ALBERT

Description: The Laser Interferometer Space Antenna (LISA) mission. a space based gravitational wave detector. uses laser metrology to measure distance fluctuations between proof masses aboard three spacecraft. LISA is unique from a mission design perspective in that the three spacecraft and their associated operations form one distributed science instrument. unlike more conventional missions where an instrument is a component of an individual spacecraft. The design of the LISA spacecraft is also tightly coupled to the design and requirements of the scientific payload; for this reason it is often referred to as a "sciencecraft." Here we describe some of the unique features of the LISA spacecraft design that help create the quiet environment necessary for gravitational wave observations.

Description: No abstract available

Description: Completed thermal-mechanical and mechanical load testing: a) 6 re-entry heating tests (3 with loading to 50% DLL), 3 hypersonic cruise tests with loading to 50% DLL and 4 high-temperature modal survey tests. b) 9 tests to 100% DLL. High-temperature modal survey results were inconclusive due to exceeding capability of some accelerometers. Overall good correlation between analysis and measured results for windward and leeward surface temperatures. Generally poor correlation between analysis and measured results for spindle area temperatures. Excellent test-to-test repeatability in strain and deflection data for 100% DLL testing. In-situ thermography images taken before and after thermal testing showed only minor changes in initial defects. Final detailed thermography tests scheduled for completion in Oct 09. In process of completing test documentation and test data analysis. Final reports complete by Dec 09. All analysis, test data, test plans, reports, photos, etc. will be made available to the technical community via the CMC Wiki.

Description: Minimizing mass and volume is critically important for space hardware. Microchannel technology can be used to decrease both of these parameters for heat exchangers. Working in concert with NASA, Pacific Northwest National Laboratories (PNNL) has developed a microchannel liquid/liquid heat exchanger that has resulted in significant mass and volume savings. The microchannel heat exchanger delivers these improvements without sacrificing thermal and pressure drop performance. A conventional heat exchanger has been tested and the performance of it recorded to compare it to the microchannel heat exchanger that PNNL has fabricated. The microchannel heat exchanger was designed to meet all of the requirements of the baseline heat exchanger, while reducing the heat exchanger mass and volume. The baseline heat exchanger was designed to have an transfer approximately 3.1 kW for a specific set of inlet conditions. The baseline heat exchanger mass was 2.7 kg while the microchannel mass was only 2.0 kg. More impressive, however, was the volumetric savings associated with the microchannel heat exchanger. The microchannel heat exchanger was an order of magnitude smaller than the baseline heat exchanger (2180cm3 vs. 311 cm3). This paper will describe the test apparatus designed to complete performance tests for both heat exchangers. Also described in this paper will be the performance specifications for the microchannel heat exchanger and how they compare to the baseline heat exchanger.

Description: NASA's Constellation Program has been developed to successfully return humans to the Lunar surface prior to 2020. The Constellation Program includes several different project offices including Altair, which is the next generation Lunar Lander. The planned Altair missions are very different than the Lunar missions accomplished during the Apollo era. These differences have resulted in a significantly different thermal control system architecture. The Altair project has employed a rather unique development approach as compared with previous manned spacecraft programs. Altair started the design process by developing a single-string (no fault tolerance), minimum functionality design. This first design and analysis cycle resulted in the baseline design for the entire process. From this point of departure, Altair continued the development process by adding vehicle functionality for the purposes of minimizing the risk of Loss Of Crew (LOC) and Loss Of Mission (LOM). Through the subsequent design and analysis cycles, the project office compared the added mass associated with the reduction of LOC/LOM and selected the most mass efficient design solutions. The current paper will summarize the Altair mission profile, the operational phases, and the LOC/LOM decisions that were made during the various design cycles. The evolution of the thermal control system design through Lunar Design and Analysis Cycle 3 (LDAC-3) will also be described in this paper.

Description: The LISA Technology Package (LTP) is the payload of the European Space Agency's LISA Pathfinder mission. LISA Pathfinder was instigated to test, in a flight environment, the critical technologies required by LISA; namely, the inertial sensing subsystem and associated control laws and micro-Newton thrusters required to place a macroscopic test mass in pure free-fall. The UP is in the late stages of development -- all subsystems are currently either in the final stages of manufacture or in test. Available flight units are being integrated into the real-time testbeds for system verification tests. This poster will describe the UP and its subsystems, give the current status of the hardware and test campaign, and outline the future milestones leading to the UP delivery.

Description: Integrated testing (such as Multi-Element Integrated Test (MEIT)) is critical to reducing risks and minimizing problems encountered during assembly, activation, and on-orbit operation of large, complex manned spacecraft. Provides the best implementation of "Test Like You Fly:. Planning for integrated testing needs to begin at the earliest stages of Program definition. Program leadership needs to fully understand and buy in to what integrated testing is and why it needs to be performed. As Program evolves and design and schedules mature, continually look for suitable opportunities to perform testing where enough components are together in one place at one time. The benefits to be gained are well worth the costs.

Description: CanSat is an international student design-build-launch competition organized by the American Astronautical Society (AAS) and American Institute of Aeronautics and Astronautics (AIAA). The competition is also sponsored by the Naval Research Laboratory (NRL), the National Aeronautics and Space Administration (NASA), AGI, Orbital Sciences Corporation, Praxis Incorporated, and SolidWorks. Specifically, the 2009 Virginia Tech CanSat Team is funded by BAE Systems, Incorporated of Manassas, Virginia. The objective of the 2009 CanSat competition is to complete remote sensing missions by designing a small autonomous sounding rocket payload. The payload designed will follow and perform to a specific set of mission requirements for the 2009 competition. The competition encompasses a complete life-cycle of one year which includes all phases of design, integration, testing, reviews, and launch.

Description: Spacecraft reaction wheel maneuvers are limited by the maximum torque and/or angular momentum that the wheels can provide. For an n-wheel configuration, the torque or momentum envelope can be obtained by projecting the n-dimensional hypercube, representing the domain boundary of individual wheel torques or momenta, into three dimensional space via the 3xn matrix of wheel axes. In this paper, the properties of the projected hypercube are discussed, and algorithms are proposed for determining this maximal torque or momentum envelope for general wheel configurations. Practical strategies for distributing a prescribed torque or momentum among the n wheels are presented, with special emphasis on configurations of four, five, and six wheels.

Description: One unique project that the Prototype lab worked on was PORT I (Post-landing Orion Recovery Test). PORT is designed to test and develop the system and components needed to recover the Orion capsule once it splashes down in the ocean. PORT II is designated as a follow up to PORT I that will utilize a mock up pressure vessel that is spatially compar able to the final Orion capsule.

Description: Recent events such as the Chinese anti-satellite missile test in January 2007 and the collision between a Russian Cosmos satellite and US Iridium satellite in February 2009 are responsible for a rapid increase in the population of orbital debris in Low Earth Orbit (LEO). Without active debris removal strategies the debris population in key orbits will continue to increase, requiring enhanced shielding capabilities to maintain allowable penetration risks. One of the more promising developments in recent years for meteoroid and orbital debris shielding (MMOD) is the application of open cell foams. Although shielding onboard the International Space Station is the most capable ever flown, the most proficient configuration (stuffed Whipple shield) requires an additional ~30% of the shielding mass for non-ballistic requirements (e.g. stiffeners, fasteners, etc.). Open cell foam structures provide similar mechanical performance to more traditional structural components such as honeycomb sandwich panels, as well as improved projectile fragmentation and melting as a result of repeated shocking by foam ligaments. In this paper, the preliminary results of an extensive hypervelocity impact test program on next generation MMOD shielding configurations incorporating open-cell metallic foams are reported.

