan consisted of artificially inducing dimensionally controlled corrosion spots in two panels and rebonding tile over the spots to model the thermal protective system of the orbiter. The corrosion spot diameter ranged from 0.100" to 0.600" inches and the depth ranged from 0.003" to 0.020". One panel consisted of a complete factorial array of corrosion spots with and without tile coverage. The second panel consisted of randomized factorial points replicated and hidden by tile. Conventional methods such as ultrasonics, infrared, eddy current and microwave methods have shortcomings. Ultrasonics and IR cannot sufficiently penetrate the tiles, while eddy current and microwaves have inadequate resolution. As such, the panels were interrogated using Backscatter Radiography and Terahertz Imaging. The terahertz system successfully detected artificially induced corrosion spots under orbiter tile and functional testing is in-work in preparation for implementation.

Description: For near-Earth application, solar electric propulsion advocates have focused on Low Earth Orbit (LEO) to Geosynchronous (GEO) low-thrust transfers because of the significant improvement in capability over chemical alternatives. While the performance gain attained from starting with a lower orbit is large, there are also increased transfer times and radiation exposure risk that has hindered the commercial advocacy for electric propulsion stages. An incremental step towards electric propulsion stages is the use of integrated solar electric propulsion systems (SEPS) for GTO to GEO transfer. Thorough analyses of electric propulsion systems options and performance are presented. Results are based on existing or near-term capabilities of Arcjets, Hall thrusters, and Gridded Ion engines. Parametric analyses based on "rubber" thruster and launch site metrics are also provided.

Description: The objectives of currently planned exploration efforts, as well as those further in the future, require significant advancements in propulsion technologies. The current Lunar exploration architecture has set goals and mission objectives that necessitate the use of new systems and the extension of existing technologies beyond present applications. In the near term, the majority of these technologies are the result of a need to apply high performing cryogenic propulsion systems to long duration in-space applications. Advancement of cryogenic propulsion to these applications is crucial to provide higher performing propulsion systems that reduce the vehicle masses; enhance the safety of vehicle systems and ground operations; and provide a path for In-situ Resource Utilization (ISRU).Use of a LOX/LH2 main propulsion system for Lunar Lander Descent is a top priority because more conventional storable propellants are far from meeting the performance needs of the current architecture. While LOX/LH2 pump feed engines have been used in flight applications for many years, these engines have limited throttle capabilities. Engines that are capable of much greater throttling while still meeting high performance goals are a necessity to achieving exploration goals. Applications of LOX/CH4 propulsion to Lander ascent propulsion systems and reaction control systems are also if interest because of desirable performance and operations improvements over conventional storable systems while being more suitable for use of in-situ produced propellants. Within the current lunar architecture, use of cryogenic propulsion for the Earth Departure Stage and Lunar Lander elements also necessitate the need for advanced Cryogenic Fluid Management technologies. These technologies include long duration propellant storage/distribution, low-gravity propellant management, cryogenic couplings and disconnects, light weight composite tanks and support structure, and subsystem integration. In addition to the propulsive and fluid management system technologies described, many component level technologies are also required to enable to the success if the integrated systems. The components include, but are not limited to, variable/throttling valves, variable position actuators, leak detectors, light weight cryogenic fluid pumps, sensor technology and others. NASA, partnering with the Aerospace Industry must endeavor to develop these, and other promising propulsion technologies, to enable the implements of the country's goals in exploration of the Moon, Mars and beyond.

Description: NASA's In-Space Propulsion Technology Program has developed the first generation of solar sail propulsion systems sufficient to accomplish inner solar system science and exploration missions. These first generation solar sails, when operational, will range in size from 40 meters to well over 100 meters in diameter and have an areal density of less than 13 grams per square meter. A rigorous, multi-year technology development effort culminated in 2005 with the testing of two different 20-m solar sail systems under thermal vacuum conditions. The first system, developed by ATK Space Systems of Goleta, California, uses rigid booms to deploy and stabilize the sail. In the second approach, L'Garde, Inc. of Tustin, California uses inflatable booms that rigidize in the coldness of space to accomplish sail deployment. This effort provided a number of significant insights into the optimal design and expected performance of solar sails as well as an understanding of the methods and costs of building and using them. In a separate effort, solar sail orbital analysis tools for mission design were developed and tested. Laboratory simulations of the effects of long-term space radiation exposure were also conducted on two candidate solar sail materials. Detailed radiation and charging environments were defined for mission trajectories outside the protection of the earth's magnetosphere, in the solar wind environment. These were used in other analytical tools to prove the adequacy of sail design features for accommodating the harsh space environment. Preceding and in conjunction with these technology efforts, NASA sponsored several mission application studies for solar sails. Potential missions include those that would be flown in the near term to study the sun and be used in space weather prediction to one that would use an evolved sail capability to support humanity's first mission into nearby interstellar space. This paper will describe the status of solar sail propulsion within NASA, nearterm solar sail mission applications, and near-term plans for further development.

Description: An approach to assessing the delamination propagation capabilities in commercial finite element codes is presented and demonstrated for one code. For this investigation, the Double Cantilever Beam (DCB) specimen and the Single Leg Bending (SLB) specimen were chosen for full three-dimensional finite element simulations. First, benchmark results were created for both specimens. Second, starting from an initially straight front, the delamination was allowed to propagate. Good agreement between the load-displacement relationship obtained from the propagation analysis results and the benchmark results could be achieved by selecting the appropriate input parameters. Selecting the appropriate input parameters, however, was not straightforward and often required an iterative procedure. Qualitatively, the delamination front computed for the DCB specimen did not take the shape of a curved front as expected. However, the analysis of the SLB specimen yielded a curved front as may be expected from the distribution of the energy release rate and the failure index across the width of the specimen. Overall, the results are encouraging but further assessment on a structural level is required.

Description: Pulsed inductive plasma accelerators are spacecraft propulsion devices in which energy is stored in a capacitor and then discharged through an inductive coil. The device is electrodeless, inducing a current sheet in a plasma located near the face of the coil. The propellant is accelerated and expelled at a high exhaust velocity (order of 10 km/s) through the interaction of the plasma current and the induced magnetic field. The Faraday Accelerator with RF-Assisted Discharge (FARAD) thruster[1,2] is a type of pulsed inductive plasma accelerator in which the plasma is preionized by a mechanism separate from that used to form the current sheet and accelerate the gas. Employing a separate preionization mechanism allows for the formation of an inductive current sheet at much lower discharge energies and voltages than those used in previous pulsed inductive accelerators like the Pulsed Inductive Thruster (PIT). A benchtop FARAD thruster was designed following guidelines and similarity performance parameters presented in Refs. [3,4]. This design is described in detail in Ref. [5]. In this paper, we present the temporally and spatially resolved measurements of the preionized plasma and inductively-accelerated current sheet in the FARAD thruster operating with a Vector Inversion Generator (VIG) to preionize the gas and a Bernardes and Merryman circuit topology to provide inductive acceleration. The acceleration stage operates on the order of 100 J/pulse. Fast-framing photography will be used to produce a time-resolved, global view of the evolving current sheet. Local diagnostics used include a fast ionization gauge capable of mapping the gas distribution prior to plasma initiation; direct measurement of the induced magnetic field using B-dot probes, induced azimuthal current measurement using a mini-Rogowski coil, and direct probing of the number density and electron temperature using triple probes.

Description: High voltage solar arrays are necessary for direct-drive solar electric propulsion, which has many advantages, including simplicity and high efficiency. Even when direct-drive is not used, the use of high voltage solar arrays leads to power transmission and conversion efficiencies in electric propulsion Power Management and Distribution. Nevertheless, high voltage solar arrays may lead to temporary power disruptions, through the so-called primary electrostatic discharges, and may permanently damage arrays, through the so-called permanent sustained discharges between array strings. Design guidance is needed to prevent these solar array discharges, and to prevent high power drains through coupling between the electric propulsion devices and the high voltage solar arrays. While most electric propulsion systems may operate outside of Low Earth Orbit, the plasmas produced by their thrusters may interact with the high voltage solar arrays in many ways similarly to Low Earth Orbit plasmas. A brief description of previous experiences with high voltage electric propulsion systems will be given in this paper. There are two new official NASA documents available free through the NASA Standards website to help in designing and testing high voltage solar arrays for electric propulsion. They are NASA-STD-4005, the Low Earth Orbit Spacecraft Charging Design Standard, and NASA-HDBK-4006, the Low Earth Orbit Spacecraft Charging Design Handbook. Taken together, they can both educate the high voltage array designer in the engineering and science of spacecraft charging in the presence of dense plasmas and provide techniques for designing and testing high voltage solar arrays to prevent electrical discharges and power drains.

Description: The J-2X program calls for the upgrade of the Apollo-era Rocketdyne J-2 engine to higher power levels, using new materials and manufacturing techniques, and with more restrictive safety and reliability requirements than prior human-rated engines in NASA history. Such requirements demand a comprehensive systems engineering effort to ensure success. Pratt & Whitney Rocketdyne system engineers performed a functional analysis of the engine to establish the functional architecture. J-2X functions were captured in six major operational blocks. Each block was divided into sub-blocks or states. In each sub-block, functions necessary to perform each state were determined. A functional engine schematic consistent with the fidelity of the system model was defined for this analysis. The blocks, sub-blocks, and functions were sequentially numbered to differentiate the states in which the function were performed and to indicate the sequence of events. The Engine System was functionally partitioned, to provide separate and unique functional operators. Establishing unique functional operators as work output of the System Architecture process is novel in Liquid Propulsion Engine design. Each functional operator was described such that its unique functionality was identified. The decomposed functions were then allocated to the functional operators both of which were the inputs to the subsystem or component performance specifications. PWR also used a novel approach to identify and map the engine functional requirements to customer-specified functions. The final result was a comprehensive Functional Flow Block Diagram (FFBD) for the J-2X Engine System, decomposed to the component level and mapped to all functional requirements. This FFBD greatly facilitates component specification development, providing a well-defined trade space for functional trades at the subsystem and component level. It also provides a framework for function-based failure modes and effects analysis (FMEA), and a rigorous baseline for the functional architecture.

