ALBERT

Description: The power point presentation includes an overview of the Hubble Space Telescope, pointing control hardware peculiarities, the two-gyro science (TGS) control system, TGS modifications to preserved hardware lifetime, chasing-down disturbance torques less than or equal to 0.002 Nm, inertia tensor optimization, and a summary with lessons learned.

Description: This Interim Standard establishes requirements for evaluation, testing, and selection of materials that are intended for use in space vehicles, associated Ground Support Equipment (GSE), and facilities used during assembly, test, and flight operations. Included are requirements, criteria, and test methods for evaluating the flammability, offgassing, and compatibility of materials.

Description: A method is provided for controlling operations in a video guidance sensor system wherein images of laser output signals transmitted by the system and returned from a target are captured and processed by the system to produce data used in tracking of the target. Six modes of operation are provided as follows: (i) a reset mode; (ii) a diagnostic mode; (iii) a standby mode; (iv) an acquisition mode; (v) a tracking mode; and (vi) a spot mode wherein captured images of returned laser signals are processed to produce data for all spots found in the image. The method provides for automatic transition to the standby mode from the reset mode after integrity checks are performed and from the diagnostic mode to the reset mode after diagnostic operations are carried out. Further, acceptance of reset and diagnostic commands is permitted only when the system is in the standby mode. The method also provides for automatic transition from the acquisition mode to the tracking mode when an acceptable target is found.

Description: The International Space Station (ISS) Program has many lessons to offer for the future of space exploration. Among these lessons of the ISS Program, three stand out as instrumental for the next generation of explorers. These include: 1) resourcefulness and the value of a strong international partnership; 2) flexibility as illustrated by the evolution of the ISS Program and 3) designing with dissimilar redundancy and simplicity of sparing. These lessons graphically demonstrate that the ISS Program can serve as a test bed for future programs. As the ISS Program builds upon the strong foundation of previous space programs, it can provide insight into the prospects for continued growth and cooperation in space exploration. As the capacity for spacefaring increases worldwide and as more nations invest in space exploration and space sector development, the potential for advancement in space exploration is unlimited. By building on its engineering and research achievements and international cooperation, the ISS Program is inspiring tomorrow s explorers today.

Description: An androgynous mating system for mating two exoatmospheric space modules comprising a first mating assembly capable of mating with a second mating assembly; a second mating assembly structurally identical to said first mating assembly, said first mating assembly comprising; a load ring; a plurality of load cell subassemblies; a plurality of actuators; a base ring; a tunnel; a closed loop control system; one or more electromagnets; and one or more striker plates, wherein said one or more electomagnets on said second mating assembly are capable of mating with said one or more striker plates on said first mating assembly, and wherein said one or more striker plates is comprised of a plate of predetermined shape and a 5-DOF mechanism capable of maintaining predetermined contact requirements during said mating of said one or more electromagnets and said one or more striker plates.

Description: Large space systems are required for a range of operational, commercial and scientific missions objectives however, current launch vehicle capacities substantially limit the size of space systems (on-orbit or planetary). Assembly and Deployment is the process of constructing a spacecraft or system from modules which may in turn have been constructed from sub-modules in a hierarchical fashion. In-situ assembly of space exploration vehicles and systems will require a broad range of operational capabilities, including: Component transfer and storage, fluid handling, construction and assembly, test and verification. Efficient execution of these functions will require supporting infrastructure, that can: Receive, store and protect (materials, components, etc.); hold and secure; position, align and control; deploy; connect/disconnect; construct; join; assemble/disassemble; dock/undock; and mate/demate.

Description: In the aerospace field spacecraft components are held together by separation systems until a specific time when they must be separated or deployed. Customarily a threaded joining bolt engages one of the components to be joined, and a threaded nut is placed on that bolt against the other component so they can be drawn together by a releasable locking assembly. The releasable locking assembly herein includes a plunger having one end coupled to one end of a plunger bolt. The other end is flanged to abut and compress a coil spring when the plunger is advanced toward the interface plane between the two components. When the plunger is so advanced toward the interface plane, the end of the plunger bolt can be connected to the joining bolt. Thus during retraction the joining bolt is drawn to one side of the interface plane by the force of the expanding spring.

Description: A key objective of NASA's Vision for Space Exploration is to revisit the lunar surface. Such an ambitious goal requires the development of a new human-rated spacecraft, the Orion Crew Exploration Vehicle (CEV), to ferry crews to low earth orbit and to the moon. The successful conclusion of both types of missions will require a thermal protection system (TPS) capable of protecting the vehicle and crew from the extreme heat of atmospheric reentry. As a part of the TPS development, various materials are being tested in arcjet tunnels; however, the combined lunar return aerothermal environment of high heat flux, shear stress, and surface pressure cannot be duplicated using only existing ground test facilities. To ensure full TPS qualification, a flight test program using sub-scale Orion capsules has been proposed to test TPS materials and heat shield construction techniques under the most stressing combination of lunar return aerothermal environments. Originally called Testing Of Reentry Capsule Heat Shield, or TORCH, but later renamed LEX, for Lunar Reentry Experiment, the proposed flight test program is presented along with the driving requirements and descriptions of the vehicle and the TPS instrumentation suite slated to conduct in-flight measurements.

Description: Concept studies for deep space missions are typically time-consuming and costly, given the variety of missions and uniqueness of each design. Yet, in an increasingly cost-constrained environment, it is critical to identify the most scientifically valuable and cost-effective designs early in the design process. Modeling is an integral part in helping to identify the most desirable design option. While some spacecraft design models currently exist for Earth-orbiting spacecraft, there has been less success with deep space missions. Instead, these missions require a modified design and modeling approach to enable the same construction of a comprehensive, yet credible, mission tradespace. This paper presents an approach for efficiently constructing such a mission tradespace. In addition to a proposed design and modeling approach, three case study missions are presented including a solar orbiter, a Europa orbiter, and a near-Earth asteroid (NEA) sample return mission.