Description: Fragment shape is an important factor for conducting reliable orbital debris damage assessments for critical space assets, such as the International Space Station. To date, seven microsatellite impact tests have been completed as part of an ongoing collaboration between Kyushu University and the NASA Orbital Debris Program Office. The target satellites ranged in size from 15 cm 15 cm 15 cm to 20 cm 20 cm 20 cm. Each target satellite was equipped with fully functional electronics, including circuits, battery, and transmitter. Solar panels and multi-layer insulation (MLI) were added to the target satellites of the last two tests. The impact tests were carried out with projectiles of different sizes and impact speeds. All fragments down to about 2 mm in size were collected and analyzed based on their three orthogonal dimensions, x, y, and z, where x is the longest dimension, y is the longest dimension in the plane perpendicular to x, and z is the longest dimension perpendicular to both x and y. Each fragment was also photographed and classified by shape and material composition. This data set serves as the basis of our effort to develop a fragment shape distribution. Two distinct groups can be observed in the x/y versus y/z distribution of the fragments. Objects in the first group typically have large x/y values. Many of them are needle-like objects originating from the fragmentation of carbon fiber reinforced plastic materials used to construct the satellites. Objects in the second group tend to have small x/y values, and many of them are box-like or plate-like objects, depending on their y/z values. Each group forms the corresponding peak in the x/y distribution. However, only one peak can be observed in the y/z distribution. These distributions and how they vary with size, material type, and impact parameters will be described in detail within the paper.

Description: LISA Pathfinder (formerly known as SMART-2) is a European Space Agency (ESA) mission designed to pave the way for the joint ESA/NASA Laser Interferometer Space Antenna (LISA) mission by testing in flight the critical technologies required for spaceborne gravitational wave detection: it will put two test masses in a near-perfect gravitational free-fall and control and measure their motion with unprecedented accuracy. LISA Pathfinder is currently in the integration and test phase of the development, and is due to be launched on a dedicated launch vehicle in late 2011, with first results on the performance of the system being available approx 6 months later. This poster will describe the mission in detail, give the current status of the spacecraft development, and highlight the future milestones in the integration and test campaign.

Description: The parachutes on the Space Transportation System (STS) Solid Rocket Booster (SRB) are the means for decelerating the SRB and allowing it to impact the water at a nominal vertical velocity of 75 feet per second. Each SRB has one pilot, one drogue, and three main parachutes. About four minutes after SRB separation, the SRB nose cap is jettisoned, deploying the pilot parachute. The pilot chute then deploys the drogue parachute. The drogue chute provides initial deceleration and proper SRB orientation prior to frustum separation. At frustum separation, the drogue pulls the frustum from the SRB and allows the main parachutes that are mounted in the frustum to unpack and inflate. These chutes are retrieved, inspected, cleaned, repaired as needed, and returned to the flight inventory and reused. Over the course of the Shuttle Program, several improvements have been introduced to the SRB main parachutes. A major change was the replacement of the small (115 ft. diameter) main parachutes with the larger (136 ft. diameter) main parachutes. Other modifications were made to the main parachutes, main parachute support structure, and SRB frustum to eliminate failure mechanisms, improve damage tolerance, and improve deployment and inflation characteristics. This reliability analysis is limited to the examination of the SRB Large Main Parachute (LMP) and drogue parachute failure history to assess the reliability of these chutes. From the inventory analysis, 68 Large Main Parachutes were used in 651 deployments, and 7 chute failures occurred in the 651 deployments. Logistic regression was used to analyze the LMP failure history, and it showed that reliability growth has occurred over the period of use resulting in a current chute reliability of R = .9983. This result was then used to determine the reliability of the 3 LMPs on the SRB, when all must function. There are 29 drogue parachutes that were used in 244 deployments, and no in-flight failures have occurred. Since there are no observed drogue chute failures, Jeffreys Prior was used to calculate a reliability of R =.998. Based on these results, it is concluded that the LMP and drogue parachutes on the Shuttle SRB are suited to their mission and changes made over their life have improved the reliability of the parachute.

Description: This paper summarizes the on-orbit structural dynamic data and the related modal analysis, model validation and correlation performed for the ISS configurations spanning ISS Stage 12A. The objective of this analysis is to validate and correlate analytical models used to verify the ISS critical interface dynamic loads and improve its fatigue life prediction. On-Orbit dynamic responses were measured during the ISS configurations throughout ISS Stage 12A by the two main ISS instrumentation systems; Internal Wireless Instrumentation System (IWIS) and the Structural Dynamic Measurement System (SDMS). These nominal on-orbit events include Russian vehicle docking and undockings. Also, the ISS photogrammetric system recorded the movements of the 2A and 4A solar arrays during a modified ISS maneuver. Modal analyses were performed on the measured data to extract modal parameters including frequency, damping and mode shapes. Correlation and comparisons between the test and analytical frequencies and mode shapes were performed to assess the accuracy of the analytical models for the ISS configurations under consideration.

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

Description: NASA Glenn Research Center hosted a 2.5-day workshop, entitled "NASA Lunar Dust Filtration and Separations Workshop" at the Ohio Aerospace Institute in Cleveland, Ohio, on November 18 to 20, 2008. The purpose of the workshop was to address the issues and challenges of particulate matter removal from the cabin atmospheres in the Altair lunar lander, lunar habitats, and in pressurized rovers. The presence of lunar regolith dust inside the pressurized volumes was a theme of particular interest. The workshop provided an opportunity for NASA, industry experts, and academia to identify and discuss the capabilities of current and developing air and gas particulate matter filtration and separations technologies as they may apply to NASA s needs. A goal of the workshop was to provide recommendations for strategic research areas in cabin atmospheric particulate matter removal and disposal technologies that will advance and/or supplement the baseline approach for these future lunar surface exploration missions.

Description: The proposed Xenia mission will, for the first time, chart the chemical and dynamical state of the majority of baryonic matter in the universe. using high-resolution spectroscopy, Xenia will collect essential information from major traces of the formation and evolution of structures from the early universe to the present time. The mission is based on innovative instrumental and observational approaches: observing with fast reaction gamma-ray bursts (GRBs) with a high spectral resolution. This enables the study of their (star-forming) environment from the dark to the local universe and the use of GRBs as backlight of large-scale cosmological structures, observing and surveying extended sources with high sensitivity using two wide field-of-view x-ray telescopes - one with a high angular resolution and the other with a high spectral resolution.

Description: The development of a robust and efficient leak detection and localization system within a space station environment presents a unique challenge. A plausible approach includes the implementation of an acoustic sensor network system that can successfully detect the presence of a leak and determine the location of the leak source. Traditional acoustic detection and localization schemes rely on the phase and amplitude information collected by the sensor array system. Furthermore, the acoustic source signals are assumed to be airborne and far-field. Likewise, there are similar applications in sonar. In solids, there are specialized methods for locating events that are used in geology and in acoustic emission testing that involve sensor arrays and depend on a discernable phase front to the received signal. These methods are ineffective if applied to a sensor detection system within the space station environment. In the case of acoustic signal location, there are significant baffling and structural impediments to the sound path and the source could be in the near-field of a sensor in this particular setting.

Description: A document describes the design of a lightweight (between 100 to 200 kg), light-tight shroud of about 3.9 meters in diameter that could be stowed into a very small volume, and be deployed to 12 meters. The shroud will consist of two concentric multi-layer blankets (MLIs) that are constructed in an accordion shape.