Description: During the Extended Life Test of the DS1 flight spare ion thruster, the engine was subjected to sensitvity testing in order to characterize the macroscopic dependence of discharge chamber sensitivity to a +\-3% vatiation in main flow, cathode flow and beam current, and to +\5% variation in beam and accelerator voltage, was determined for the minimum- (THO), half- (TH8) and full power (TH15) throttle levels. For each power level investigared, 16 high/low operating conditions were chosen to vary the flows, beam current, and grid voltages in in a matrix that mapped out the entire parameter space. The matrix of data generated was used to determine the partial derivative or senitivity of the dependent parameters--discharge voltage, discharge current, discharge loss, double-to-single-ion current ratio, and neutralizer-keeper voltage--to the variation in the independent parameters--main flow, cathode flow, beam current, and beam voltage. The sensititivities of each dependent parameter with respect to each independent parameter were determined using a least-square fit routine. Variation in these sensitivities with thruster runtime was recorded over the duration of the ELT, to detemine if discharge performance changed with thruster wear. Several key findings have been ascertained from the sensitivity testing. Discharge operation is most sensitve to changes in cathode flow and to a lesser degree main flow. The data also confirms that for the NSTAR configuration plasma production is limited by primary electron input due to the fixed neutral population. Key sensitivities along with their change with thruster wear (operating time) will be presented. In addition double ion content measurements with an ExB probe will also be presented to illustrate beam ion production and content sensitivity to the discharge chamber operating parameteres.

Description: The NASA's Evolutionary Xenon Thruster (NEXT) ion propulsion system has been in advanced technology development under the NASA In-Space Propulsion Technology project. The highest fidelity hardware planned has now been completed by the government/industry team, including a flight prototype model (PM) thruster, an engineering model (EM) power processing unit, EM propellant management assemblies, a breadboard gimbal, and control unit simulators. Subsystem and system level technology validation testing is in progress. To achieve the objective Technology Readiness Level 6, environmental testing is being conducted to qualification levels in ground facilities simulating the space environment. Additional tests have been conducted to characterize the performance range and life capability of the NEXT thruster. This paper presents the status and results of technology validation testing accomplished to date, the validated subsystem and system capabilities, and the plans for completion of this phase of NEXT development.

Description: Florida is home to NASA's Launch Operations Center. Since its establishment in July 1962, the spaceport has served as the departure gate for every American manned mission and hundreds of advanced scientific spacecraft under the Launch Services Program. The center was renamed the John F. Kennedy Space Center in late 1963 to honor the president who put America on the path to the moon. Today, NASA is on the edge of a bold new chaIlenge: the ConsteIlation Program. ConsteIlation is a NASA program to create a new generation of spacecraft for human spaceflight, consisting primarily of the Ares I and Ares V launch vehicles, the Orion crew capsule, the Earth Departure stage and the Lunar access module. These spacecraft will be capable of performing a variety of missions, from Space Station resupply to lunar landings. The ambitious new endeavor caIls for NASA to return human explorers to the moon and then venture even farther, to Mars and beyond. As the nation's premier spaceport, Kennedy Space Center (KSC) will playa critical role in this new chapter in exploration, particularly in the conversion of the launch facilities to accommodate the new launch vehicles. To prepare for this endeavor, the launch site and facilities for the next generation of crew and cargo vehicles must be redesigned, assembled and tested. One critical factor that is being carefuIly considered during the renovation is protecting the new facilities and structures from corrosion and deterioration.

Description: TUFROC (Toughened Uni-piece Fibrous Reinforced Oxidation-resistant Composite) has been developed as a new thermal protection system (TPS) material for wing leading edge and nose cap applications. The composite withstands temperatures up to 1,970 K, and consists of a toughened, high-temperature surface cap and a low-thermal-conductivity base, and is applicable to both sharp and blunt leading edge vehicles. This extends the possible application of fibrous insulation to the wing leading edge and/or nose cap on a hypersonic vehicle. The lightweight system comprises a treated carbonaceous cap composed of ROCCI (Refractory Oxidation-resistant Ceramic Carbon Insulation), which provides dimensional stability to the outer mold line, while the fibrous base material provides maximum thermal insulation for the vehicle structure.

Description: A document proposes a CO2-reduction primary electrochemical cell as a building block of batteries to supply electric power on the surface of Venus. The basic principle of the proposed cell is similar to that of terrestrial Zn-air batteries, the major differences being that (1) the anode metal would not be Zn and (2) CO2, which is about 96.5 mole percent of the Venusian atmosphere, would be used, instead of O2, as the source of oxygen.

Description: Spray chemical vapor deposition (spray CVD) processes of a special type have been investigated for use in making CuInS2 absorber layers of thin-film solar photovoltaic cells from either of two subclasses of precursor compounds: [(PBu3) 2Cu(SEt)2In(SEt)2] or [(PPh3)2Cu(SEt)2 In(SEt)2]. The CuInS2 films produced in the experiments have been characterized by x-ray diffraction, scanning electron microscopy, energy-dispersive spectroscopy, and four-point-probe electrical tests.

Description: A material consisting of a barium calcium aluminosilicate glass reinforced with 4 weight percent of boron nitride nanotubes (BNNTs) has shown promise for use as a sealant in planar solid oxide fuel cells (SOFCs).

Description: The engineering design and analysis of air-breathing propulsion systems relies heavily on zero- or one-dimensional properties (e.g. thrust, total pressure recovery, mixing and combustion efficiency, etc.) for figures of merit. The extraction of these parameters from experimental data sets and/or multi-dimensional computational data sets is therefore an important aspect of the design process. A variety of methods exist for extracting performance measures from multi-dimensional data sets. Some of the information contained in the multi-dimensional flow is inevitably lost when any one-dimensionalization technique is applied. Hence, the unique assumptions associated with a given approach may result in one-dimensional properties that are significantly different than those extracted using alternative approaches. The purpose of this effort is to examine some of the more popular methods used for the extraction of performance measures from multi-dimensional data sets, reveal the strengths and weaknesses of each approach, and highlight various numerical issues that result when mapping data from a multi-dimensional space to a space of one dimension.

Description: Ferroelectrics:Polymer composites can be considered an established substitute for conventional electroceramics and ferroelectric polymers. The composites have a unique blend of polymeric properties such as mechanical flexibility, high strength, formability, and low cost, with the high electro-active properties of ceramic materials. They have attracted considerable interest because of their potential use in pyroelectric infrared detecting devices and piezoelectric transducers. These flexible sensors and transducers may eventually be useful for their health monitoring applications for NASA crew launch vehicles and crew exploration vehicles being developed. In the light of many technologically important applications in this field, it is worthwhile to present an overview of the pyroelectric infrared detector theory, models to predict dielectric behavior and pyroelectric coefficient, and the concept of connectivity and fabrication techniques of biphasic composites. An elaborate review of Pyroelectric-Polymer composite materials investigated to date for their potential use in pyroelectric infrared detectors is presented.

Description: This paper examines the use of a continuum damage model to predict strength and size effects in notched carbon-epoxy laminates. The effects of size and the development of a fracture process zone before final failure are identified in an experimental program. The continuum damage model is described and the resulting predictions of size effects are compared with alternative approaches: the point stress and the inherent flaw models, the Linear-Elastic Fracture Mechanics approach, and the strength of materials approach. The results indicate that the continuum damage model is the most accurate technique to predict size effects in composites. Furthermore, the continuum damage model does not require any calibration and it is applicable to general geometries and boundary conditions.

Description: During 2003 and 2004, the Johnson Space Center's White Sands Testing Facility in Las Cruces, New Mexico conducted hypervelocity impact tests on the space shuttle wing leading edge. Hypervelocity impact tests were conducted to determine if Micro-Meteoroid/Orbital Debris impacts could be reliably detected and located using simple passive ultrasonic methods. The objective of Target RCC16R was to study hypervelocity impacts through the reinforced carbon-carbon (RCC) panels of the Wing Leading Edge. Impact damage was detected using lightweight, low power instrumentation capable of being used in flight.

Description: The resin content, and by extension the fiber volume, of carbon fiber/cyanate ester composite laminates are measured using thermogravimetric analysis (TGA). Conventional measurement involves acid digestion of the laminate to determine resin content. The mean difference between techniques is 0.03%. In addition to eliminating the hazards and environmental impact of standard acid digestion, the TGA technique allows quantification of errors associated with fiber volume measurements, e.g. incomplete resin removal or fiber degradation. An additional benefit of the TGA technique is a reduction in sample size requirements, allowing the examination of fiber volume changes in complex shapes.

Description: Interlaminar fracture mechanics has proven useful for characterizing the onset of delaminations in composites and has been used successfully primarily to investigate onset in fracture toughness specimens and laboratory size coupon type specimens. Future acceptance of the methodology by industry and certification authorities, however, requires the successful demonstration of the methodology on the structural level. For this purpose, a panel was selected that is reinforced with stiffeners. Shear loading causes the panel to buckle, and the resulting out-of-plane deformations initiate skin/stiffener separation at the location of an embedded defect. A small section of the stiffener foot, web and noodle as well as the panel skin in the vicinity of the delamination front were modeled with a local 3D solid model. Across the width of the stiffener foot, the mixedmode strain energy release rates were calculated using the virtual crack closure technique. A failure index was calculated by correlating the results with a mixed-mode failure criterion of the graphite/epoxy material. Computed failure indices were compared to corresponding results where the entire web was modeled with shell elements and only a small section of the stiffener foot and panel were modeled locally with solid elements. Including the stiffener web in the local 3D solid model increased the computed failure index. Further including the noodle and transition radius in the local 3D solid model changed the local distribution across the width. The magnitude of the failure index decreased with increasing transition radius and noodle area. For the transition radii modeled, the material properties used for the noodle area had a negligible effect on the results. The results of this study are intended to be used as a guide for conducting finite element and fracture mechanics analyses of delamination and debonding in complex structures such as integrally stiffened panels.