Description: NASA is seeking to embark on a new set of human and robotic exploration missions back to the Moon, to Mars, and destinations beyond. Key strategic technical challenges will need to be addressed to realize this new vision for space exploration, including improvements in safety and reliability to improve robustness of space operations. Under sponsorship by NASA's Exploration Systems Mission, the Jet Propulsion Laboratory (JPL), together with its partners in government (NASA Johnson Space Center) and industry (Boeing, Vacco Industries, Ashwin-Ushas Inc.) is developing an ultra-low mass (〈3.0 kg) free-flying micro-inspector spacecraft in an effort to enhance safety and reduce risk in future human and exploration missions. The micro-inspector will provide remote vehicle inspections to ensure safety and reliability, or to provide monitoring of in-space assembly. The micro-inspector spacecraft represents an inherently modular system addition that can improve safety and support multiple host vehicles in multiple applications. On human missions, it may help extend the reach of human explorers, decreasing human EVA time to reduce mission cost and risk. The micro-inspector development is the continuation of an effort begun under NASA's Office of Aerospace Technology Enabling Concepts and Technology (ECT) program. The micro-inspector uses miniaturized celestial sensors; relies on a combination of solar power and batteries (allowing for unlimited operation in the sun and up to 4 hours in the shade); utilizes a low-pressure, low-leakage liquid butane propellant system for added safety; and includes multi-functional structure for high system-level integration and miniaturization. Versions of this system to be designed and developed under the H&RT program will include additional capabilities for on-board, vision-based navigation, spacecraft inspection, and collision avoidance, and will be demonstrated in a ground-based, space-related environment. These features make the micro-inspector design unique in its ability to serve crewed as well as robotic spacecraft, well beyond Earth-orbit and into arenas such as robotic missions, where human teleoperation capability is not locally available.

Description: A prototype system for monitoring spacecraft operations and control, including an alert system, is highlighted.

Description: This paper describes the attitude controller for the atmospheric entry of the Mars Science Laboratory (MSL). The controller will command 8 RCS thrusters to control the 3- axis attitude of the entry capsule. The Entry Controller is formulated as three independent channels in the control frame, which is nominally aligned with the stability frame. Each channel has a feedfoward and a feedback path. The feedforward path enables fast response to large bank commands. The feedback path stabilizes the vehicle angle of attack and sideslip around its trim position, and tracks bank commands. The feedback path has a PD/D control structure with deadbands that minimizes fuel usage. The performance of this design is demonstrated via computer simulations.

Description: The Mars Reconnaissance Orbiter is the most complex spacecraft that has ever been sent to investigate the Red Planet. A major part of what makes this mission so complex is the suite of instruments that were selected. The instruments on MRO vary from a simple imaging system, not much larger than a pocket knife to the largest camera ever flown to another planet. Not only does the size of the instruments vary, so do the scientific investigations associated with each instrument. In order to ensure that this payload suite would be able to satisfy all of its science objectives, a major effort was put forth by the MRO Project to ensure these instruments were well calibrated prior to the start of the Primary Science Phase. The in-flight calibration plan for MRO proved to be quite challenging, given the often conflicting requirements due to the varying capability of each of the instruments and the desire to constrain the workload on the Mission Operations personnel. The quality of data returned by MRO since the start of the Primary Science Phase is a tribute to the effort that was put forth to characterize the in-flight performance of the instruments. This paper will describe the challenges associated with the planning and implementation of the various calibration events on MRO, and will exhibit some of the results from those calibrations.

Description: In order for solar sail propulsion technologies to be considered as a viable option for a wide range of near term practical missions a predictable, stable, reliable, manufactureable, scaleable, and cost effective system must be developed and tested first on earth and then on orbit. The design and development of a Scaleable Square Solar Sail System (S^4) is well underway a t AEC-Able Engineering Co. Inc., and the design and production of the Solar Sails for this system is being carried out by SRS Technologies. In April and May of 2004 a single quadrant 10-meter system was tested at NASA LARC's vacuum chamber and a four quadrant 20-meter system has been designed and built for deployment and testing in the Spring of 2005 at NASA/Glenn Research Center's Plumb Brook Facility. SRS has developed an effective and efficient design for triangular sail quadrants that are supported are three points and provide a flat reflective surface with a high fill factor. This sail design is robust enough for deployments in a one atmosphere, one gravity environment and incorporates several advanced features including adhesiveless seaming of membrane strips, compliant edge borders to allow for film membrane cord strain mismatch without causing wrinkling and low mass (3% of total sail mass) ripstop. This paper will outline the sail design and fabrication process, the lessons learned and the resulting mature production, packaging and deployment processes that have been developed. It will also highlight the scalability of the equipment and processes that were developed to fabricate and package the sails. Based on recent experience, SRS is confidant that flight worthy solar sails in the 40-120-meter size range with areal density in the 4-5g/sq m (sail minus structure) range can be produced with existing technology. Additional film production research will lead to further reductions in film thickness to less than 1 micron enabling production of sails with areal densities as low as 20 g/sq m using the current design resulting in a system areal density of as low as 5.3g/sq m. These areal densities are low enough to allow nearly all of the Solar Sail missions that have been proposed by the scientific community and the fundamental technology required to produce these sails has been demonstrated on the ground test sails that have recently been built. These demonstrations have shown that the technology is mature enough to build sails needed to support critical science missions. Solar Sails will be an enabling technology for NASA's Vision for Space Exploration by allowing communication satellite orbits that can maintain continuous communication with the polar regions of the Moon and Mars and to support solar weather monitoring to provide early warning of solar flares and storms that could threaten the safety of astronauts and other spacecraft.