Description: It is well recognized that water handling systems used in a spacecraft are prone to failure caused by biofouling and mineral scaling, which can clog mechanical systems and degrade the performance of capillary-based technologies. Long duration spaceflight applications, such as extended stays at a Lunar Outpost or during a Mars transit mission, will increasingly benefit from hardware that is generally more robust and operationally sustainable overtime. This paper presents potential design and testing considerations for improving the reliability of water handling technologies for exploration spacecraft. Our application of interest is to devise a spacecraft wastewater management system wherein fouling can be accommodated by design attributes of the management hardware, rather than implementing some means of preventing its occurrence.

Description: A study of hybrid material couples using the Spiral Orbit Tribometer (SOT) was initiated to investigate both lubricated (Pennzane X2000 and Brayco 815Z) and unlubricated Si3N4, 440C SS, Rex 20, Cronidur X30 and X40 plates with Cerbec SN-101-C (Si3N4) and 440C balls. The hybrid wheel/bearing assembly will be used on the Linear Optical Delay Line (LODL) stage as an element of the NASA Space Interferometry Mission (SIM). SIM is an orbiting interferometer linking a pair of telescopes within the spacecraft and, by using an interferometry technique and several precision optical stages, is able to measure the motions of known stars much better than current ground or space based systems. This measurement will provide the data to "infer" the existence of any plants, undetectable by other methods, orbiting these known stars.

Description: The Dawn Ion Propulsion System is the ninth project in NASA s Discovery Program. The Dawn spacecraft is being developed to enable the scientific investigation of the two heaviest main-belt asteroids, Vesta and Ceres. Dawn is the first mission to orbit two extraterrestrial bodies, and the first to orbit a main-belt asteroid. The mission is enabled by the onboard Ion Propulsion System (IPS) to provide the post-launch delta-V. The three Ion Engines of the IPS are mounted on Thruster Gimbal Assembly (TGA), with only one engine operating at a time for this 10-year mission. The three TGAs weigh 14.6 kg.

Description: This paper describes a series of studies to assess the potential risk associated with the failure of one of three gaseous hydrogen flow control valves in the orbiter's main propulsion system during the launch of Shuttle Endeavour (STS-126) in November 2008. The studies focused on critical issues associated with the possibility of combustion resulting from release of gaseous hydrogen from the external tank into the atmosphere during assent. The Shuttle Program currently assumes hydrogen venting from the external tank will result in a critical failure. The current effort was conducted to increase understanding of the risk associated with venting hydrogen given the flow control valve failure scenarios being considered in the Integrated In-Flight Anomaly Investigation being conducted by NASA.

Description: Small spacecraft have been increasing in popularity because of their low cost, short turnaround and relative efficiency. In the past, small spacecraft have been primarily used for technology demonstrations, but advances in technology have made the miniaturization of space science possible [1,2]. PharmaSat is a low cost, small three cube size spacecraft, with a biological experiment on board, built at NASA (National Aeronautics and Space Administration) Ames Research Center. The thermal design of small spacecraft presents challenges as their smaller surface areas translate into power and thermal constraints. The spacecraft is thermally designed to run colder in the Low Earth Orbit space environment, and heated to reach the temperatures required by the science payload. The limited power supply obtained from the solar panels on small surfaces creates a constraint in the power used to heat the payload to required temperatures. The pressurized payload is isolated with low thermally conductance paths from the large ambient temperature changes. The thermal design consists of different optical properties of section surfaces, Multi Layer Insulation (MLI), low thermal conductance materials, flexible heaters and thermal spreaders. The payload temperature is controlled with temperature sensors and flexible heaters. Finite Element Analysis (FEA) and testing were used to aid the thermal design of the spacecraft. Various tests were conducted to verify the thermal design. An infrared imager was used on the electronic boards to find large heat sources and eliminate any possible temperature runaways. The spacecraft was tested in a thermal vacuum chamber to optimize the thermal and power analysis and qualify the thermal design of the spacecraft for the mission.

Description: An investigation of the aeroheating environment of the Project Orion Crew Exploration Vehicle was performed in the Arnold Engineering Development Center Hypervelocity Wind Tunnel No. 9 Mach 8 and Mach 10 nozzles and in the NASA Langley Research Center 20 - Inch Mach 6 Air Tunnel. Heating data were obtained using a thermocouple-instrumented approx.0.035-scale model (0.1778-m/7-inch diameter) of the flight vehicle. Runs were performed in the Tunnel 9 Mach 10 nozzle at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 20x10(exp 6)/ft, in the Tunnel 9 Mach 8 nozzle at free stream unit Reynolds numbers of 8 x 10(exp 6)/ft to 48x10(exp 6)/ft, and in the 20-Inch Mach 6 Air Tunnel at free stream unit Reynolds numbers of 1x10(exp 6)/ft to 7x10(exp 6)/ft. In both facilities, enthalpy levels were low and the test gas (N2 in Tunnel 9 and air in the 20-Inch Mach 6) behaved as a perfect-gas. These test conditions produced laminar, transitional and turbulent data in the Tunnel 9 Mach 10 nozzle, transitional and turbulent data in the Tunnel 9 Mach 8 nozzle, and laminar and transitional data in the 20- Inch Mach 6 Air Tunnel. Laminar and turbulent predictions were generated for all wind tunnel test conditions and comparisons were performed with the experimental data to help define the accuracy of computational method. In general, it was found that both laminar data and predictions, and turbulent data and predictions, agreed to within less than the estimated 12% experimental uncertainty estimate. Laminar heating distributions from all three data sets were shown to correlate well and demonstrated Reynolds numbers independence when expressed in terms of the Stanton number based on adiabatic wall-recovery enthalpy. Transition onset locations on the leeside centerline were determined from the data and correlated in terms of boundary-layer parameters. Finally turbulent heating augmentation ratios were determined for several body-point locations and correlated in terms of the boundary-layer momentum Reynolds number.

Description: Since 1996, NASA has been developing a docking system that will simplify operations and reduce risks associated with mating spacecraft. This effort has focused on developing and testing an original, reconfigurable, active, closed-loop, force-feedback controlled docking system using modern technologies. The primary objective of this effort has been to design a docking interface that is tunable to the unique performance requirements for all types of mating operations (i.e. docking and berthing, autonomous and piloted rendezvous, and in-space assembly of vehicles, modules and structures). The docking system must also support the transfer of crew, cargo, power, fluid, and data. As a result of the past 10 years of docking system advancement, the Low Impact Docking System or LIDS was developed. The current LIDS design incorporates the lessons learned and development experiences from both previous and existing docking systems. LIDS feasibility was established through multiple iterations of prototype hardware development and testing. Benefits of LIDS include safe, low impact mating operations, more effective and flexible mission implementation with an anytime/anywhere mating capability, system level redundancy, and a more affordable and sustainable mission architecture with reduced mission and life cycle costs. In 1996 the LIDS project, then known as the Advanced Docking Berthing System (ADBS) project, launched a four year developmental period. At the end of the four years, the team had built a prototype of the soft-capture hardware and verified the control system that will be used to control the soft-capture system. In 2001, the LIDS team was tasked to work with the X- 38 Crew Return Vehicle (CRV) project and build its first Engineering Development Unit (EDU).

Description: No abstract available

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: 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: 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: 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: On 31 July 2008 the National Aeronautics and Space Administration Crew Exploration Vehicle Parachute Assembly System team conducted the final planned cluster test of the first generation parachute recovery system design. The two primary test objectives were to demonstrate the operation of the complete parachute system deployed from a full scale capsule simulator and to demonstrate the test technique of separating the capsule simulator from the Low Velocity Air Drop pallet used to extract the test article from a United States Air Force C-17 aircraft. The capsule simulator was the Parachute Test Vehicle with an accurate heat shield outer mold line and forward bay compartment of the Crew Exploration Vehicle Command Module. The Parachute Test Vehicle separated cleanly from the pallet following extraction, but failed to reach test conditions resulting in the failure of the test and the loss of the test assets. No personnel were injured. This paper will discuss the design of the test and the findings of the team that investigated the test, including a discussion of what were determined to be the root causes of the failure.