Description: Free-piston Stirling power conversion has been considered a candidate for radioisotope power systems for space for more than a decade. Prior to the free-piston Stirling architecture, systems were designed with kinematic Stirling engines with rotary alternators to convert heat to electricity. These systems were proposed with lightly loaded linkages to achieve the necessary life. When the free-piston configuration was initially proposed, it was thought to be attractive due to the relatively high conversion efficiency, acceptable mass, and the potential for long life and high reliability. These features have consistently been recognized by teams that have studied technology options for radioisotope power systems. Since free-piston Stirling power conversion was first considered for space power applications, there have been major advances in three general areas of development: demonstration of life and reliability, the success achieved by Stirling cryocoolers in flight, and the overall developmental maturity of the technology for both flight and terrestrial applications. Based on these advances, free-piston Stirling convertors are currently being developed for a number of terrestrial applications. They commonly operate with the power, efficiency, life, and reliability as intended, and much of the development now centers on system integration. This paper will summarize the accomplishments of free-piston Stirling power conversion technology over the past decade, review the status, and discuss the challenges that remain.

Description: In order to investigate the fundamental thermal noise limit of a torsion pendulum using a fused silica fiber, we systematically measured and modeled the mechanical losses of thin fused silica fibers coated by electrically conductive thin metal films. Our results indicate that it is possible to achieve a thermal noise limit for coated silica lower by a factor between 3 and 9, depending on the silica diameter, compared to the best tungsten fibers available. This will allow a corresponding increase in sensitivity of torsion pendula used for weak force measurements, including the gravitational constant measurement and ground-based force noise testing for the Laser Interferometer Space Antenna (LISA) mission.

Description: Multiple liquid-metal pump options are reviewed for the purpose of determining the technologies that are best suited for inclusion in a nuclear reactor thermal simulator intended to test prototypical space nuclear system components. Conduction, induction, and thermoelectric electromagnetic pumps are evaluated based on their performance characteristics and the technical issues associated with incorporation into a reactor system. The thermoelectric pump is recommended for inclusion in the planned system at NASA MSFC based on its relative simplicity, low power supply mass penalty, flight heritage, and the promise of increased pump efficiency over earlier flight pump designs through the use of skutterudite thermoelectric elements.

Description: The tensile mechanical properties of ceramic matrix composites (CMC) in directions off the primary axes of the reinforcing fibers are important for architectural design of CMC components that are subjected to multi-axial stress states. In this study, 2D-woven melt-infiltrated (MI) SiC/SiC composite panels with balanced fiber content in the 0 degree and 90 degree directions were tensile loaded in-plane in the 0 degree direction and at 45 degree to this direction. In addition, a 2D triaxially-braided MI composite panel with balanced fiber content in the plus or minus 67 degree bias directions and reduced fiber content in the axial direction was tensile loaded perpendicular to the axial direction tows (i.e., 23 degrees from the bias fibers). Stress-strain behavior, acoustic emission, and optical microscopy were used to quantify stress-dependent matrix cracking and ultimate strength in the panels. It was observed that both off-axis loaded panels displayed higher composite onset stresses for through-thickness matrix cracking than the 2D-woven 0/90 panels loaded in the primary 0 degree direction. These improvements for off-axis cracking strength can in part be attributed to higher effective fiber fractions in the loading direction, which in turn reduces internal stresses on critical matrix flaws for a given composite stress. Also for the 0/90 panel loaded in the 45 degree direction, an improved distribution of matrix flaws existed due to the absence of fiber tows perpendicular to the loading direction. In addition, for the +67/0/-67 braided panel, the axial tows perpendicular to the loading direction were not only low in volume fraction, but were also were well separated from one another. Both off-axis oriented panels also showed relatively good ultimate tensile strength when compared to other off-axis oriented composites in the literature, both on an absolute strength basis as well as when normalized by the average fiber strength within the composites. Initial implications are discussed for constituent and architecture design to improve the directional cracking of SiC/SiC CMC components with MI matrices.

Description: Prototype designs of two separate pumps for use in electric propulsion systems with liquid lithium and bismuth propellants are presented. Both pumps are required to operate at elevated temperatures, and the lithium pump must additionally withstand the corrosive nature of the propellant. Compatibility of the pump materials and seals with lithium and bismuth were demonstrated through proof-of-concept experiments followed by post-experiment visual inspections. The pressure rise produced by the bismuth pump was found to be linear with input current and ranged from 0-9 kPa for corresponding input current levels of 0-30 A, showing good quantitative agreement with theoretical analysis.

Description: The United States (U.S.) Vision for Space Exploration has directed NASA to develop two new launch vehicles for sending humans to the Moon, Mars, and beyond. In January 2006, NASA streamlined its hardware development approach for replacing the Space Shuttle after it is retired in 2010. Benefits of this approach include reduced programmatic and technical risks and the potential to return to the Moon by 2020 by developing the Ares I Crew Launch Vehicle (CLV) propulsion elements now, with full extensibility to future Ares V Cargo Launch Vehicle (CaLV) lunar systems. The Constellation Program selected the Pratt & Whitney Rocketdyne J-2X engine to power the Ares I Upper Stage Element and the Ares V Earth Departure Stage (EDS). This decision was reached during the Exploration Systems Architecture Study and confirmed after the Exploration Launch Projects Office performed a variety of risk analyses, commonality assessments, and trade studies. This paper narrates the evolution of that decision; describes the performance capabilities expected of the J-2X design, including potential commonality challenges and opportunities between the Ares I and Ares V launch vehicles; and provides a current status of J-2X design, development, and hardware testing activities. This paper also explains how the J-2X engine effort mitigates risk by testing existing engine hardware and designs; building on the Apollo Program (1961 to 1975), the Space Shuttle Program (1972 to 2010); and consulting with Apollo era experts to derive other lessons learned to deliver a human-rated engine that is on an aggressive development schedule, with its first demonstration flight in 2012.

Description: Propulsion system efficiency increases as operating temperatures are increased. Some very high-temperature materials are being developed, including refractory metal alloys, carbides, borides, and silicides. System design requires data for materials properties at operating temperatures. Materials property data are not available for many materials of interest at the desired operating temperatures (up to approx. 3000 K). The objective of this work is to provide important physical property data at ultra-high temperatures. The MSFC Electrostatic levitation (ESL) facility can provide measurements of thermophysical properties which include: creep strength, density and thermal expansion for materials being developed for propulsion applications. The ESL facility uses electrostatic fields to position samples between electrodes during processing and characterization studies. Because the samples float between the electrodes during studies, they are free from any contact with a container or test apparatus. This provides a high purity environment for the study of high-temperature, reactive materials. ESL can be used to process a wide variety of materials including metals, alloys, ceramics, glasses and semiconductors. The MSFC ESL has provided non-contact measurements of properties of materials up to 3400 C. Density and thermal expansion are measured by analyzing digital images of the sample at different temperatures. Our novel, non-contact method for measuring creep uses rapid rotation to deform the sample. Digital images of the deformed samples are analyzed to obtain the creep properties, which match those obtained using ASTM Standard E-139 for Nb at 1985 C. Data from selected ESL-based characterization studies will be presented. The ESL technique could support numerous propulsion technologies by advancing the knowledge base and the technology readiness level for ultra-high temperature materials. Applications include non-eroding nozzle materials and lightweight, high-temperature alloys for turbines and structures.

Description: T:hree non-toxic demonstration reaction control engines (RCE) were successfully tested at the Aerojet Sacramento facility under a technology contract sponsored by the National Aeronautics and Space Administration's (NASA) Marshall Space Flight Center (MSFC). The goals of the NASA MSFC contract (NAS8-01109) were to develop and expand the technical maturity of a non-toxic, on-orbit auxiliary propulsion system (APS) thruster under the auspices of the Exploration Systems Mission Directorate. The demonstration engine utilized Liquid Oxygen (LOX) and Ethanol as propellants to produce 870 lbf thrust. The Aerojet RCE's were successfully acceptance tested over a broad range of operating conditions. Steady state tests evaluated engine response to varying chamber pressures and mixture ratios. In addition to the steady state tests, a variety of pulsing tests were conducted over a wide range of electrical pulse widths (EPW). Each EPW condition was also tested over a range of percent duty cycles (DC), and bit impulse and pulsing specific impulse were determined for each of these conditions. Subsequent to acceptance testing at Aerojet, these three engines were delivered to the NASA White Sands Test Facility (WSTF) in April 2005 for incorporation into a cryogenic Auxiliary Propulsion System Test Bed (APSTB). The APSTB is a test article that will be utilized in an altitude test cell to simulate anticipated mission applications. The objectives of this APSTB testing included evaluation of engine performance over an extended duty cycle map of propellant pressure and temperature, as well as engine and system performance at typical mission duty cycles over extended periods of time. This paper provides acceptance test results and a status of the engine performance as part of the system level testing.

Description: A self-healing system for an insulation material initiates a self-repair process by rupturing a plurality of microcapsules disposed on the insulation material. When the plurality of microcapsules are ruptured reactants witlun the plurality of microcapsules react to form a replacement polymer in a break of the insulation material. This self-healing system has the ability to repair multiple breaks in a length of insulation material without exhausting the repair properties of the material.

Description: A method for providing for thermal conduction using an array of carbon nanotubes (CNTs). An array of vertically oriented CNTs is grown on a substrate having high thermal conductivity, and interstitial regions between adjacent CNTs in the array are partly or wholly filled with a filler material having a high thermal conductivity so that at least one end of each CNT is exposed. The exposed end of each CNT is pressed against a surface of an object from which heat is to be removed. The CNT-filler composite adjacent to the substrate provides improved mechanical strength to anchor CNTs in place and also serves as a heat spreader to improve diffusion of heat flux from the smaller volume (CNTs) to a larger heat sink.