Description: Forward Attached Inflatable Decelerators, more commonly known as inflatable aeroshells, provide an effective, cost efficient means of decelerating spacecrafts by using atmospheric drag for aerocapture or planetary entry instead of conventional liquid propulsion deceleration systems. Entry into planetary atmospheres results in significant heating and aerodynamic pressures which stress aeroshell systems to their useful limits. Incorporation of lightweight inflatable decelerator surfaces with increased surface-area footprints provides the opportunity to reduce heat flux and induced temperatures, while increasing the payload mass fraction. Furthermore, inflatable aeroshell decelerators provide the needed deceleration at considerably higher altitudes and Mach numbers when compared with conventional rigid aeroshell entry systems. Inflatable aeroshells also provide for stowage in a compact space, with subsequent deployment of a large-area, lightweight heatshield to survive entry heating. Use of a deployable heatshield decelerator enables an increase in the spacecraft payload mass fraction and may eliminate the need for a spacecraft backshell.

Description: This paper describes the structural dynamic tests conducted in-vacuum on the Scalable Square Solar Sail (S(sup 4)) System 20-meter test article developed by ATK Space Systems as part of a ground demonstrator system development program funded by NASA's In-Space Propulsion program1-3. These tests were conducted for the purpose of validating analytical models that would be required by a flight test program to predict in space performance4. Specific tests included modal vibration tests on the solar sail system in a 1 Torr vacuum environment using various excitation locations and techniques including magnetic excitation at the sail quadrant corners, piezoelectric stack actuation at the mast roots, spreader bar excitation at the mast tips, and bi-morph piezoelectric patch actuation on the sail cords. The excitation methods were evaluated for their suitability to in-vacuum ground testing and their traceability to the development of on-orbit flight test techniques. The solar sail masts were also tested in ambient atmospheric conditions and these results are also discussed.

Description: Solar sail tip-mounted, lightweight pulsed plasma thrusters (PPTs) are proposed for a secondary (or backup) attitude control system (ACS) of a 160-m, 450-kg solar sail spacecraft of the Solar Polar Imager (SPI) mission. A propellantless primary ACS of the SPI sailcraft employs trim control masses running along mast lanyards for pitch/yaw control together with roll stabilizer bars at the mast tips for quadrant tilt (roll) control. The robustness of such a propellantless primary ACS would be further enhanced by a secondary ACS utilizing tip-mounted, lightweight PPTs. The microPPT-based ACS is intended mainly for attitude recovery maneuvers from various off-nominal conditions that cannot be reliably handled by the propellantless primary ACS. However, it can also be employed for: i) the checkout or standby mode prior to and during sail deployment, ii) the post-deployment transition mode (prior to the propellantless primary ACS mode operation), iii) the solar sailing cruise mode of a trimmed sailcraft, and iv) the spin-stabilized, sun-pointing, safe mode. Although a conventional bus ACS is required for the SPI mission as the sail is jettisoned at the start of its science mission phase, the microPPT-based ACS option promises greater redundancy and robustness for the SPI mission. For other sailing missions, where the sail is never jettisoned, this secondary ACS provides a lower-cost, lower-mass propulsion for deployment control and greater redundancy than any traditional reaction-jet control system. This paper presents an overview nf the state--of-the--art microPPT technology, the design requirements of microPPTs for solar sail attitude control, and the preliminary ACS design and simulation results.

Description: On September 27, 2007, a Delta II rocket carrying the Dawn spacecraft lifted off from Kennedy Space Center. Part of NASAs Discovery program, the $370 million Dawn mission began its three-billion-mile voyage to the asteroid belt to study the asteroid Vesta and Ceres, a dwarf planet. The spacecraft is scheduled to reach Vesta in 2011. After spending nine months measuring the composition, shape, and topography of that body, it will travel a billion miles to carry out a similar analysis of Ceres in 2015. The Important Lessons: The demands of Dawn and other challenging missions have taught some important lessons for successful program and project management. These are the main ones: a) Program management, particularly of uncoupled and loosely coupled projects, should be more about enabling than controlling. You're working with motivated, high-performing teams and institutions with a track record of quality and success. Emphasize commander's intent over rudder control; let them know where you want to go and when you want to be there, then let them figure out how to get there. b) Open and honest discussion of issues is essential. People fill the void of the unknown with their worst fears. Get folks around the table and have open, honest, and frank dialogue. I've seldom seen this fail to get to the root of issues. c) You have to earn your seat at the table, proving that you are competent, trustworthy, and dedicated to the success of the mission. d) Know when to fold 'em. Your pride can get rolled up in making a milestone or launch date, but you have to make a judgment based on the realities of the situation and not wear down the team trying to meet an increasingly impossible deadline. e) The NASA governance model that gives a voice to the concerns of engineers and safety experts works-trust it and use it.

Description: The International Space Station (ISS) is a great international, technological, and political achievement. It is the latest step in humankind's quest to explore and live in space. The research done on the ISS may advance our knowledge in various areas of science, enable us to improve life on this planet, and give us the experience and increased understanding that can eventually equip us to journey to other worlds. As a result of the Station s complexity, few understand its configuration, its design and component systems, or the complex operations required in its construction and operation. This book provides high-level insight into the ISS. The ISS is in orbit today, operating with a crew of three. Its assembly will continue through 2010. As the ISS grows, its capabilities will increase, thus requiring a larger crew. Currently, 16 countries are involved in this venture. This CD-ROM includes multimedia files and animations.