Description: Propellant-optimal trajectories, relative sensors and navigation, and docking/capture mechanisms are rendezvous disciplines that receive much attention in the technical literature. However, other areas must be considered. These include absolute navigation, maneuver targeting, attitude control, power generation, software development and verification, redundancy management, thermal control, avionics integration, robotics, communications, lighting, human factors, crew timeline, procedure development, orbital debris risk mitigation, structures, plume impingement, logistics, and in some cases extravehicular activity. While current and future spaceflight programs will introduce new technologies and operations concepts, the complexity of integrating multiple systems on multiple spacecraft will remain. The systems integration task may become more difficult as increasingly complex software is used to meet current and future automation, autonomy, and robotic operation requirements.

Description: During the ascent flight phase of NASA s Constellation Program, the Ares launch vehicle propels the Orion crew vehicle to an agreed to insertion target. If a failure occurs at any point in time during ascent then a system must be in place to abort the mission and return the crew to a safe landing with a high probability of success. To achieve continuous abort coverage one of two sets of effectors is used. Either the Launch Abort System (LAS), consisting of the Attitude Control Motor (ACM) and the Abort Motor (AM), or the Service Module (SM), consisting of SM Orion Main Engine (OME), Auxiliary (Aux) Jets, and Reaction Control System (RCS) jets, is used. The LAS effectors are used for aborts from liftoff through the first 30 seconds of second stage flight. The SM effectors are used from that point through Main Engine Cutoff (MECO). There are two distinct sets of Guidance and Control (G&C) algorithms that are designed to maximize the performance of these abort effectors. This paper will outline the necessary inputs to the G&C subsystem, the preliminary design of the G&C algorithms, the ability of the algorithms to predict what abort modes are achievable, and the resulting success of the abort system. Abort success will be measured against the Preliminary Design Review (PDR) abort performance metrics and overall performance will be reported. Finally, potential improvements to the G&C design will be discussed.

Description: An overview of significant Micrometeoroid and Orbital Debris (MMOD) impacts on the Payload Bay Door radiators, wing leading edge reinforced carbon-carbon panels and crew module windows will be presented, along with a discussion of the techniques NASA has implemented to reduce the risk from MMOD impacts. The concept of "Late Inspection" of the Nose Cap and Wing leading Edge (WLE) Reinforced Carbon Carbon (RCC) regions will be introduced. An alternative mated attitude with the International Space Station (ISS) on shuttle MMOD risk will also be presented. The significant threat mitigation effect of these two techniques will be demonstrated. The wing leading edge impact detection system, on-orbit repair techniques and disabled vehicle contingency plans will also be discussed.

Description: This slide presentation explores some of the issues that concern the engineers and planners of the Altair Lunar landing module. Particular attention is paid to the issues concerning Lunar dust, and attempts that will be made to test the Altair systems using Lunar dust simulants.

Description: To effectively explore the lunar surface, astronauts will need a transportation vehicle which can traverse all types of terrain. Currently, the National Aeronautics and Space Administration s (NASA) is investigating two lunar rover configurations to meet such a requirement. Under the Lunar Electric Rover (LER) project, a comparison study between the unpressurized rover (UPR) and the small pressurized rover (SPR) was conducted at the Black Point Lava Flow in Arizona. The objective of the study was to obtain human-in-the-loop performance data on the vehicles with respect to human-machine interfaces, vehicle impacts on crew productivity, and scientific observations. Four male participants took part in four, one-day field tests using the exact same terrain and scientific sites for an accurate comparison between vehicle configurations. Subjective data was collected using several human factors performance measures. Results indicate either vehicle configuration was generally acceptable for a lunar mission; however, the SPR configuration was preferred over the UPR configuration primarily for the SPR s ability to cause less fatigue and enabling greater crew productivity.

Description: The Magnetospheric MultiScale Mission is a four spacecraft formation flying mission designed to study the Earth s magnetosphere. The spacecraft fly in highly elliptical orbits, forming a tetrahedron at apogee. Each spacecraft spins at 3 RPM and is equipped with a star scanner, slit sun sensor, and accelerometer. The purpose of this work is to develop an Extended Kalman Filter to simultaneously estimate the attitude, angular velocity, angular acceleration, and center of mass of each spacecraft.

Description: A key objective of NASA s Orbital Debris program office at Johnson Space Center (JSC) is to characterize the debris environment by way of assessing the physical properties (type, mass, density, and size) of objects in orbit. Knowledge of the geosynchronous orbit (GEO) debris environment in particular can be used to determine the hazard probability at specific GEO altitudes and aid predictions of the future environment. To calculate an optical size from an intensity measurement of an object in the GEO regime, a 0.175 albedo is assumed currently. However, identification of specific material type or types could improve albedo accuracy and yield a more accurate size estimate for the debris piece. Using spectroscopy, it is possible to determine the surface materials of space objects. The study described herein used the NASA Infrared Telescope Facility (IRTF) to record spectral data in the 0.6 to 2.5 micron regime on eight catalogued space objects. For comparison, all of the objects observed were in GEO or near-GEO. The eight objects consisted of two intact spacecraft, three rocket bodies, and three catalogued debris pieces. Two of the debris pieces stemmed from Titan 3C transtage breakup and the third is from COSMOS 2054. The reflectance spectra of the Titan 3C pieces share similar slopes (increasing with wavelength) and lack any strong absorption features. The COSMOS debris spectra is flat and has no absorption features. In contrast, the intact spacecraft show classic absorption features due to solar panels with a strong band gap feature near 1 micron. The two spacecraft are spin-stabilized objects and therefore have solar panels surrounding the outer surface. Two of the three rocket bodies are inertial upper stage (IUS) rocket bodies and have similar looking spectra. The slopes flatten out near 1.5 microns with absorption features in the near-infrared that are similar to that of white paint. The third rocket body has a similar flattening of slope but with fewer features of white paint - indicating that the surface paint on the SL-12 may be different than the IUS. This study shows that the surface materials of debris appear different spectrally than intact rocket bodies and spacecraft and therefore are not believed to be solar panel material or pristine white paint. Further investigation is necessary in order to eliminate materials as possible choices for the debris pieces.

Description: This paper summarizes the on-orbit structural dynamic data and the related modal analysis, model validation and correlation performed for the ISS configurations spanning ISS Stage 12A. The objective of this analysis is to validate and correlate analytical models used to verify the ISS critical interface dynamic loads and improve its fatigue life prediction. On-Orbit dynamic responses were measured during the ISS configurations throughout ISS Stage 12A by the two main ISS instrumentation systems; Internal Wireless Instrumentation System (IWIS) and the Structural Dynamic Measurement System (SDMS). These nominal on-orbit events include Russian vehicle docking and undockings. Also, the ISS photogrammetric system recorded the movements of the 2A and 4A solar arrays during a modified ISS maneuver. Modal analyses were performed on the measured data to extract modal parameters including frequency, damping and mode shapes. Correlation and comparisons between the test and analytical frequencies and mode shapes were performed to assess the accuracy of the analytical models for the ISS configurations under consideration.