Description: The need for a high specific impulse LOX/LH2 pump-fed lunar lander engine has been established by NASA for the new Exploration architecture. Preliminary studies indicate that a 4 engine cluster in the thrust range of 9,000-lbf each is a likely configuration for the main propulsion of the manned lunar lander vehicle. The main Lunar Surface Access Module engines will likely be responsible for mid-course correction burns, lunar orbit insertion burns, a deorbit burn, and the powered descent to the lunar surface. This multi-task engine philosophy imposes a wide throttling requirement on the engines in the range of 10:1. Marshall Space Flight Center has initiated an internal effort to mature the technologies needed for full scale development of such a LOX/LH2 pump-fed engine. In particular, a fuel turbopump is being designed and fabricated at MSFC to address the issues that a small high speed turbopump of this class will face. These issues include adequate throttling performance of the pump and turbine over a very wide operating range. The small scale of the hardware presents issues including performance scaling, and manufacturing issues like that will challenge the traditional methods we have used to fabricate and assemble larger scale turbopumps. The small high speed turbopump being developed at MSFC will operate at speeds greater than 100,000-rpm. These speeds create issues that include structural dynamics and high cycle fatigue as well as rotordynamic stability. The fuel turbopump development at MSFC will address these issues, and plans are in work for component level testing as well as operation in a test bed engine environment. The fuel turbopump design is nearing completion and described herein.

Description: This paper describes the results of a series of experiments aimed at quantifying the plasma propulsion testing capabilities of a 12-ft diameter vacuum facility (12V) at USAF-Arnold Engineering Development Center (AEDC). Vacuum is maintained in the 12V facility by cryogenic panels lining the interior of the chamber. The pumping capability of these panels was shown to be great enough to support plasma thrusters operating at input electrical power 〉20 kW. In addition, a series of plasma diagnostics inside the chamber allowed for measurement of plasma parameters at different spatial locations, providing information regarding the chamber's effect on the global plasma thruster flowfield. The plasma source used in this experiment was Hall thruster manufactured by Busek Co. The thruster was operated at up to 20 kW steady-state power in both a lower current and higher current mode. The vacuum level in the chamber never rose above 9 x 10(exp -6) torr during the course of testing. Langmuir probes, ion flux probes, and Faraday cups were used to quantify the plasma parameters in the chamber. We present the results of these measurements and estimates of pumping speed based on the background pressure level and thruster propellant mass flow rate.

Description: An international team, lead by Tokyo Metropolitan University, is developing a mission concept for a suborbital test of orbital-motion-limited (OML) bare-wire anode current collection for application to electrodynamic tether propulsion. The tether is a tape with a 50-mm width, 0.05-mm thickness, and 1-km length. This will be the first space test of the OML theory. In addition, by being an engineering demonstration (of space tethers), the mission will demonstrate electric beam generation for "sounding" determination of the neutral density profile in the ionospheric "E-layer." If selected by the Institute of Space and Astronautical Science/Japanese Aerospace Exploration Agency (JAXA), the mission will launch in early 2009 using an $520 Sounding Rocket. During ascent, and above =100 km in attitude, the 1-km tape tether will be deployed at a rate of 8 m/s. Once deployed, the tape tether will serve as an anode, collecting ionospheric electrons. The electrons will be expelled into space by a hollow cathode device, thereby completing the circuit and allowing current to flow.This paper will describe the objectives of the proposed mission, the technologies to be employed, and the application of the results to future space missions using electrodynamic tethers for propulsion or power generation.

Description: Solar sails provide a propellant free form for space propulsion. These are large flat surfaces that generate thrust when they are impacted by light. When attached to a space vehicle, the thrust generated can propel the space vehicle to great distances at significant speeds. For optimal performance the sail must be kept from excessive vibration. Active control techniques can provide the best performance. However, they require an external power-source that may create significant parasitic mass to the solar sail. However, solar sails require low mass for optimal performance. Secondly, active control techniques typically require a good system model to ensure stability and performance. However, the accuracy of solar sail models validated on earth for a space environment is questionable. An alternative approach is passive vibration techniques. These do not require an external power supply, and do not destabilize the system. A third alternative is referred to as semi-active control. This approach tries to get the best of both active and passive control, while avoiding their pitfalls. In semi-active control, an active control law is designed for the system, and passive control techniques are used to implement it. As a result, no external power supply is needed so the system is not destabilize-able. Though it typically underperforms active control techniques, it has been shown to out-perform passive control approaches and can be unobtrusively installed on a solar sail boom. Motivated by this, the objective of this research is to study the suitability of a Piezoelectric (PZT) patch actuator/sensor based semi-active control system for the vibration suppression problem of solar sail booms. Accordingly, we develop a suitable mathematical and computer model for such studies and demonstrate the capabilities of the proposed approach with computer simulations.

Description: A double vacuum bag molding assembly with improved void management and laminate net shape control which provides a double vacuum enviromnent for use in fabricating composites from prepregs containing air and/or volatiles such as reactive resin matrix composites or composites from solvent containing prepregs with non-reactive resins matrices. By using two vacuum environments during the curing process, a vacuum can be drawn during a B-stage of a two-step cycle without placing the composite under significant relative pressure. During the final cure stage, a significant pressure can be applied by releasing the vacuum in one of the two environments. Inner and outer bags are useful for creating the two vacuum environments with a perforated tool intermediate the two. The composite is placed intermediate a tool plate and a caul plate in the first environment with the inner bag and tool plate defining the first environment. The second environment is characterized by the outer bag which is placed over the inner bag and the tool plate.

Description: A method for increasing the compressive modulus of aerogels comprising: providing aerogel substrate comprising a bubble matrix in a chamber; providing monomer to the chamber, the monomer comprising vapor phase monomer which polymerizes substantially free of polymerization byproducts; depositing monomer from the vapor phase onto the surface of the aerogel substrate under deposition conditions effective to produce a vapor pressure sufficient to cause the vapor phase monomer to penetrate into the bubble matrix and deposit onto the surface of the aerogel substrate, producing a substantially uniform monomer film; and, polymerizing the substantially uniform monomer film under polymerization conditions effective to produce polymer coated aerogel comprising a substantially uniform polymer coating substantially free of polymerization byproducts.Polymer coated aerogel comprising aerogel substrate comprising a substantially uniform polymer coating, said polymer coated aerogel comprising porosity and having a compressive modulus greater than the compressive modulus of the aerogel substrate, as measured by a 100 lb. load cell at 1 mm/minute in the linear range of 20% to 40% compression.

Description: The invention provides polymer blends containing polyhydroxyamide and one or more flammable polymers. The polymer blends are flame retardant and have improved durability and heat stability compared to the flammable polymer portion of the blends. Articles containing the polymer blends are also provided.

Description: A solid carbon has CNTs dispersed therein and is formed about three-dimensionally ordered spherical voids arranged in an opal-like lattice.

Description: The NASA In-Space Propulsion Technology Project Office initiated a preliminary study to evaluate the performance benefits of a solar electric propulsion (SEP) upper-stage with existing and near-term small launch vehicles. The analysis included circular and elliptical Low Earth Orbit (LEO) to Geosynchronous Earth Orbit (GEO) transfers, and LEO to Low Lunar Orbit (LLO) applications. SEP subsystem options included state-of-the-art and near-term solar arrays and electric thrusters. In-depth evaluations of the Aerojet BPT-4000 Hall thruster and NEXT gridded ion engine were conducted to compare performance, cost and revenue potential. Preliminary results indicate that Hall thruster technology is favored for low-cost, low power SEP stages, while gridded-ion engines are favored for higher power SEP systems unfettered by transfer time constraints. A low-cost point design is presented that details one possible stage configuration and outlines system limitations, in particular fairing volume constraints. The results demonstrate mission enhancements to large and medium class launch vehicles, and mission enabling performance when SEP system upper stages are mounted to low-cost launchers such as the Minotaur and Falcon 1. Study results indicate the potential use of SEP upper stages to double GEO payload mass capability and to possibly enable launch on demand capability for GEO assets. Transition from government to commercial applications, with associated cost/benefit analysis, has also been assessed. The sensitivity of system performance to specific impulse, array power, thruster size, and component costs are also discussed.

Description: The Floating Potential Measurement Unit (FPMU) is a multi-probe package designed to measure the floating potential of the 1nternational Space Station (ISS) as well as the density and temperature of the local ionospheric plasma environment. The role oj the FPMU is to provide direct measurements of ISS spacecraft charging as continuing construction leads to dramatic changes in ISS size and configuration. FPMU data are used for refinement and validation of the ISS spacecraft charging models used to evaluate the severity and frequency of occurrence of ISS charging hazards. The FPMU data and the models are also used to evaluate the effectiveness of proposed hazard controls. The FPMU consists of four probes: a floating potential probe, two Langmuir probes. and a plasma impedance probe. These probes measure the floating potential of the ISS, plasma density, and electron temperature. Redundant measurements using different probes support data validation by inter-probe comparisons. The FPMU was installed by ISS crewmembers, during an ExtraVehicular Activity, on the starboard (Sl) truss of the ISS in early August 2006, when the ISS incorporated only one 160V US photovoltaic (PV) array module. The first data campaign began a few hours after installation and continued for over five days. Additional data campaigns were completed in 2007 after a second 160V US PV array module was added to the ISS. This paper discusses the general performance characteristics of the FPMU as integrated on ISS, the functional performance of each probe, the charging behavior of the ISS before and after the addition of a second 160V US PV array module, and initial results from model comparisons.

Description: Neural network-based adaptive control is considered for active control of a highly flexible truss structure which may be used to support solar sail membranes. The objective is to suppress unwanted vibrations in SAFE (Solar Array Flight Experiment) boom, a test-bed located at NASA. Compared to previous tests that restrained truss structures in planar motion, full three dimensional motions are tested. Experimental results illustrate the potential of adaptive control in compensating for nonlinear actuation and modeling error, and in rejecting external disturbances.