Description: Solar sails employ a unique form of propulsion, gaining momentum from incident and reflected photons. However, the momentum transferred by an individual photon is extremely small. Consequently, a solar sail must have an extremely large surface area and also be extremely light. The flexibility of the sail then must be considered when designing or evaluating control laws. In this paper, solar sail flexibility and its influence on control effectiveness is considered using idealized two-dimensional models to represent physical phenomena rather than a specific design. Differential equations of motion are derived for a distributed parameter model of a flexible solar sail idealized as a rotating central hub with two opposing flexible booms. This idealization is appropriate for solar sail designs in which the vibrational modes of the sail and supporting booms move together allowing the sail mass to be distributed along the booms in the idealized model. A reduced analytical model of the flexible response is considered. Linear feedback torque control is applied at the central hub. Two translational disturbances and a torque disturbance also act at the central hub representing the equivalent effect of deflecting sail shape about a reference line. Transient simulations explore different control designs and their effectiveness for controlling orientation, for reducing flexible motion and for disturbance rejection. A second model also is developed as a two-dimensional "pathfinder" model to calculate the effect of solar sail shape on the resultant thrust, in-plane force and torque at the hub. The analysis is then extended to larger models using the finite element method. The finite element modeling approach is verified by comparing results from a two-dimensional finite element model with those from the analytical model. The utility of the finite element modeling approach for this application is then illustrated through examples based on a full finite element model.

Description: Under the direction of the NASA In-Space Propulsion Technology Office, the team of L Garde, NASA Jet Propulsion Laboratory, Ball Aerospace, and NASA Langley Research Center has been developing a scalable solar sail configuration to address NASA's future space propulsion needs. Prior to a flight experiment of a full-scale solar sail, a comprehensive phased test plan is currently being implemented to advance the technology readiness level of the solar sail design. These tests consist of solar sail component, subsystem, and sub-scale system ground tests that simulate the vacuum and thermal conditions of the space environment. Recently, two solar sail test articles, a 7.4-m beam assembly subsystem test article and a 10-m four-quadrant solar sail system test article, were tested in vacuum conditions with a gravity-offload system to mitigate the effects of gravity. This paper presents the structural analyses simulating the ground tests and the correlation of the analyses with the test results. For programmatic risk reduction, a two-prong analysis approach was undertaken in which two separate teams independently developed computational models of the solar sail test articles using the finite element analysis software packages: NEiNastran and ABAQUS. This paper compares the pre-test and post-test analysis predictions from both software packages with the test data including load-deflection curves from static load tests, and vibration frequencies and mode shapes from vibration tests. The analysis predictions were in reasonable agreement with the test data. Factors that precluded better correlation of the analyses and the tests were uncertainties in the material properties, test conditions, and modeling assumptions used in the analyses.

Description: On a spring day in 1996, at their research center in the Maryland countryside, representatives from the Johns Hopkins University Applied Physics Laboratory (APL) presented Administrator Daniel S. Goldin of the National Aeronautics and Space Administration (NASA) with a check for $3.6 million. 1 Two and a half years earlier, APL officials had agreed to develop a spacecraft capable of conducting an asteroid rendezvous and to do so for slightly more than $122 million. This was a remarkably low sum for a spacecraft due to conduct a planetaryclass mission. By contrast, the Mars Observer spacecraft launched in 1992 for an orbital rendezvous with the red planet had cost $479 million to develop, while the upcoming Cassini mission to Saturn required a spacecraft whose total cost was approaching $1.4 billion. In an Agency accustomed to cost overruns on major missions, the promise to build a planetary-class spacecraft for about $100 million seemed excessively optimistic.

Description: The last decade has seen a significant increase in the number and the capabilities of remote sensing satellites launched by the international community. A relatively new approach has been the launching of satellites into heterogeneous constellations. Constellations provide the scientists a capability to acquire science data, not only from specific instruments on a single satellite, but also from instruments on other satellites that fly in the same orbit. Initial results from the A-Train (especially following the CALIPSO/CloudSat launch) attest to the tremendous scientific value of constellation flying. This paper provides a history of the constellations (particularly the A-Train) and how the A-Train mission design was driven by science requirements. The A-Train has presented operational challenges which had not previously been encountered. Operations planning had to address not only how the satellites of each constellation operate safely together, but also how the two constellations fly in the same orbits without interfering with each other when commands are uplinked or data are downlinked to their respective ground stations. This paper discusses the benefits of joining an on-orbit constellation. When compared to a single, large satellite, a constellation infrastructure offers more than just the opportunities for coincidental science observations. For example, constellations reduce risks by distributing observing instruments among numerous satellites; in contrast, a failed launch or a system failure in a single satellite would lead to loss of all observations. Constellations allow for more focused, less complex satellites. Constellations distribute the development, testing, and operations costs among various agencies and organizations for example, the Morning and Afternoon Constellations involve several agencies within the U.S. and in other countries. Lastly, this paper addresses the need to plan for the long-term evolution of a constellation. Agencies need to have a replenishment strategy as some satellites age and eventually leave the constellation. This will ensure overlap of observations, thus providing continuous, calibrated science data over a much longer time period. Thoughts on the evolution of the A-Train will also be presented.