Description: NASA Johnson Space Center White Sands Test Facility (WSTF) performed particle impact testing to determine whether there is a particle impact ignition hazard in the quick disconnects (QDs) in the Environmental Control and Life Support System (ECLSS) on the International Space Station (ISS). Testing included standard supersonic and subsonic particle impact tests on 15-5 PH stainless steel, as well as tests performed on a QD simulator. This paper summarizes the particle impact tests completed at WSTF. Although there was an ignition in Test Series 4, it was determined the ignition was caused by the presence of a machining imperfection. The sum of all the test results indicates that there is no particle impact ignition hazard in the ISS ECLSS QDs. KEYWORDS: quick disconnect, high pressure, particle impact testing, stainless steel

Description: LS-DYNA finite element simulations of a rigid Orion Crew Module (CM) were used to investigate the CM impact behavior on eight different soil models. Ten different landing conditions, characterized by the combination of CM vertical and horizontal velocity, hang angle, and roll angle were simulated on the eight different soils. The CM center of gravity accelerations, pitch angle, kinetic energy, and soil contact forces were the outputs of interest. The simulation results are presented, with comparisons of the CM behavior on the different soils. The soils analyzed in this study can be roughly categorized as soft, medium, or hard, according to the CM accelerations that occur when landing on them. The soft group is comprised of the Carson Sink Wet soil and the Kennedy Space Center (KSC) Low Density Dry Sand. The medium group includes Carson Sink Dry, the KSC High Density In-Situ Moisture Sand and High Density Flooded Sand, and Cuddeback B. The hard soils are Cuddeback A and the Gantry Unwashed Sand. The softer soils were found to produce lower peak accelerations, have more stable pitch behavior, and to be less sensitive to the landing conditions. This investigation found that the Cuddeback A soil produced the highest peak accelerations and worst stability conditions, and that the best landing performance was achieved on the KSC Low Density Dry Sand.

Description: The presence of honeycomb cells in a dual-wall structure is advantageous for mechanical performance and low weight in spacecraft primary structures but detrimental for shielding against impact of micrometeoroid and orbital debris particles (MMOD). The presence of honeycomb cell walls acts to restrict the expansion of projectile and bumper fragments, resulting in the impact of a more concentrated (and thus lethal) fragment cloud upon the shield rear wall. The Multipurpose Laboratory Module (MLM) is a Russian research module scheduled for launch and ISS assembly in 2011 (currently under review). Baseline shielding of the MLM is expected to be predominantly similar to that of the existing Functional Energy Block (FGB), utilizing a baseline triple wall configuration with honeycomb sandwich panels for the dual bumpers and a thick monolithic aluminum pressure wall. The MLM module is to be docked to the nadir port of the Zvezda service module and, as such, is subject to higher debris flux than the FGB module (which is aligned along the ISS flight vector). Without upgrades to inherited shielding, the MLM penetration risk is expected to be significantly higher than that of the FGB module. Open-cell foam represents a promising alternative to honeycomb as a sandwich panel core material in spacecraft primary structures as it provides comparable mechanical performance with a minimal increase in weight while avoiding structural features (i.e. channeling cells) detrimental to MMOD shielding performance. In this study, the effect of replacing honeycomb sandwich panel structures with metallic open-cell foam structures on MMOD shielding performance is assessed for an MLM-representative configuration. A number of hypervelocity impact tests have been performed on both the baseline honeycomb configuration and upgraded foam configuration, and differences in target damage, failure limits, and derived ballistic limit equations are discussed.

Description: Objectives include: a) Define major Command and Service Module (CSM) design considerations; b) List Command Module (CM) RCS failures and lessons learned; and c) List Service Module (SM) RCS failures and lessons learned.

Description: Objectives include: a) Understand some of the design considerations that went into creating the Saturn V launch vehicle; b) Gain an appreciation for some of the manufacturing issues concerning the Saturn V; and c) Review three major problems that affected Saturn V launches.

Description: This slide presentation reviews the first part of the post-touchdown lunar surface and system checkout tasks. A stay/no stay decision for the lunar lander was made based on the questions: "Is the Lunar Module (LM) stable on the lunar surface?"; "Are there any time critical systems failures or trends indicating impending loss of capability to ascent and achieve a safe lunar orbit?"; and "Is there loss of capability in critical LM systems?" The sequence of these decisions is given as a time after touchdown on the surface of the moon. After the decision to stay is made the next task is to checkout status of the lunar module. While the status of the lunar module is checking out certain conditions, the Command Service Module was also engaged in certain checkout activities.

Description: Objectives include: a) Become familiar with the Saturn V Stage I (S-IC) major structural components: Forward Skirt, Oxidizer Tank, Intertank, Fuel Tank, and Thrust Structure. b) Gain a general understanding of the Stage I subsystems: Fuel, Oxidizer, Instrumentation, Flight Control, Environmental Control, Electrical, Control Pressure, and Ordinance.

Description: Objectives include: a) Describe LM Electrical System original specifications; b) Describe the decision to change from fuel cells to batteries and other changes; c) Describe the Electrical system; and d) Describe the Apollo 13 failure from the LM perspective.

Description: Design considerations and failure modes for the Lunar Module (LM) Environmental Control System (ECS) are described. An overview of the the oxygen supply and cabin pressurization, atmosphere revitalization, water management and heat transport systems are provided. Design considerations including reliability, flight instrumentation, modularization and the change to the use of batteries instead of fuel cells are discussed. A summary is provided for the LM ECS general testing regime.

Description: The Crew Exploration Vehicle (CEV) is an element of the Constellation Program that includes launch vehicles, spacecraft, and ground systems needed to embark on a robust space exploration program. As an anchoring capability of the Constellation Program, the CEV shall be human-rated and will carry human crews and cargo from Earth into space and back again. Coupled with transfer stages, landing vehicles, and surface exploration systems, the CEV will serve as an essential component of the architecture that supports human voyages to the Moon and beyond. In addition, the CEV will be modified, as required, to support International Space Station (ISS) mission requirements for crewed and pressurized cargo configurations. Headed by Johnson Space Center (JSC), NASA selected Jacobs Engineering as the support contractor to manage the overall CEV Parachute Assembly System (CPAS) program development. Airborne Systems was chosen to develop the parachute system components. General Dynamics Ordnance and Tactical Systems (GD-OTS) was subcontracted to Airborne Systems to provide the mortar systems. Thus the CPAS development team of JSC, Jacobs, Airborne Systems and GD-OTS was formed. The CPAS team has completed the first phase, or Generation I, of the design, fabrication, and test plan. This paper presents an overview of the CPAS program including system requirements and the development of the second phase, known as the Engineering Development Unit (EDU) architecture. We also present top level results of the tests completed to date. A significant number of ground and flight tests have been completed since the last CPAS presentation at the 2007 AIAA ADS Conference.

Description: This viewgraph presentation seeks to describe the Lunar Module Environmental Control System (ECS) subsystem testing and redesign and seeks to summarize the in-flight failures of the Lunar Module (LM) Environmental Control System (ECS).

Description: NASA's development of new concepts for the Crew Exploration Vehicle Orion presents many similar challenges to those worked in the sixties during the Apollo program. However, with improved modeling capabilities, new challenges arise. For example, the use of the commercial code LS-DYNA, although widely used and accepted in the technical community, often involves high-dimensional, time consuming, and computationally intensive simulations. The challenge is to capture what is learned from a limited number of LS-DYNA simulations to develop models that allow users to conduct interpolation of solutions at a fraction of the computational time. This paper presents a description of the LS-DYNA model, a brief summary of the response surface techniques, the analysis of variance approach used in the sensitivity studies, equations used to estimate impact parameters, results showing conditions that might cause injuries, and concluding remarks.