Description: Operation of Hall thrusters with bismuth propellant has been shown to be a promising path toward high-power, high-performance, long-lifetime electric propulsion for spaceflight missions [1]. There has been considerable effort in the past three years aimed at resuscitating this promising technology and validating earlier experimental results indicating the advantages of a bismuth-fed Hall thruster. A critical element of the present effort is the precise metering of propellant to the thruster, since performance cannot be accurately assessed without an accurate accounting of mass flow rate. Earlier work used a pre./post-test propellant weighing scheme that did not provide any real-time measurement of mass flow rate while the thruster was firing, and makes subsequent performance calculations difficult. The motivation of the present work is to develop a precision liquid bismuth Propellant Management System (PMS) that provides hot, molten bismuth to the thruster while simultaneously monitoring in real-time the propellant mass flow rate. The system is a derivative of our previous propellant feed system [2], but the present system represents a more compact design. In addition, all control electronics are integrated into a single unit and designed to reside on a thrust stand and operate in the relevant vacuum environment where the thruster is operating, significantly increasing the present technology readiness level of liquid metal propellant feed systems. The design of various critical components in a bismuth PMS are described. These include the bismuth reservoir and pressurization system, 'hotspot' flow sensor, power system and integrated control system. Particular emphasis is given to selection of the electronics employed in this system and the methods that were used to isolate the power and control systems from the high-temperature portions of the feed system and thruster. Open loop calibration test results from the 'hotspot' flow sensor are reported, and results of integrated thruster/PMS tests demonstrate operation of the feed system in the relevant environment.

Description: Radiation or passively cooled thrust chambers are used for a variety of chemical propulsion functions including apogee insertion, reaction control for launch vehicles, and primary propulsion for planetary spacecraft. The performance of these thrust chambers is limited by the operating temperature of available materials. Improved oxidation resistance and increased operating temperatures can be achieved with the use of thermal barrier coatings such as zirconium oxide (ZrO2) and hafnium oxide (HfO2). However, previous attempts to include these materials showed cracking and spalling of the oxide layer due to poor bonding. Current research at NASA's Marshall Space Flight Center (MSFC) has generated unique, high temperature material options for in-space thruster designs that are capable of up to 2500 C operating temperatures. The research is focused on fabrication technologies to form low cost Iridium,qF_.henium (Ir/Re) components with a ceramic hot wall created as an integral, functionally graded material (FGM). The goal of this effort is to further de?celop proven technologies for embedding a protective ceramic coating within the Ir/Re liner to form a robust functional gradient material. Current work includes the fabrication and testing of subscale samples to evaluate tensile, creep, thermal cyclic/oxidation, and thermophysical material properties. Larger test articles have also being fabricated and hot-fire tested to demonstrate the materials in prototype thrusters at 1O0 lbf thrust levels.

Description: The Tethered Satellite System Space Shuttle missions, TSS-1 in 1993 and TSS-1R in 1996, were the height of space tether technology development. Since NASA's investment of some $200M and two Shuttle missions in those two pioneering missions, there have been several smaller tether flight experiments, but interest in this promising technology has waned within NASA as well as the DOD agencies. This is curious in view of the unique capabilities of space tether systems and the fact that they have been flight validated and shown to perform as, or better than, expected in earth orbit. While it is true that the TSS-1, TSS-1R and SEDS-2 missions experienced technical difficulties, the causes of these early developmental problems are now known to be design or materials flaws that are (1) unrelated to the basic viability of space tether technology, and (2) they are readily corrected. The purpose of this paper is to review the dynamic and electrodynamic fundamentals of space tethers and the unique capabilities they afford (that are enabling to certain types of space missions); to elucidate the nature, cause, and solution of the early developmental problems; and to provide an update on progress made in development of the technology. Finally, it is shown that (1) all problems experienced during early development of the technology now have solutions; and (2) the technology has been matured by advances made in strength and robustness of tether materials, high voltage engineering in the space environment, tether health and status monitoring, and the elimination of the broken tether hazard. In view of this, it is inexplicable why this flight-validated technology has not been utilized in the past decade, considering the powerful and unique capabilities that space tethers can afford that are, not only required to carryout, otherwise, unobtainable missions, but can also greatly reduce the cost of certain on-going space operations.

Description: A method of growing carbon nanotubes uses a synthesized mesoporous si lica template with approximately cylindrical pores being formed there in. The surfaces of the pores are coated with a carbon nanotube precu rsor, and the template with the surfaces of the pores so-coated is th en heated until the carbon nanotube precursor in each pore is convert ed to a carbon nanotube.

Description: A method for fabricating a piezoelectric macro-fiber composite actuator comprises making a piezoelectric fiber sheet by providing a plurality of wafers of piezoelectric material, bonding the wafers together with an adhesive material to from a stack of alternating layers of piezoelectric material and adhesive material, and cutting through the stack in a direction substantially parallel to the thickness of the stack and across the alternating layers of piezoelectric material and adhesive material to provide at least one piezoelectric fiber sheet having two sides comprising a plurality of piezoelectric fibers in juxtaposition to the adhesive material. The method further comprises bonding two electrically conductive films to the two sides of the piezoelectric fiber sheet. At least one conductive film has first and second conductive patterns formed thereon which are electrically isolated from one another and in electrical contact with the piezoelectric fiber sheet.

Description: This viewgraph presentation reviews the use of the binding procurement process in purchasing Aerospace Flight Battery Systems. NASA Engineering and Safety Center (NESC) requested NASA Aerospace Flight Battery Systems Working Group to develop a set of guideline requirements document for Binding Procurement Contracts.

Description: In support of the President's Vision for Space Exploration, the Nuclear Thermal Rocket (NTR) concept is being evaluated as a potential propulsion technology for human expeditions to the moon and Mars. The need for exceptional propulsion system performance in these missions has been documented in numerous studies, and was the primary focus of a considerable effort undertaken during the 1960's and 1970's. The NASA Glenn Research Center is leveraging this past NTR investment in their vehicle concepts and mission analysis studies with the aid of the Nuclear Engine System Simulation (NESS) code. This paper presents the additional capabilities and upgrades made to this code in order to perform higher fidelity NTR propulsion system analysis and design.

Description: Studies were conducted for the In-Space Propulsion (ISP) Ultralightweight Ballute Technology Development Program to increase the technical readiness level of inflatable decelerator systems for planetary aerocapture. The present experimental study was conducted to develop the capability for testing lightweight, flexible materials in hypersonic facilities. The primary objectives were to evaluate advanced polymer film materials in a high-temperature, high-speed flow environment and provide experimental data for comparisons with fluid-structure interaction modeling tools. Experimental testing was conducted in the Langley Aerothermodynamics Laboratory 20-Inch Hypersonic CF4 and 31-Inch Mach 10 Air blowdown wind tunnels. Quantitative flexure measurements were made for 60 degree half angle afterbody-attached ballutes, in which polyimide films (1-mil and 3- mil thick) were clamped between a 1/2-inch diameter disk and a base ring (4-inch and 6-inch diameters). Deflection measurements were made using a parallel light silhouette of the film surface through an existing schlieren optical system. The purpose of this paper is to discuss these results as well as free-flying testing techniques being developed for both an afterbody-attached and trailing toroidal ballute configuration to determine dynamic fluid-structural stability. Methods for measuring polymer film temperature were also explored using both temperature sensitive paints (for up to 370 C) and laser-etched thin-film gages.

Description: NASA LaRC has established a new protocol for visualizing the nanomaterials in structural polymer matrix resins. Using this new technique and reconstructing the 3D distribution of the nanomaterials allows us to compare this distribution against a theoretically perfect distribution. Additional tertiary structural information can now be obtained and quantified with the electron tomography studies. These tools will be necessary to establish the structural-functional relationships between the nano and the bulk. This will also help define the critical length scales needed for functional properties. Field ready tool development and calibration can begin by using these same samples and comparing the response. i.e. gold standards of good and bad dispersion.

Description: This viewgraph presentation reviews the lunar lander propulsion requirements. It includes discussion on: Lander Project Overview, Project Evolution/Design Cycles, Lunar Architecture & Lander Reference Missions, Lander Concept Configurations, Descent and Ascent propulsion reviews, and a review of the technology requirements.

Description: An actively pumped alkali metal flow circuit, designed and fabricated at the NASA Marshall Space Flight Center, is currently undergoing testing in the Early Flight Fission Test Facility (EFF-TF). Sodium potassium (NaK), which was used in the SNAP-10A fission reactor, was selected as the primary coolant. Basic circuit components include: simulated reactor core, NaK to gas heat exchanger, electromagnetic (EM) liquid metal pump, liquid metal flowmeter, load/drain reservoir, expansion reservoir, test section, and instrumentation. Operation of the circuit is based around a 37-pin partial-array core (pin and flow path dimensions are the same as those in a full core), designed to operate at 33 kWt. NaK flow rates of greater than 1 kg/sec may be achieved, depending upon the power applied to the EM pump. The heat exchanger provides for the removal of thermal energy from the circuit, simulating the presence of an energy conversion system. The presence of the test section increases the versatility of the circuit. A second liquid metal pump, an energy conversion system, and highly instrumented thermal simulators are all being considered for inclusion within the test section. This paper summarizes the capabilities and ongoing testing of the Fission Surface Power Primary Test Circuit (FSP-PTC).

Description: Fracture control requirements have been developed to address damage tolerance of composites for manned space flight hardware. The requirements provide the framework for critical and noncritical hardware assessment and testing. The need for damage threat assessments, impact damage protection plans, and nondestructive evaluation are also addressed. Hardware intended to be damage tolerant have extensive coupon, sub-element, and full-scale testing requirements in-line with the Building Block Approach concept from the MIL-HDBK-17, Department of Defense Composite Materials Handbook.

Description: The energy requirements of hypothetical, propellant-less space drives are compared to rockets. This serves to provide introductory estimates for potential benefits and to suggest analytical approaches for further study. A "space drive" is defined as an idealized form of propulsion that converts stored potential energy directly into kinetic energy using only the interactions between the spacecraft and its surrounding space. For Earth-to-orbit, the space drive uses 3.7 times less energy. For deep space travel, energy is proportional to the square of delta-v, whereas rocket energy scales exponentially. This has the effect of rendering a space drive 150-orders-of-magnitude better than a 17,000-s Specific Impulse rocket for sending a modest 5000 kg probe to traverse 5 ly in 50 years. Indefinite levitation, which is impossible for a rocket, could conceivably require 62 MJ/kg for a space drive. Assumption sensitivities and further analysis options are offered to guide further inquires.