Description: As a space faring nation, we are at a critical juncture in the evolution of space exploration. NASA has announced its Vision for Space Exploration, a vision of returning humans to the Moon, sending robots and eventually humans to Mars, and exploring the outer solar system via automated spacecraft. However, mission concepts have become increasingly complex, with the potential to yield a wealth of scientific knowledge. Meanwhile, there are significant resource challenges to be met. Launch costs remain a barrier to routine space flight; the ever-changing fiscal and political environments can wreak havoc on mission planning; and technologies are constantly improving, and systems that were state of the art when a program began can quickly become outmoded before a mission is even launched. This Conference Publication describes the workshop and featured presentations by world-class experts presenting leading-edge technologies and applications in the areas of power and propulsion; communications; automation, robotics, computing, and intelligent systems; and transformational techniques for space activities. Workshops such as this one provide an excellent medium for capturing the broadest possible array of insights and expertise, learning from researchers in universities, national laboratories, NASA field Centers, and industry to help better our future in space.

Description: Future NASA Exploration goals are difficult to meet using current launch vehicle implementations and techniques. We introduce a concept of On-Orbit Staging (OOS) using multiple launches into a Low Earth orbit (LEO) staging area to increase payload mass and reduce overall cost for exploration initiative missions. This concept is a forward-looking implementation of ideas put forth by Oberth and Von Braun to address the total mission design. Applying staging throughout the mission and utilizing technological advances in propulsion efficiency and architecture enable us to show that exploration goals can be met in the next decade. As part of this architecture, we assume the readiness of automated rendezvous, docking, and assembly technology.

Description: No abstract available

Description: No abstract available

Description: A Space Shuttle Columbia main engine controller 14 AWG wire short circuited during the launch of STS-93. Post-flight examination divulged that the wire had electrically arced against the head of a nearby bolt. More extensive inspection revealed additional damage to the subject wire, and to other wires as well from the mid-body of Columbia. The shorted wire was to have been constructed from nickel-plated copper conductors surrounded by the polyimide insulation Kapton, top-coated with an aromatic polyimide resin. The wires were analyzed via scanning electron microscope (SEM), energy dispersive X-Ray spectroscopy (EDX), and electron spectroscopy for chemical analysis (ESCA); differential scanning calorimetry (DSC) and thermal gravimetric analysis (TGA) were performed on the polyimide. Exemplar testing under laboratory conditions was performed to replicate the mechanical damage characteristics evident on the failed wires. The exemplar testing included a step test, where, as the name implies, a person stepped on a simulated wire bundle that rested upon a bolt head. Likewise, a shear test that forced a bolt head and a torque tip against a wire was performed to attempt to damage the insulation and conductor. Additionally, a vibration test was performed to determine if a wire bundle would abrade when vibrated against the head of a bolt. Also, an abrasion test was undertaken to determine if the polyimide of the wire could be damaged by rubbing against convolex helical tubing. Finally, an impact test was performed to ascertain if the use of the tubing would protect the wire from the strike of a foreign object.

Description: Most Low Earth Orbit (LEO) debris lies in a limited number of inclination "bands" associated with launch latitudes, or with specific useful orbit inclinations (such as polar orbits). Such narrow inclination bands generally have a uniform spread over all possible Right Ascensions of Ascending Node (RAANs), creating a different orbit plane for nearly every piece of debris. This complicates concept of rendezvous and capture for debris removal. However, a low-orbiting satellite will always phase in RAAN faster than debris objects in higher orbits at the same inclination, potentially solving the problem. Such a base can serve as a single space-based launch facility (a "mother ship") that can tend and then send tiny individual catcher devices for each debris object, as the facility drifts into the same RAAN as the higher object. This presentation will highlight characteristic system requirements of such an architecture, including structural and navigation requirements, power, mass and dV budgets for both the mother ship and the mass-produced common catcher devices that would clean out selected inclination bands. The altitude and inclination regime over which a band is to be cleared, the size distribution of the debris, and the inclusion of additional mission priorities all affect the sizing of the system. It is demonstrated that major LEO hazardous debris reductions can be realized in each band with a single LEO launch of a single mother ship, with simple attached catchers of total mass less than typical commercial LEO launch capability.

Description: Integrated vehicle testing will be critical to ensuring proper vehicle integration of the Ares I crew launch vehicle and Ares V cargo launch vehicle. The Ares Projects, based at Marshall Space Flight Center in Alabama, created the Flight and Integrated Test Office (FITO) as a separate team to ensure that testing is an integral part of the vehicle development process. As its name indicates, FITO is responsible for managing flight testing for the Ares vehicles. FITO personnel are well on the way toward assembling and flying the first flight test vehicle of Ares I, the Ares I-X. This suborbital development flight will evaluate the performance of Ares I from liftoff to first stage separation, testing flight control algorithms, vehicle roll control, separation and recovery systems, and ground operations. Ares I-X is now scheduled to fly in summer 2009. The follow-on flight, Ares I-Y, will test a full five-segment first stage booster and will include cryogenic propellants in the upper stage, an upper stage engine simulator, and an active launch abort system. The following flight, Orion 1, will be the first flight of an active upper stage and upper stage engine, as well as the first uncrewed flight of an Orion spacecraft into orbit. The Ares Projects are using an incremental buildup of flight capabilities prior to the first operational crewed flight of Ares I and the Orion crew exploration vehicle in 2015. In addition to flight testing, the FITO team will be responsible for conducting hardware, software, and ground vibration tests of the integrated launch vehicle. These efforts will include verifying hardware, software, and ground handling interfaces. Through flight and integrated testing, the Ares Projects will identify and mitigate risks early as the United States prepares to take its next giant leaps to the Moon and beyond.

Description: A Japanese led international team is developing a suborbital test of orbital-motion-limited (OML) bare wire anode current collection for application to electrodynamic tether (EDT) propulsion. The tether is a tape with a width of 25 mm, thickness of 0.05 mm, and is 300 m in length. This will be the first space test of OML theory. The mission will launch in the summer of 2009 using an S520 Sounding Rocket. During ascent, and above approx. 100 km in attitude, the tape tether will be deployed at a rate of approx. 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. The total amount of current collected will be used to assess the validity of OML theory. 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 EDTs for propulsion or power generation.