Description: Synchronized Position Hold, Engage, Reorient, Experimental Satellites (SPHERES) are bowling-ball sized spherical satellites. They will be used inside the space station to test a set of well-defined instructions for spacecraft performing autonomous rendezvous and docking maneuvers. Three free-flying spheres will fly within the cabin of the station, performing flight formations. Each satellite is self-contained with power, propulsion, computers and navigation equipment. The results are important for satellite servicing, vehicle assembly and formation flying spacecraft configurations. SPHERES is a testbed for formation flying by satellites, the theories and calculations that coordinate the motion of multiple bodies maneuvering in microgravity. To achieve this inside the ISS cabin, bowling-ball-sized spheres perform various maneuvers (or protocols), with one to three spheres operating simultaneously . The Synchronized Position Hold, Engage, Reorient, Experimental Satellites (SPHERES) experiment will test relative attitude control and station-keeping between satellites, re-targeting and image plane filling maneuvers, collision avoidance and fuel balancing algorithms, and an array of geometry estimators used in various missions. SPHERES consists of three self-contained satellites, which are 18 sided polyhedrons that are 0.2 meter in diameter and weigh 3.5 kilograms. Each satellite contains an internal propulsion system, power, avionics, software, communications, and metrology subsystems. The propulsion system uses CO2, which is expelled through the thrusters. SPHERES satellites are powered by AA batteries. The metrology subsystem provides real-time position and attitude information. To simulate ground station-keeping, a laptop will be used to transmit navigational data and formation flying algorithms. Once these data are uploaded, the satellites will perform autonomously and hold the formation until a new command is given.

Description: Space Acceleration Measurement System (SAMS-II) is an ongoing study of the small forces (vibrations and accelerations) on the ISS that result from the operation of hardware, crew activities, as well as dockings and maneuvering. Results will be used to generalize the types of vibrations affecting vibration-sensitive experiments. Investigators seek to better understand the vibration environment on the space station to enable future research.

Description: Dual RF Astrodynamic GPS Orbital Navigator Satellite (DRAGONSat) will demonstrate autonomous rendezvous and docking (ARD) in low Earth orbit (LEO) and gather flight data with a global positioning system (GPS) receiver strictly designed for space applications. ARD is the capability of two independent spacecraft to rendezvous in orbit and dock without crew intervention. DRAGONSat consists of two picosatellites (one built by the University of Texas and one built by Texas A and M University) and the Space Shuttle Payload Launcher (SSPL); this project will ultimately demonstrate ARD in LEO.

Description: In the early morning of January 15, 2006, the Stardust Sample Return Capsule (SRC) successfully delivered its precious cargo of cometary particles to the awaiting recovery team at the Utah Test and Training Range (UTTR). As the SRC entered at 12.8 km/s, the fastest manmade object to traverse the atmosphere, a team of researchers imaged the event aboard the NASA DC-8 airborne observatory. At SRC entry, the airplane was at an altitude of 11.9 km positioned within 6.4 km of the prescribed, preferred target view location. The incoming SRC was first acquired approximately 18 seconds (s) after atmospheric interface and tracked for approximately 60 s, an observation period that is roughly centered in time around predicted peak heating.

Description: In the early morning of January 15, 2006, the Stardust Sample Return Capsule (SRC) successfully delivered its precious cargo of cometary particles to the awaiting recovery team at the Utah Test and Training Range (UTTR). As the SRC entered at 12.8 km/s, the fastest manmade object to traverse the atmosphere, a team of researchers imaged the event aboard the NASA DC-8 airborne observatory. At SRC entry, the airplane was at an altitude of 11.9 km positioned within 6.4 km of the prescribed, preferred target view location. The incoming SRC was first acquired approximately 18 seconds (s) after atmospheric interface and tracked for approximately 60 s, an observation period that is roughly centered in time around predicted peak heating.

Description: The NESC eras requested by the NASA Jet Propulsion Laboratory (JPL) to conduct an independent review of the Mars Reconnaissance Orbiter (MRO) Thermal/Vacuum (T/V) Anomaly Assessment. Because the anomaly resulted in the surface contamination of the MRO, selected members of the Materials Super Problem Resolution Team (SPRT) and the NASA technical community having technical expertise relative to contamination issues were chosen for the independent review. The consultation consisted of a review of the MRO Project's reported response to the assessment findings, a detailed review of JPL technical assessment final report, and detailed discussions with the JPL assessment team relative to their findings.

Description: The Deep Impact spacecraft was launched on a Boeing Delta II rocket from Cape Canaveral Air Force Station (CCAFS) on January 12, 2005. Prior to the launch, the Director of the Office of Safety and Mission Assurance (OS&MA) requested the NASA Engineering and Safety Center (NESC) lead a team to render an independent opinion on the rationale for flight and the risk code assignments for the hazard of cracked Thick Film Assemblies (TFAs) in the E-packages of the Delta II launch vehicle for the Deep Impact Mission. The results of the evaluation are contained in this report.

Description: NASA HYP M&S is pursuing the development of SITPS: 1) Working with HYP MDAO to formulate methodology to incorporate SITPS into hypersonic vehicle design trades. 2) SITPS-0 to SITPS-1 (FY10): a) Manufacturing development and weight reduction (5.8 to 3.1 lb(sub m)/sq ft); b) Structural testing to mature SITPS model. 3) SITPS-2 (FY11): a) Focus on panel closeout, panel-to-panel load transfer, and panel curvature. 4) Extend fabrication technology to include alternate cores and insulations (FY12).

Description: A Williams International FJ44-3A 3000-lb thrust class turbofan engine was used as a demonstrator for a Foam-Metal Liner (FML) installed in close proximity to the fan. Two FML designs were tested and compared to the hardwall baseline. Traditional single degree-of-freedom liner designs were also evaluated to provide a comparison. Farfield acoustic levels and limited engine performance results are presented in this paper. The results show that the FML achieved up to 5 dB Acoustic Power Level (PWL) overall attenuation in the forward quadrant, equivalent to the traditional liner design. An earlier report presented the test set-up and conditions.

Description: Spacecraft testing specifications differ greatly in the criteria they specify for stability in thermal balance tests. Some specify a required temperature stabilization rate (the change in temperature per unit time, dT/dt), some specify that the final steady-state temperature be approached to within a specified difference, delta T , and some specify a combination of the two. The particular values for temperature stabilization rate and final temperature difference also vary greatly between specification documents. A one-size-fits-all temperature stabilization rate requirement does not yield consistent results for all test configurations because of differences in thermal mass and heat transfer to the environment. Applying a steady-state temperature difference requirement is problematic because the final test temperature is not accurately known a priori, especially for powered configurations. In the present work, a simplified, lumped-mass analysis has been used to explore the applicability of these criteria. A new, user-friendly, physics-based approach is developed that allows the thermal engineer to determine when an acceptable level of temperature stabilization has been achieved. The stabilization criterion can be predicted pre-test but must be refined during test to allow verification that the defined level of temperature stabilization has been achieved.

Description: The standards for materials and processes surrounding spacecraft are discussed. Presentation focused on minimum requirements for Materials and Processes (M&P) used in design, fabrication, and testing of flight components for NASA manned, unmanned, robotic, launch vehicle, lander, in-space and surface systems, and spacecraft program/project hardware elements.Included is information on flammability, offgassing, compatibility requirements, and processes; both metallic and non-metallic materials are mentioned.

Description: This slide presentation reviews the status and progress made in the Constellation Program's work towards the goal of lunar and Martian exploration flights. It includes views of the various components of the program, and reviews the status of the engine tests, and the development of the Ares I-X towards test launch, the Orion Crew Module, the launch abort system, and the ground operations facilities.