Description: The layered morphology of silicate clay provides an effective barrier to oxidative degradation of the matrix resin. However, as resin thermal stability continues to reach higher limits, development of an organic modification with comparable temperature capabilities becomes a challenge. Typically, phyllosilicates used in polymer nanocomposites are modified with an alkyl ammonium ion. Such organic modifiers are not suited for incorporation into high temperature polymers as they commonly degrade below 200oC. Therefore, the development of nanoparticle specifically suited for high temperature applications is necessary. Several nanoparticles were investigated in this study, including pre-exfoliated synthetic clay, an organically modified clay, and carbon nanofiber. Dispersion of the layered silicate increases the onset temperature of matrix degradation as well as slows oxidative degradation. The thermally stable carbon nanofibers are also observed to significantly increase the resin thermal stability.

Description: As carbon/epoxy materials became more prevalent in the aerospace industry, microstructural analysis demanded specimen preparation techniques that led to better polished surfaces, achievable in a shorter time, and using fewer steps. The desire to use image analysis for material characterization also helped drive the goal for defect free surfaces. At NASA-Langley (LaRC), carbon/epoxy specimens had been historically prepared in 1 inch diameter Bakelite mounts. Carbon/epoxy specimens that were 1/8 to 1/4 inch thick were not affected by the heat and pressure required for mounting in Bakelite, however thinner specimens were crushed during mounting. A two-part room temperature curing epoxy was chosen as an alternative but sometimes voids developed between the specimen and the mounting material. This was prevented by either heating the epoxy to 140 degrees F to lower the viscosity of the epoxy or by using a vacuum impregnation apparatus. Both techniques helped facilitate flow and allowed the epoxy to penetrate crevices.

Description: Over the past several years, efforts have been under way to design and develop an operationally flexible research facility for investigating the use of cross-field MHD accelerators as a potential thrust augmentation device for thermal propulsion systems, The baseline configuration for this high-power experimental facility utilizes a 1,5-MW, multi-gas arc-heater as a thermal driver for a 2-MW, MHD accelerator, which resides in a large-bore 2-tesla electromagnet. A preliminary design study using NaK seeded nitrogen as the working fluid led to an externally diagonalized segmented MHD channel configuration based on an expendable beat-sink design concept. The current status report includes a review of engineering/design work and performance optimization analyses and summarizes component hardware fabrication and development efforts, preliminary testing results, and recent progress toward full-up assembly and testing

Description: A rectangular, variable-stiffness panel with tow overlaps was manufactured using an advanced tow placement machine. The cured panel had large anticlastic imperfections, with measured amplitudes of over two times the average panel thickness. These imperfections were not due to the overall steered-fiber layup or the tow overlaps, but instead resulted from local asymmetries in the laminate that were caused by a manufacturing oversight. In the nominal panel layup, fiber angles vary linearly from 60 degrees on the panel axial centerline to 30 degrees on the parallel edges. A geometrically nonlinear analysis was performed with a -280 degree Fahrenheit thermal load to simulate the postcure cooldown to room temperature. The predicted geometric imperfections correlated well with the measured panel shape. Unique structural test fixtures were then developed which greatly reduced these imperfections, but they also caused prestresses in the panel. Surface imperfections measured after the panel was installed in the test fixtures were used with nonlinear finite element analyses to predict these fixturing-induced prestresses. These prestresses were also included in structural analyses of panel end compression to failure, and the analytical results compared well with test data when both geometric and material nonlinearities were included.

Description: A methodology that accounts for both delaminaton onset and growth in composite structural components is proposed for improved fatigue life prediction to reduce life cycle costs and improve accept/reject criteria for manufacturing flaws. The benefits of using a Delamination Onset Threshold (DOT) approach in combination with a Modified Damage Tolerance (MDT) approach is highlighted. The use of this combined approach to establish accept/reject criteria, requiring less conservative initial manufacturing flaw sizes, is illustrated.

Description: This paper presents a mathematical analysis of the heat transfer processes taking place in a radiator for a closed cycle gas turbine (CCGT), also referred to as a Closed Brayton Cycle (CBC) space power system. The resulting equations and relationships have been incorporated into a radiator sub-routine of a numerical triple objective CCGT optimization program to determine operating conditions yielding maximum cycle efficiency, minimum radiator area and minimum overall systems mass. Study results should be of interest to numerical modeling of closed cycle Brayton space power systems and to the design of fluid cooled radiators in general.

Description: Effective thermal conductivity of a high porosity unbonded silica fibrous insulation specimen was measured over a pressure range of 0.001 to 750 torr (0.1 to 101.3 x 10(exp 3) Pa), and with large temperature gradients maintained across the sample thickness: hot side temperature range of 360 to 1360 K, with the cold side at room temperature. The measurements were compared with the theoretical solution of combined radiation/conduction heat transfer. The previously developed radiation heat transfer model used in this study is based on a modified diffusion approximation, and uses deterministic parameters that define the composition and morphology of the medium: distributions of fiber size and orientation, fiber volume fractions, and the spectral complex refractive index of the fibers. The close agreement between experimental and theoretical data further verifies the theoretical model over a wide range of temperatures and pressures.

Description: The development of flexible, thermally conductive fabrics and plastic tubes for the Liquid Cooling and Ventilation Garment (LCVG) are needed to reduce weight and improve the mobility, comfort, and performance of future spacesuits. Such improvements would allow astronauts to operate more efficiently and safely for extended extravehicular activities. As a continuation of our work on the improvement of thermal conductivity (TC) of polymeric materials, nanocomposites were prepared from copoly(ethylene vinyl acetate), trade name Elvax 260TradeMark), metallized carbon nanofibers (CNFs), nickel (Ni) nanostrands, boron nitride both alone and as mixtures with aluminum powder. The nanocomposites were prepared by melt mixing at various loading levels and subsequently fabricated into several material forms (i.e., ribbons, tubes, and compression molded plaques) for analysis. Ribbons and tubes were extruded to form samples in which the nanoparticles were aligned in the direction of flow. The degree of dispersion and alignment of the nanoparticles were investigated using high-resolution scanning electron microscopy. Tensile properties of the aligned samples were determined at room temperature. TC measurements were performed using a laser flash (Nanoflash(TradeMark) technique. The TC of the samples was measured in both the direction of alignment as well as transverse. Tubing of comparable dimensions to that used in the LCVG was extruded from select compositions and the thermal conductivities of the tubes measured.

Description: The contribution made by orbiting solid rocket motors (SRMs) to the orbital debris environment is both potentially significant and insufficiently studied. A combination of rocket motor design and the mechanisms of the combustion process can lead to the emission of sufficiently large and numerous by-products to warrant assessment of their contribution to the orbital debris environment. These particles are formed during SRM tail-off, or the termination of burn, by the rapid expansion, dissemination, and solidification of the molten Al2O3 slag pool accumulated during the main burn phase of SRMs utilizing immersion-type nozzles. Though the usage of SRMs is low compared to the usage of liquid fueled motors, the propensity of SRMs to generate particles in the 100 m and larger size regime has caused concern regarding their contributing to the debris environment. Particle sizes as large as 1 cm have been witnessed in ground tests conducted under vacuum conditions and comparable sizes have been estimated via ground-based telescopic and in-situ observations of sub-orbital SRM tail-off events. Using sub-orbital and post recovery observations, a simplistic number-size-velocity distribution of slag from on-orbit SRM firings was postulated. In this paper we have developed more elaborate distributions and emission scenarios and modeled the resultant orbital population and its time evolution by incorporating a historical database of SRM launches, propellant masses, and likely location and time of particulate deposition. From this analysis a more comprehensive understanding has been obtained of the role of SRM ejecta in the orbital debris environment, indicating that SRM slag is a significant component of the current and future population.

Description: A unique probabilistic theory is described to predict the properties of nanocomposites. The simulation is based on composite micromechanics with progressive substructuring down to a nanoscale slice of a nanofiber where all the governing equations are formulated. These equations have been programmed in a computer code. That computer code is used to simulate uniaxial strengths properties of a mononanofiber laminate. The results are presented graphically and discussed with respect to their practical significance. These results show smooth distributions.

Description: Composite Overwrapped Pressure Vessels (COPVs) are often used for storing pressurant gases onboard spacecraft. Kevlar (DuPont), glass, carbon and other more recent fibers have all been used as overwraps. Due to the fact that overwraps are subjected to sustained loads for an extended period during a mission, stress rupture failure is a major concern. It is therefore important to ascertain the reliability of these vessels by analysis, since the testing of each flight design cannot be completed on a practical time scale. The present paper examines specifically a Weibull statistics based stress rupture model and considers the various uncertainties associated with the model parameters. The paper also examines several reliability estimate measures that would be of use for the purpose of recertification and for qualifying flight worthiness of these vessels. Specifically, deterministic values for a point estimate, mean estimate and 90/95 percent confidence estimates of the reliability are all examined for a typical flight quality vessel under constant stress. The mean and the 90/95 percent confidence estimates are computed using Monte-Carlo simulation techniques by assuming distribution statistics of model parameters based also on simulation and on the available data, especially the sample sizes represented in the data. The data for the stress rupture model are obtained from the Lawrence Livermore National Laboratories (LLNL) stress rupture testing program, carried out for the past 35 years. Deterministic as well as probabilistic sensitivities are examined.

Description: From 1999 to 2006, the NASA Glenn Research Center (GRC) supported the development of a high-efficiency, nominal 110-We Stirling Radioisotope Generator (SRG110) for potential use on NASA missions, including deep space missions, Mars rovers, and lunar applications. Lockheed Martin (LM) was the system integrator for the SRG110, under contract to the Department of Energy (DOE). Infinia Corporation (formerly Stirling Technology Company) developed the Stirling convertor, first as a contractor to DOE and then under subcontract to LM. The SRG110 development has been redirected, and recent program changes have been made to significantly increase the specific power of the generator. System development of an Advanced Stirling Radioisotope Generator (ASRG) has now begun, using a lightweight, advanced convertor from Sunpower, Inc. This paper summarizes the results of the supporting technology effort that GRC completed for the SRG110. GRC tasks included convertor extended-duration testing in air and thermal vacuum environments, heater head life assessment, materials studies, permanent magnet aging characterization, linear alternator evaluations, structural dynamics testing, electromagnetic interference (EMI) and electromagnetic compatibility (EMC) characterization, organic materials evaluations, reliability studies, and development of an end-to-end system dynamic model. Related efforts are now continuing in many of these areas to support ASRG development.