Description: The National Aeronautics and Space Administration (NASA) is developing new launch systems and preparing to retire the Space Shuttle by 2010, as directed in the United States (U.S.) Vision for Space Exploration. The Ares I Crew Launch Vehicle (CLV) and the Ares V heavy-lift Cargo Launch Vehicle (CaLV) systems will build upon proven, reliable hardware derived from the Apollo-Saturn and Space Shuttle programs to deliver safe, reliable, affordable space transportation solutions. This approach leverages existing aerospace talent and a unique infrastructure, as well as legacy knowledge gained from nearly 50 years' experience developing space hardware. Early next decade, the Ares I will launch the new Orion Crew Exploration Vehicle (CEV) to the International Space Station (ISS) or to low-Earth orbit for trips to the Moon and, ultimately, Mars. Late next decade, the Ares V's Earth Departure Stage will carry larger payloads such as the lunar lander into orbit, and the Crew Exploration Vehicle will dock with it for missions to the Moon, where astronauts will explore new territories and conduct science and technology experiments. Both Ares I and Ares V are being designed to support longer future trips to Mars. The Exploration Launch Projects Office is designing, developing, testing, and evaluating both launch vehicle systems in partnership with other NASA Centers, Government agencies, and industry contractors. This paper provides top-level information regarding the genesis and evolution of the baseline configuration for the Ares V heavy-lift system. It also discusses riskbased, management strategies, such as building on powerful hardware and promoting common features between the Ares I and Ares V systems to reduce technical, schedule, and cost risks, as well as development and operations costs. Finally, it summarizes several notable accomplishments since October 2005, when the Exploration Launch Projects effort officially kicked off, and looks ahead at work planned for 2007 and beyond.

Description: This paper describes the development and initial demonstration of a Portable Health Algorithms Test (PHALT) System that is being developed by researchers at the NASA Glenn Research Center (GRC). The PHALT System was conceived as a means of evolving the maturity and credibility of algorithms developed to assess the health of aerospace systems. Comprising an integrated hardware-software environment, the PHALT System allows systems health management algorithms to be developed in a graphical programming environment; to be tested and refined using system simulation or test data playback; and finally, to be evaluated in a real-time hardware-in-the-loop mode with a live test article. In this paper, PHALT System development is described through the presentation of a functional architecture, followed by the selection and integration of hardware and software. Also described is an initial real-time hardware-in-the-loop demonstration that used sensor data qualification algorithms to diagnose and isolate simulated sensor failures in a prototype Power Distribution Unit test-bed. Success of the initial demonstration is highlighted by the correct detection of all sensor failures and the absence of any real-time constraint violations.

Description: NASA's Global Precipitation Measurement (GPM) mission is an ongoing Goddard Space Flight Center (GSFC) project whose basic objective is to improve global precipitation measurements. It has been decided that the GPM spacecraft is to be a "design for demise" spacecraft. This requirement resulted in the need for a propellant tank that would also demise or ablate to an appropriate degree upon re-entry. This paper will describe GSFC-performed spacecraft and tankage demise analyses, vendor conceptual design studies, and vendor performed hydrazine compatibility and wettability tests performed on 6061 and 2219 aluminum alloys.

Description: A tether system for providing thrust to or power subsystems of an artificial satellite in a low earth orbit. The tether has three main sections, an insulated section connected to the satellite, a conducting section connected to the insulating section for drawing in and releasing electrons from the space plasma and a non-conducting section for providing a tension to the other sections of the tether. An oxygen resistant coating is applied to the bare wire of the conducting section as well as the insulated wires of the insulated section that prevents breakdown during tether operations in the space plasma. The insulated and bare wire sections also surround a high tensile flexible polymer core to prevent any debris from breaking the tether during use.

Description: The Momentum Exchange/Electrodynamic Reboost (MXER) tether facility is a transformational concept that significantly reduces the fuel requirements (and associated costs) in transferring payloads above low earth orbit (LEO). Facility reboost is accomplished without propellant by driving current against a voltage created by a conducting tether's interaction with the Earth's magnetic field (electrodynamic reboost). This system can be used for transferring a variety of payloads (scientific, cargo, and human space vehicles) to multiple destinations including geosynchronous transfer orbit, the Moon or Mars. MXER technology advancement requires development in two key areas: survivable, high tensile strength non-conducting tethers and reliable, lightweight payload catch/release mechanisms. Fundamental requirements associated with the MXER non-conducting strength tether and catch mechanism designs will be presented. Key requirements for the tether design include high specific-strength (tensile strength/material density), material survivability to the space environment (atomic oxygen and ultraviolet radiation), and structural survivability to micrometeoroid/orbital debris (MM/OD) impacts. The driving mechanism key,gequirements include low mass-to-capture-volume ratio, positional and velocity error tolerance, and operational reliability. Preliminary tether and catch mechanism design criteria are presented, which have been used as guidelines to "screen" and down-select initial concepts. Candidate tether materials and protective coatings are summarized along with their performance in simulated space environments (e.g., oxygen plasma, thermal cycling). A candidate catch mechanism design concept is presented along with examples of demonstration hardware.

Description: To protect spacecraft and their contents from excessive heat thermal protection system are essential. For such thermal protection, metal coatings, ceramic materials, ablative materials, and various matrix materials have all been tried, but none have been found entirely satisfactory. The basis for this thermal protection system is the fact that the heat required to melt a substance is 80 to 100 times larger than the heat required to raise its temperature one degree. This led to the use herein of solid-liquid phase change materials. Unlike conventional heat storage materials, when phase change materials reach the temperature at which they change phase they absorb large amounts of heat without getting hotter. By this invention, then, a coating composition is provided for application to substrates subjected to temperatures above 100 F. The coating composition includes a phase change material.