Description: A new generation of spacecraft is now under development by NASA to replace the Space Shuttle and return astronauts to the Moon. These spacecraft will have a manual control capability for several mission tasks, and the ease and precision with which pilots can execute these tasks will have an important effect on mission risk and training costs. This paper focuses on the handling qualities of a spacecraft based on dynamics similar to that of the Crew Exploration Vehicle, during the last segment of the docking task with a space station in low Earth orbit. A previous study established that handling qualities for this task degrade significantly as the level of translation-into-rotation coupling increases. The goal of this study is to evaluate the efficacy of various pilot aids designed to mitigate the handling qualities degradation caused by this coupling. Four pilot tools were ev adluaetead:d-band box/indicator, flight-path marker, translation guidance cues, and feed-forward control. Each of these pilot tools improved handling qualities, generally with greater improvements resulting from using these tools in combination. A key result of this study is that feedforward control effectively counteracts coupling effects, providing solid Level 1 handling qualities for the spacecraft configuration evaluated.

Description: This slide presentation reviews selected lessons that were learned during the design, development, assembly and operation of the International Space Station. The critical importance of standards and common interfaces is emphasized to create a common operation environment that can lead to flexibility and adaptability.

Description: The ability to predict SMART active trailing edge flap rotor loads is explored in this study. Full-scale wind tunnel data recently acquired in the NASA Ames 40- by 80- Foot Wind Tunnel are compared with analytical results from CAMRAD II. For the 5-bladed rotor, two high-speed forward flight cases are considered, namely, a 0 deg flap deflection case and a 5P, 2 deg flap deflection case. Overall, the correlation is reasonable, with the following exceptions: the torsion moment frequency and the chordwise bending moment are under predicted. In general, the effect of the 5P, 2 deg flap motion is captured by the analysis, though there is over prediction in the neighborhood of the 105 deg and 120 deg azimuthal locations. Changes to the flexbeam torsion stiffness are also briefly considered in this study, as this stiffness will be updated in the future. Finally, the indication is that compressibility effects are important, and this suggests that computational fluid dynamics might improve the current correlation.

Description: The advent of the new Ares I launch vehicle has highlighted the need for advanced dynamic analysis tools for variable mass, flexible structures. This system is composed of interconnected flexible stages or components undergoing rapid mass depletion through the consumption of solid or liquid propellant. In addition to large rigid body configuration changes, the system simultaneously experiences elastic deformations. In most applications, the elastic deformations are compatible with linear strain-displacement relationships and are typically modeled using the assumed modes technique. The deformation of the system is approximated through the linear combination of the products of spatial shape functions and generalized time coordinates. Spatial shape functions are traditionally composed of normal mode shapes of the system or even constraint modes and static deformations derived from finite element models of the system. Equations of motion for systems undergoing coupled large rigid body motion and elastic deformation have previously been derived through a number of techniques [1]. However, in these derivations, the mode shapes or spatial shape functions of the system components were considered constant. But with the Ares I vehicle, the structural characteristics of the system are changing with the mass of the system. Previous approaches to solving this problem involve periodic updates to the spatial shape functions or interpolation between shape functions based on system mass or elapsed mission time. These solutions often introduce misleading or even unstable numerical transients into the system. Plus, interpolation on a shape function is not intuitive. This paper presents an approach in which the shape functions are held constant and operate on the changing mass and stiffness matrices of the vehicle components. Each vehicle stage or component finite element model is broken into dry structure and propellant models. A library of propellant models is used to describe the distribution of mass in the fuel tank or Solid Rocket Booster (SRB) case for various propellant levels. Based on the mass consumed by the liquid engine or SRB, the appropriate propellant model is coupled with the dry structure model for the stage. Then using vehicle configuration data, the integrated vehicle model is assembled and operated on by the constant system shape functions. The system mode shapes and frequencies can then be computed from the resulting generalized mass and stiffness matrices for that mass configuration. The rigid body mass properties of the vehicle are derived from the integrated vehicle model. The coupling terms between the vehicle rigid body motion and elastic deformation are also updated from the constant system shape functions and the integrated vehicle model. This approach was first used to analyze variable mass spinning beams and then prototyped into a generic dynamics simulation engine. The resulting code was tested against Crew Launch Vehicle (CLV-)class problems worked in the TREETOPS simulation package and by Wilson [2]. The Ares I System Integration Laboratory (SIL) is currently being developed at the Marshall Space Flight Center (MSFC) to test vehicle avionics hardware and software in a hardware-in-the-loop (HWIL) environment and certify that the integrated system is prepared for flight. The Ares I SIL utilizes the Ares Real-Time Environment for Modeling, Integration, and Simulation (ARTEMIS) tool to simulate the launch vehicle and stimulate avionics hardware. Due to the presence of vehicle control system filters and the thrust oscillation suppression system, which are tuned to the structural characteristics of the vehicle, ARTEMIS must incorporate accurate structural models of the Ares I launch vehicle. The ARTEMIS core dynamics simulation models the highly coupled nature of the vehicle flexible body dynamics, propellant slosh, and vehicle nozzle inertia effects combined with mass and flexible body properties that vary significant with time during the flight. All forces that act on the vehicle during flight must be simulated, including deflected engine thrust force, spatially distributed aerodynamic forces, gravity, and reaction control jet thrust forces. These forces are used to excite an integrated flexible vehicle, slosh, and nozzle dynamics model for the vehicle stack that simulates large rigid body translations and rotations along with small elastic deformations. Highly effective matrix math operations on a distributed, threaded high-performance simulation node allow ARTEMIS to retain up to 30 modes of flex for real-time simulation. Stage elements that separate from the stack during flight are propagated as independent rigid six degrees of freedom (6DOF) bodies. This paper will present the formulation of the resulting equations of motion, solutions to example problems, and describe the resulting dynamics simulation engine within ARTEMIS.

Description: The Tschauner-Hempel equations are widely used to correct the separation distance drifts between a pair of satellites within a constellation in highly elliptical orbits [1]. This set of equations was discretized in the true anomaly angle [1] to be used in a digital steady-state hierarchical controller [2]. This controller [2] performed the drift correction between a pair of satellites within the constellation. The objective of a discretized system is to develop a simple algorithm to be implemented in the computer onboard the satellite. The main advantage of the discrete systems is that the computational time can be reduced by selecting a suitable sampling interval. For this digital system, the amount of data will depend on the sampling interval in the true anomaly angle [3]. The purpose of this paper is to implement the discrete Tschauner-Hempel equations and the steady-state hierarchical controller in the computer onboard the satellite. This set of equations is expressed in the true anomaly angle in which a relation will be formulated between the time and the true anomaly angle domains.

Description: Compartment venting analyses have been performed for the Ares I first stage systems tunnel using both the lumped parameter method and the three-dimensional (31)) transient computational fluid dynamics (CFD) approach. The main objective of venting analyses is to predict the magnitudes of differential pressures across the skin so the integrity of solid walls can be evaluated and properly designed. The lumped parameter method assumes the gas pressure and temperature inside the systems tunnel are spatially uniform, which is questionable since the tunnel is about 1,700 in. long and 4 in. wide. Therefore, 31) transient CFD simulations using the commercial CFD code FLUENT are performed in order to examine the gas pressure and temperature variations inside the tunnel. It was found that the uniform pressure and temperature assumptions inside the systems tunnel are valid during ascent. During reentry, the uniform pressure assumption is also reasonable but the uniform temperature assumption is not valid. Predicted pressure and temperature inside the systems tunnel using CFD are also compared with those from the lumped parameter method using the NASA code CHCHVENT. In general, the average pressure and temperature inside the systems tunnel from CFD are between the burst and crush results from CHCHVENT during both ascent and reentry. The skin differential pressure and pressure inside the systems tunnel relative to freestream pressure from CHCHVENT as well as velocity vectors and streamlines are also discussed in detail.