Description: Oxides having magnetoplumbite structure are promising candidate materials for applications as high temperature thermal barrier coatings because of their high thermal stability, high thermal expansion, and low thermal conductivity. In this study, powders of LaMgAl11O19, GdMgAl11O19, SmMgAl11O19, and Gd0.7Yb0.3MgAl11O19 magnetoplumbite oxides were synthesized by citric acid sol-gel method and hot pressed into disk specimens. The thermal expansion coefficients (CTE) of these oxide materials were measured from room temperature to 1500 C. The average CTE value was found to be approx.9.6x10(exp -6)/C. Thermal conductivity of these magnetoplumbite-based oxide materials was also evaluated using steady-state laser heat flux test method. The effects of doping on thermal properties were also examined. Thermal conductivity of the doped Gd0.7Yb0.3MgAl11O19 composition was found to be lower than that of the undoped GdMgAl11O19. In contrast, thermal expansion coefficient was found to be independent of the oxide composition and appears to be controlled by the magnetoplumbite crystal structure. Thermal conductivity testing of LaMgAl11O19 and LaMnAl11O19 magnetoplumbite oxide coatings plasma sprayed on NiCrAlY/Rene N5 superalloy substrates indicated resistance of these coatings to sintering even at temperatures as high as 1600 C.

Description: A parametric study using a two-flux approximation of the radiative transfer equation was performed to examine the effects of an opaque reflective layer on the thermal behavior of a typical semitransparent thermal barrier coating on an opaque substrate. Some ceramic materials are semitransparent in the wavelength ranges where thermal radiation is important. Even with an opaque layer on each side of the semitransparent thermal barrier coating, scattering and absorption can have an effect on the heat transfer. In this work, a thermal barrier coating that is semitransparent up to a wavelength of 5 micrometers is considered. Above 5 micrometers wavelength, the thermal barrier coating is opaque. The absorption and scattering coefficient of the thermal barrier was varied. The thermal behavior of the thermal barrier coating with an opaque reflective layer is compared to a thermal barrier coating without the reflective layer. For a thicker thermal barrier coating with lower convective loading, which would be typical of a combustor liner, a reflective layer can significantly decrease the temperature in the thermal barrier coating and substrate if the scattering is weak or moderate and for strong scattering if the absorption is large. The layer without the reflective coating can be about as effective as the layer with the reflective coating if the absorption is small and the scattering strong. For low absorption, some temperatures in the thermal barrier coating system can be slightly higher with the reflective layer. For a thin thermal barrier coating with high convective loading, which would be typical of a blade or vane that sees the hot sections of the combustor, the reflective layer is not as effective. The reflective layer reduces the surface temperature of the reflective layer for all conditions considered. For weak and moderate scattering, the temperature of the TBC-substrate interface is reduced but for strong scattering, the temperature of the substrate is increased slightly.

Description: The NASA Engineering and Safety Center (NESC) has been conducting an independent technical assessment to address safety concerns related to the known stress rupture failure mode of filament wound pressure vessels in use on Shuttle and the International Space Station. The Shuttle's Kevlar-49 fiber overwrapped tanks are of particular concern due to their long usage and the poorly understood stress rupture process in Kevlar-49 filaments. Existing long term data show that the rupture process is a function of stress, temperature and time. However due to the presence of load sharing liners and the complex manufacturing procedures, the state of actual fiber stress in flight hardware and test articles is not clearly known. Indeed non-conservative life predictions have been made where stress rupture data and lifing procedures have ignored the contribution of the liner in favor of applied pressure as the controlling load parameter. With the aid of analytical and finite element results, this paper examines the fundamental mechanical response of composite overwrapped pressure vessels including the influence of elastic-plastic liners and degraded/creeping overwrap properties. Graphical methods are presented describing the non-linear relationship of applied pressure to Kevlar-49 fiber stress/strain during manufacturing, operations and burst loadings. These are applied to experimental measurements made on a variety of vessel systems to demonstrate the correct calibration of fiber stress as a function of pressure. Applying this analysis to the actual qualification burst data for Shuttle flight hardware revealed that the nominal fiber stress at burst was in some cases 23% lower than what had previously been used to predict stress rupture life. These results motivate a detailed discussion of the appropriate stress rupture lifing philosophy for COPVs including the correct transference of stress rupture life data between dissimilar vessels and test articles.

Description: It has been well established that a few ppmw sulfur impurity may segregate to the interface of thermally grown alumina scales and the underlying substrate, resulting in bond degradation and premature spallation. This has been shown for NiAl and NiCrAl-based alloys, bare single crystal superalloys, or coated superalloys. The role of reactive elements (especially Y) has been to getter the sulfur in the bulk and preclude interfacial segregation. Pt additions are also very beneficial, however a similar thermodynamic explanation does not apply. The purpose of the present discussion is to highlight some observations of these effects on Rene'142, Rene'N5, PWA1480, and PWA1484. For PWA1480, we have mapped cyclic oxidation and spallation in terms of potential sulfur interfacial layers and found that a cumulative amount of about one monolayer is sufficient to degrade long term adhesion. Depending on substrate thickness, optimum performance occurs if sulfur is reduced below about 0.2-0.5 ppmw. This is accomplished in the laboratory by hydrogen annealing or commercially by melt-fluxing. Excellent 1150 C cyclic oxidation is thus demonstrated for desulfurized Rene'142, Rene'N5, and PWA1484. Alternatively, a series of N5 alloys provided by GE-AE have shown that as little as 15 ppmw of Y dopant was effective in providing remarkable scale adhesion. In support of a Y-S gettering mechanism, hydrogen annealing was unable to desulfurize these alloys from their initial level of 5 ppmw S. This impurity and critical doping level corresponds closely to YS or Y2S3 stoichiometry. In many cases, Y-doped alloys or alloys with marginal sulfur levels exhibit an oxidative sensitivity to the ambient humidity called Moisture-Induced Delayed Spallation (MIDS). After substantial scale growth, coupled with damage from repeated cycling, cold samples may spall after a period of time, breathing on them, or immersing them in water. While stress corrosion arguments may apply, we propose that the underlying cause is related to a hydrogen embrittlement reaction: Al alloy + 3 H2O = Al(OH)3 + 3H(+) + 3e(-). This mechanism is derived from an analogous moisture-induced hydrogen embrittlement mechanism originally shown for Ni3Al and FeAl intermetallics. Consequently, a cathodic hydrogen charging technique was used to demonstrate that electrolytic de-scaling occurs for these otherwise adherent alumina scales formed on Y-doped Rene'N5, in support of hydrogen effects. Finally, some TBC observations are discussed in light of all of the above. Plasma sprayed 8YSZ coatings, produced on PWA1484 without a bond coat, were found to survive more than 1000 1-hr cycles at 1100 C when desulfurized to below 0.1 ppmw. At higher sulfur (1.2 ppmw) levels, moisture sensitivity and delayed TBC failure, referred to as Desk Top Spallation, occurred at just 200 hr. Despite a large degree of scatter, a factor of 5 in life improvement is indicated for desulfurized samples in cyclic furnace tests, confirming the beneficial effect of low sulfur alloys on model TBC systems. (DTS and moisture effects are also observed on commercially applied PVD 7YSZ coatings on Rene'N5+Y with Pt-aluminide bond coats). These types of catastrophic failure were subverted on the model system by segmenting the substrate into a network of 0.010 high ribs, spaced in. apart, prior to plasma spraying. No failures occurred after 1000 cycles at 1150 C or after 2000 cycles at 1100 C, even after water immersion. The benefit is described in terms of elasticity models and a critical buckling stress.

Description: Weight, durability and performance are all major concerns for any NASA mission. Use of lightweight materials, such as fiber reinforced polymer matrix composites can lead to significant reductions in vehicle weight and improvements in vehicle performance. Research in the Polymeric Materials Branch at NASA Glenn is focused on improving the durability, properties, processability and performance of polymeric materials by utilizing both conventional polymer science and engineering as well as nanotechnology and bioinspired approaches. This presentation will provide an overview of these efforts and highlight recent progress.

Description: First, we will review our most important research accomplishments from a five year study concerned with the prediction of mechanical properties of unidirectional and woven graphite/polyimide composites based on T650-35, M40J and M60J fibers embedded in either PMR-15 or PMR-II-50 polyimide resins. Then, an aging model recently developed for the composites aged in nitrogen will be proposed and experimentally verified on an eight harness satin (8HS) woven T650-35/PMR-15 composite aged in nitrogen at 315 C for up to 1500 hours. The study was supported jointly between 1999 and 2005 by the AFOSR, the NASA Glenn Research Center, and the National Science Foundation.

Description: Spacecraft radiators reject heat to their surroundings. Radiators can be deployable or mounted on the body of the spacecraft. NASA's Crew Exploration Vehicle is to use body mounted radiators. Coatings play an important role in heat rejection. The coatings provide the radiator surface with the desired optical properties of low solar absorptance and high infrared emittance. These specialized surfaces are applied to the radiator panel in a number of ways, including conventional spraying, plasma spraying, or as an applique. Not specifically designed for a weathering environment, little is known about the durability of conventional paints, coatings, and appliques upon exposure to weathering and subsequent exposure to solar wind and ultraviolet radiation exposure. In addition to maintaining their desired optical properties, the coatings must also continue to adhere to the underlying radiator panel. This is a challenge, as new composite radiator panels are being considered as replacements for the aluminum panels used previously. Various thermal control paints, coatings, and appliques were applied to aluminum and isocyanate ester composite coupons and were exposed for 30 days at the Atmospheric Exposure Site of the Kennedy Space Center s Beach Corrosion Facility for the purpose of identifying their durability to weathering. Selected coupons were subsequently exposed to simulated solar wind and vacuum ultraviolet radiation to identify the effect of a simulated space environment on the as-weathered surfaces. Optical properties and adhesion testing were used to document the durability of the paints and coatings. The purpose of this paper is to present the results of the weathering testing and to summarize the durability of several thermal control paints, coatings, and appliques to weathering and postweathering environments.