Description: Contents include the following: Introduction. Capability Breakdown Structure. Decelerator Functions. Candidate Solutions. Performance and Technology. Capability State-of-the-Art. Performance Needs. Candidate Configurations. Possible Technology Roadmaps. Capability Roadmaps.

Description: Capability Roadmap Team. Capability Description, Scope and Capability Breakdown Structure. Benefits of the HPLS. Roadmap Process and Approach. Current State-of-the-Art, Assumptions and Key Requirements. Top Level HPLS Roadmap. Capability Presentations by Leads. Mission Drivers Requirements. "AEDL" System Engineering. Communication & Navigation Systems. Hypersonic Systems. Super to Subsonic Decelerator Systems. Terminal Descent and Landing Systems. A Priori In-Situ Mars Observations. AEDL Analysis, Test and Validation Infrastructure. Capability Technical Challenges. Capability Connection Points to other Roadmaps/Crosswalks. Summary of Top Level Capability. Forward Work.

Description: Contents include the following: Capability Description. Some Initial Thoughts. Capability State-of-the-Art, Gaps and Requirements. Capability Roadmap. Candidate Technologies. Metrics.

Description: Contents include the following: Capability Description, Benefits, Current State-of-the-Art. Capability Requirements and Assumptions. Maturity Level - Capabilities. Maturity Level - Technologies. Metrics. Roadmap for Capability.

Description: Contents include the following: NASA capability roadmap activity. Advanced modeling, simulation, and analysis overview. Scientific modeling and simulation. Operations modeling. Multi-special sensing (UV-gamma). System integration. M and S Environments and Infrastructure.

Description: It has been over 35 years since NASA developed new human spaceflight capabilities. As NASA builds vehicles to once again venture beyond Earth's orbit, it has the advantage of a powerful legacy of seasoned professionals who have already been there. Apollo-era veterans are lending their knowledge and expertise to nearly every aspect of the new Ares I crew launch vehicle and the Ares V cargo launch vehicle, from management to design and manufacturing techniques. Through group discussions, personal interviews, and consultant relationships, these talented individuals are sharing their "lessons lived" to help a new generation of engineers repeat the successes and avoid some of the pitfalls of America's first journeys to the Moon. In addition to learning from resident and retired experts, Ares will draw on legacy facilities, tooling, and hardware like the J-2 engine from the Apollo era and the Reusable Solid Rocket Boosters from the Space Shuttle Program. NASA needs to re-learn the skills required to send astronauts to the Moon, Mars, and beyond. The new Ares team is training with the best and building on the work of their eminent predecessors. They are standing on the shoulders of giants to see a future that is bright with possibilities on the space frontier.

Description: The NUMIT 1-dimensional bulk charging model is used as a screening to ol for evaluating time-dependent bulk internal or deep dielectric) ch arging of dielectrics exposed to penetrating electron environments. T he code is modified to accept time dependent electron flux time serie s along satellite orbits for the electron environment inputs instead of using the static electron flux environment input originally used b y the code and widely adopted in bulk charging models. Application of the screening technique ts demonstrated for three cases of spacecraf t exposure within the Earth's radiation belts including a geostationa ry transfer orbit and an Earth-Moon transit trajectory for a range of orbit inclinations. Electric fields and charge densities are compute d for dielectric materials with varying electrical properties exposed to relativistic electron environments along the orbits. Our objectiv e is to demonstrate a preliminary application of the time-dependent e nvironments input to the NUMIT code for evaluating charging risks to exposed dielectrics used on spacecraft when exposed to the Earth's ra diation belts. The results demonstrate that the NUMIT electric field values in GTO orbits with multiple encounters with the Earth's radiat ion belts are consistent with previous studies of charging in GTO orb its and that potential threat conditions for electrostatic discharge exist on lunar transit trajectories depending on the electrical proper ties of the materials exposed to the radiation environment.

Description: This slide presentation reviews the uses of wires in the Aerospace industry. The vision is to minimize cables and connectors and increase functionality across the aerospace industry by providing reliable lower cost modular and higher performance alternatives to wired data connectivity to benefit the entire vehicle and program

Description: Upon observing an abnormal closure of the Space Shuttle s External Tank Doors (ETD), a dynamic model was created in MSC/ADAMS to conduct deflection analyses for assessing whether the Door Drive Mechanism (DDM) was subjected to excessive additional stress, and more importantly, to evaluate the magnitude of the induced step or gap with respect to shuttle s body tiles. To model the flexibility of the DDM, a lumped parameter approximation was used to capture the compliance of individual parts within the drive linkage. These stiffness approximations were then validated using FEA and iteratively updated in the model to converge on the actual distributed parameter equivalent stiffnesses. The goal of the analyses is to determine the deflections in the mechanism and whether or not the deflections are in the region of elastic or plastic deformation. Plastic deformation may affect proper closure of the ETD and would impact aero-heating during re-entry.