Description: Results from a numerical and experimental study that illustrate the effects of laminate orthotropy on the buckling and failure response of compression-loaded composite cylindrical shells with a cutout are presented. The effects of orthotropy on the overall response of compression-loaded shells is described. In general, preliminary numerical results appear to accurately predict the buckling and failure characteristics of the shell considered herein. In particular, some of the shells exhibit stable post-local-buckling behavior accompanied by interlaminar material failures near the free edges of the cutout. In contrast another shell with a different laminate stacking sequence appears to exhibit catastrophic interlaminar material failure at the onset of local buckling near the cutout and this behavior correlates well with corresponding experimental results.

Description: Variation in constraint through the thickness of a specimen effects the cyclic crack-tip-opening displacement (DELTA CTOD). DELTA CTOD is a valuable measure of crack growth behavior, indicating closure development, constraint variations and load history effects. Fatigue loading with a continual load reduction was used to simulate the load history associated with fatigue crack growth threshold measurements. The constraint effect on the estimated DELTA CTOD is studied by carrying out three-dimensional elastic-plastic finite element simulations. The analysis involves numerical simulation of different standard fatigue threshold test schemes to determine how each test scheme affects DELTA CTOD. The American Society for Testing and Materials (ASTM) prescribes standard load reduction procedures for threshold testing using either the constant stress ratio (R) or constant maximum stress intensity (K(sub max)) methods. Different specimen types defined in the standard, namely the compact tension, C(T), and middle cracked tension, M(T), specimens were used in this simulation. The threshold simulations were conducted with different initial K(sub max) values to study its effect on estimated DELTA CTOD. During each simulation, the DELTA CTOD was estimated at every load increment during the load reduction procedure. Previous numerical simulation results indicate that the constant R load reduction method generates a plastic wake resulting in remote crack closure during unloading. Upon reloading, this remote contact location was observed to remain in contact well after the crack tip was fully open. The final region to open is located at the point at which the load reduction was initiated and at the free surface of the specimen. However, simulations carried out using the constant Kmax load reduction procedure did not indicate remote crack closure. Previous analysis results using various starting K(sub max) values and different load reduction rates have indicated DELTA CTOD is independent of specimen size. A study of the effect of specimen thickness and geometry on the measured DELTA CTOD for various load reduction procedures and its implication in the estimation of fatigue crack growth threshold values is discussed.

Description: The influence of delamination on the progressing damage path and initial failure load in composite laminates is investigated. Results are presented from a numerical and an experimental study of center-notched tensile-loaded coupons. The numerical study includes two approaches. The first approach considers only intralaminar (fiber breakage and matrix cracking) damage modes in calculating the progression of the damage path. In the second approach, the model is extended to consider the effect of interlaminar (delamination) damage modes in addition to the intralaminar damage modes. The intralaminar damage is modeled using progressive damage analysis (PDA) methodology implemented with the VUMAT subroutine in the ABAQUS finite element code. The interlaminar damage mode has been simulated using cohesive elements in ABAQUS. In the experimental study, 2-3 specimens each of two different stacking sequences of center-notched laminates are tensile loaded. The numerical results from the two different modeling approaches are compared with each other and the experimentally observed results for both laminate types. The comparisons reveal that the second modeling approach, where the delamination damage mode is included together with the intralaminar damage modes, better simulates the experimentally observed damage modes and damage paths, which were characterized by splitting failures perpendicular to the notch tips in one or more layers. Additionally, the inclusion of the delamination mode resulted in a better prediction of the loads at which the failure took place, which were higher than those predicted by the first modeling approach which did not include delaminations.

Description: The NASA White Sands Test Facility (WSTF) routinely operates hypergolic propulsion systems. Some of the onsite activities include performing long duration studies on the operational life of these systems. A few of them have been in use for over twenty years. During this span of time contamination has built up in the propellant and some of the distribution infrastructure. This study investigated the nature of this contamination, the pathology of its generation, and developed a process for removal of the contamination that was cost efficient with minimal waste generation.

Description: This viewgraph document reviews the use of electric batteries in space applications. Batteries are high energy devices that are used to power hardware for space applications The applications include IVA (Intra-Vehicular Activity) and EVA (Extra-Vehicular Activity) use. High energy batteries pose hazards such as cell/battery venting leading to electrolyte (liquid or gas) leakage, high temperatures, fire and explosion (shrapnel). It reviews the process of certifying of Commercial batteries for space applications in view of the multi-national purchasing for the International Space Station. The documentation used in the certification is reviewed.

Description: Previous composite cryotank designs have relied on the use of conventional composite materials to reduce microcracking and permeability. However, revolutionary advances in nanotechnology derived materials may enable the production of ultra-lightweight cryotanks with significantly enhanced durability and damage tolerance, as well as reduced propellant permeability. Layered silicate nanocomposites are especially attractive in cryogenic storage tanks based on results that have been reported for epoxy nanocomposite systems. These materials often exhibit an order of magnitude reduction in gas permeability when compared to the base resin. In addition, polymer-silicate nanocomposites have been shown to yield improved dimensional stability, strength, and toughness. The enhancement in material performance of these systems occurs without property trade-offs which are often observed in conventionally filled polymer composites. Research efforts at NASA Glenn Research Center have led to the development of epoxy-clay nanocomposites with 70% lower hydrogen permeability than the base epoxy resin. Filament wound carbon fiber reinforced tanks made with this nanocomposite had a five-fold lower helium leak rate than the corresponding tanks made without clay. The pronounced reduction observed with the tank may be due to flow induced alignment of the clay layers during processing. Additionally, the nanocomposites showed CTE reductions of up to 30%, as well as a 100% increase in toughness.

Description: S-Glass/epoxy [0/90/plus or minus theta]s for theta =45 deg., 60 deg., and 75 deg. laminated fiber-reinforced composite stiffened plate was simulated to investigated for damage and fracture progression under uniform pressure. An integrated computer code was augmented for the simulation of the damage initiation, growth, accumulation, and propagation to fracture and to structural collapse. Results show in detail the damage progression sequence and structural fracture resistance during different degradation stages. Damage through the thickness of the laminate initiated first at [0/90/plus or minus 45]s at 15.168 MPa (2200 psi), followed by [0/90/plus or minus 60]s at 16.96 MPa (2460 psi) and finally by [0/90/plus or minus 75]s at 19.3 MPa (2800 psi). After damage initiation happened the cracks propagate rapidly to structural fracture.

Description: White Sands Test Facility (WSTF) was requested to perform ambient temperature hydrostatic pressurization testing of a Space Transportation System (STS) 40-in. Kevlar Composite Overwrapped Pressure Vessel (COPV). The 40-in. vessel was of the same design and approximate age as the STS Main Propulsion System (MPS) and Orbiter Maneuvering System (OMS) vessels. The NASA Engineering Safety Center (NESC) assembled a team of experts and conducted an assessment that involved a review of national Kevlar COPY data. During the review, the STS COPVs were found to be beyond their original certification of ten years. The team observed that the likelihood of STS COPV Stress rupture, a catastrophic burst before leak failure mode, was greater than previously believed. Consequently, a detailed assessment of remaining stress rupture life became necessary. Prior to STS-114, a certification deviation was written for two flights of OV-103 (Discovery) and OV-104 (Atlantis) per rationale that was based on an extensive review of the Lawrence Livermore National Laboratories, COPV data, and revisions to the STS COPV stress levels. In order to obtain flight rationale to extend the certification deviation through the end of the Program, the Orbiter Project Office has directed an interagency COPV team to conduct further testing and analysis to investigate conservatism in the stress rupture model and evaluate material age degradation. Additional analysis of stress rupture life requires understanding the fiber stresses including stress that occurs due to thru-wall composite compression in COPV components. Data must be obtained at both zero gauge pressure (pre-stress) and at the component operating pressure so that this phenomenon can be properly evaluated. The zero gauge pressure stresses are predominantly a result of the autofrettage process used during vessel manufacture. Determining these pre-stresses and the constitutive behavior of the overwrap at pressure will provide necessary information to better predict the remaining life of the STS COPVs. The primary test objective is obtaining data to verify the hypothesis of a radially oriented thru-thickness stress-riser in the COPV composite whose magnitude is a function of the applied pressure and the load history. The anticipated load dependent response follows from the constitutive behavior of the composite overwrap so data to quantify its nonlinear and time dependent response will be sought. The objective of the Fiber Braggs Gratings (FBGs) were to advance the state-of-the-art by developing techniques using FBG sensors that are capable of assessing stress-rupture degradation in Kevlar COPVs in a health monitoring mode (1). Moreover, they sought to answer questions of how embedded sensors affect overall integrity of the structure. And lastly, they sought to provide an important link in the overall stress rupture study that will help close the loop on the COPV fabrication process. NDE inspection methods will be used from start to finish and FBG will be an integral link within the overall chain.

Description: A series of polyimide resins with low-melt viscosities in the range of 10-30 poise and high glass transition temperatures (Tg s) of 330-370 C were developed for resin transfer molding (RTM) applications. These polyimide resins were formulated from 2,3,3 ,4 -biphenyltetracarboxylic dianhydride (a-BPDA) with 4-phenylethynylphthalic anhydride endcaps along with either 3,4 - oxyaniline (3,4 -ODA), 3,4 -methylenedianiline, (3,4 -MDA) or 3,3 -methylenedianiline (3,3 -MDA). These polyimides had pot lives of 30-60 minutes at 260-280 C, enabling the successful fabrication of T650-35 carbon fiber reinforced composites via RTM process. The viscosity profiles of the polyimide resins and the mechanical properties of the polyimide carbon fiber composites will be discussed.