Description: Ground crew veterans at Kennedy Space Center still talk about what they call "the summer of hydrogen"-the long, frustrating months in 1990 when the shuttle fleet was grounded by an elusive hydrogen leak that foiled our efforts to fill the orbiter's external fuel tank. Columbia (STS-35) was on Launch Pad A for a scheduled May 30 launch when we discovered the hydrogen leak during - tanking. The external fuel tank is loaded through the orbiter. Liquid hydrogen flows through a 17-inch umbilical between the orbiter and the tank. During fueling, we purge the aft fuselage with gaseous nitrogen to reduce the risk of fire, and we have a leak-detection system in the mobile launch platform, which samples (via tygon tubing) the atmosphere in and around the vehicle, drawing it down to a mass spectrometer that analyzes its composition. When we progressed to the stage of tanking where liquid hydrogen flows through the vehicle, the concentration of hydrogen approached four percent-the limit above which it would be dangerously flammable. We had a leak. We did everything we could think of to find it, and the contractor who supplied the flight hardware was there every day, working alongside us. We did tanking tests, which involved instrumenting the suspected leak sources, and cryo-loaded the external tank to try to isolate precisely where the leak originated. We switched out umbilicals; we replaced the seals between the umbilical and the orbiter. We inspected the seals microscopically and found no flaws. We replaced the recirculation pumps, and we found and replaced a damaged teflon seal in a main propulsion system detent cover, which holds the prevalve-the main valve supplying hydrogen to Space Shuttle Main Engine 3 -in the open position. The seal passed leak tests at ambient temperature but leaked when cryogenic temperatures were applied. We added new leak sensors-up to twenty at a time and tried to be methodical in our placements to narrow down the possible sources of the problem. We even switched orbiters, sending Columbia back to the Vehicle Assembly Building and bringing out Atlantis, scheduled to fly as STS-38. Two shuttles on their mobile launchers passing in the night was a majestic sight, but not one you want to see if you're trying to get an orbiter launched. None of this told us where the leak was, or if we were dealing with more than one leak source.

Description: In response to the Vision for Space Exploration, the National Aeronautics and Space Administration (NASA) has defined a new space exploration architecture to return humans to the Moon and prepare for human exploration of Mars. One of the first new developments will be the Ares I Crew Launch Vehicle (CLV), which will carry the Orion Crew Exploration Vehicle (CEV), into Low Earth Orbit (LEO) to support International Space Station (ISS) missions and, later, support lunar missions. As part of Ares I development, NASA will perform a series of Ares I flight tests. The tests will provide data that will inform the engineering and design process and verify the flight hardware and software. The data gained from the flight tests will be used to certify the new Ares/Orion vehicle for human space flight. The primary objectives of this first flight test (Ares I-X) are the following: Demonstrate control of a dynamically similar integrated Ares CLV/Orion CEV using Ares CLV ascent control algorithms; Perform an in-flight separation/staging event between an Ares I-similar First Stage and a representative Upper Stage; Demonstrate assembly and recovery of a new Ares CLV-like First Stage element at Kennedy Space Center (KSC); Demonstrate First Stage separation sequencing, and quantify First Stage atmospheric entry dynamics and parachute performance; and Characterize the magnitude of the integrated vehicle roll torque throughout the First Stage (powered) flight. This paper will provide an overview of the Ares I-X flight test process and details of the individual flight tests.

Description: The NASA In-Space Propulsion Technology (ISPT) Projects Office has been sponsoring 2 separate, independent system design and development hardware demonstration activities during 2002-2005. ATK Space Systems of Goleta, CA was the prime contractor for one development team and L'Garde, Inc. of Tustin, CA was the prime contractor for the other development team. The goal of these activities was to advance the technology readiness level (TRL) of solar sail propulsion from 3 towards 6 by the year 2006. Component and subsystem fabrication and testing were completed successfully, including the ground deployment of 10-meter and 20-meter ground demonstration hardware systems under vacuum conditions. The deployment and structural testing of the 20-meter solar sail systems was conducted in the 30 meter diameter Space Power Facility thermal-vacuum chamber at NASA Glenn Plum Brook in April though August, 2005. This paper will present the results of the TRL assessment following the solar sail technology development activities associated with the design, development, analysis and testing of the 20-meter system ground demonstrators. Descriptions of the system designs for both the ATK and L'Garde systems will be presented. Changes, additions and evolution of the system designs will be highlighted. A description of the modeling and analyses activities performed by both teams, as well as testing conducted to raise the TRL of solar sail technology will be presented. A summary of the results of model correlation activities will be presented. Finally, technology gaps identified during the assessment and gap closure plans will be presented, along with "lessons learned", subsequent planning activities and validation flight opportunities for solar sail propulsion technology.

Description: The Heliostorm (also referred to as Geostorm) mission has been regarded as the best choice for the first application of solar sail technology. The objective of Heliostorm is to obtain data from an orbit station slightly displaced from the ecliptic at or nearer to the Sun than 0.98 AU, which places it twice as close to the sun as Earth's natural L1 point at 0.993 AU. The maintenance of such an orbit location would require prohibitive amounts of propellants using chemical or electric propulsion systems; however, a solar sailcraft is ideally suited for this purpose because it relies solely on the propulsive force from photons for orbit maintenance. Heliostorm has been the subject of several mission studies over the past decade, with the most complete study conducted in 1999 in conjunction with a proposed New Millennium Program (NMP) Space Technology 5 (ST-5) flight opportunity. Recently, over a two and one-half year period dating from 2003 through 2005, NASA's In-Space Propulsion Technology Program (ISTP) matured solar sail technology from laboratory components to full systems, demonstrated in as relevant a space environment as could feasibly be simulated on the ground. Work under this program has yielded promising results for enhanced Heliostorm mission performance. This enhanced performance is achievable principally through reductions in the sail areal density. These reductions are realized through the use of lower linear mass density booms, a thinner sail membrane, and increased sail area. Advancements in sailcraft vehicle system design also offer potential mass reductions and hence improved performance. This paper will present the preliminary results of an updated Heliostorm mission design study including the enhancements incorporated during the design, development, analysis and testing of the system ground demonstrator.