Description: Air bags were evaluated as the landing attenuation system for earth landing of the Orion Crew Module (CM). Analysis conducted to date shows that airbags are capable of providing a graceful landing of the CM in nominal and off-nominal conditions such as parachute failure, high horizontal winds, and unfavorable vehicle/ground angle combinations, while meeting crew and vehicle safety requirements. The analyses and associated testing presented here surround a second generation of the airbag design developed by ILC Dover, building off of relevant first-generation design, analysis, and testing efforts. In order to fully evaluate the second generation air bag design and correlate the dynamic simulations, a series of drop tests were carried out at NASA Langley s Landing and Impact Research (LandIR) facility in Hampton, Virginia. The tests consisted of a full-scale set of air bags attached to a full-scale test article representing the Orion Crew Module. The techniques used to collect experimental data, develop the simulations, and make comparisons to experimental data are discussed.

Description: This paper introduces the parachutes that have been drop tested in support of the Ares I first stage deceleration system development. The results of the tests show that the reefing ratios for these quarter spherical ribbon parachutes provide the same reefed drag area as historical conical ribbon parachutes. Two sources are investigated for properly normalizing the parachutes relative to their suspension line length, and one is found to be superior.

Description: The Next Generation Advanced Video Guidance Sensor (NGAVGS) is the latest in a line of sensors that have flown four times in the last 10 years. The NGAVGS has been under development for the last two years as a long-range proximity operations and docking sensor for use in an Automated Rendezvous and Docking (AR&D) system. The first autonomous rendezvous and docking in the history of the U.S. Space Program was successfully accomplished by Orbital Express, using the Advanced Video Guidance Sensor (AVGS) as the primary docking sensor. That flight proved that the United States now has a mature and flight proven sensor technology for supporting Crew Exploration Vehicles (CEV) and Commercial Orbital Transport Systems (COTS) Automated Rendezvous and Docking (AR&D). NASA video sensors have worked well in the past: the AVGS used on the Demonstration of Autonomous Rendezvous Technology (DART) mission operated successfully in "spot mode" out to 2 km, and the first generation rendezvous and docking sensor, the Video Guidance Sensor (VGS), was developed and successfully flown on Space Shuttle flights in 1997 and 1998. This paper presents the flight heritage and results of the sensor technology, some hardware trades for the current sensor, and discusses the needs of future vehicles that may rendezvous and dock with the International Space Station (ISS) and other Constellation vehicles. It also discusses approaches for upgrading AVGS to address parts obsolescence, and concepts for minimizing the sensor footprint, weight, and power requirements. In addition, the testing of the various NGAVGS development units will be discussed along with the use of the NGAVGS as a proximity operations and docking sensor.

Description: The Glory observatory is the current incarnation of the Vegetation Canopy Lidar (VCL) mission spacecraft bus. The VCL spacecraft bus, having been cancelled for programmatic reasons in 2000, was nearly integrated when it was put into storage for possible future use. The Glory mission was a suitable candidate for using this spacecraft and in 2006 an effort to recertify the two axis solar array gimbal drive after its extended storage was begun. What was expected to be a simple performance validation of the two dual axis gimbal stepper motors became a serious test, diagnosis and repair task once questions arose on the flight worthiness of the hardware. A significant test program logic flow was developed which identified decisions that could be made based on the results of individual recertification tests. Without disassembling the bi-axial gimbals, beginning with stepper motor threshold voltage measurements and relating these to powered drive torque measurements, both performed at the spacecraft integrator s facility, a confusing picture of the health of the actuators came to light. Tests at the gimbal assembly level and tests of the disassembled actuators were performed by the manufacturer to validate our results and torque discrepancies were noted. Further disassembly to the component level of the actuator revealed the source of the torque loss.

Description: During the past decade, within the United States, NASA Marshall Space Flight Center (MSFC) was heavily engaged in the development of revolutionary new technologies for in-space propulsion. One of the major in-space propulsion technologies developed was a solar sail propulsion system. Solar sail propulsion uses the solar radiation pressure exerted by the momentum transfer of reflected photons to generate a net force on a spacecraft. To date, solar sail propulsion systems have been designed for large spacecraft in the tens to hundreds of kilograms mass range. Recently, however, MSFC has been investigating the application of solar sails for small satellite propulsion. Likewise, NASA Ames Research Center (ARC) has been developing small spacecraft missions that have a need for amass-efficient means of satisfying deorbit requirements. Hence, a synergistic collaboration was established between these two NASA field Centers with the objective of conducting a flight demonstration of solar sail technologies for small satellites. The NanoSail-D mission flew onboard the ill-fated Falcon Rocket launched August 2, 2008, and, due to the failure of that rocket, never achieved orbit. The NanoSail-D flight spare is ready for flight and a suitable launch arrangement is being actively pursued. Both the original sailcraft and the flight spare are hereafter referred to as NanoSail-D. The sailcraft consists of a sail subsystem stowed in a three-element CubeSat. Shortly after deployment of the NanoSail-D, the solar sail will deploy and mission operations will commence. This demonstration flight has two primary technical objectives: (1) to successfully stow and deploy the sail and (2) to demonstrate deorbit functionality. Given a near-term opportunity for launch on Falcon, the project was given the challenge of delivering the flight hardware in 6 mo, which required a significant constraint on flight system functionality. As a consequence, passive attitude stabilization of the spacecraft will be achieved using permanent magnets to detumble and orient the body with the magnetic field lines and then rely on atmospheric drag to passively stabilize the sailcraft in an essentially maximum drag attitude. This paper will present an introduction to solar sail propulsion systems and an overview of the NanoSail-D spacecraft.

Description: The constraint force equation (CFE) methodology provides a framework for modeling constraint forces and moments acting at joints that connect multiple vehicles. With implementation in Program to Optimize Simulated Trajectories II (POST 2), the CFE provides a capability to simulate end-to-end trajectories of launch vehicles, including stage separation. In this paper, the CFE/POST2 methodology is applied to the Shuttle-SRB separation problem as a test and validation case. The CFE/POST2 results are compared with STS-1 flight test data.

Description: In the past, the orbital debris environment was modeled as consisting entirely of aluminum particles. As a consequence, most of the impact test database on spacecraft micro-meteoroid and orbital debris (MMOD) shields, and the resulting ballistic limit equations used to predict shielding performance, has been based on using aluminum projectiles. Recently, data has been collected from returned spacecraft materials and other sources that indicate higher and lower density components of orbital debris also exist. New orbital debris environment models such as ORDEM2008 provide predictions of the fraction of orbital debris in various density bins (high = 7.9 g/cu cm, medium = 2.8 g/cu cm, and low = 0.9-1.1 g/cu cm). This paper describes impact tests to assess the effects of projectile density on the performance capabilities of typical MMOD shields. Updates to shield ballistic limit equations are provided based on results of tests and analysis.

Description: There is often skepticism about the need for Conjunction Assessment from mission operators that invest in the "big sky theory", which states that the likelihood of a collision is so small that it can be neglected. On 10 February 2009, the collision between Iridium 3; and Cosmos 2251 provided an indication that this theory is invalid and that a CA process should be considered for all missions. This paper presents statistics of the effect of the Iridium/Cosmos collision on NASA's Earth Science Constellation as well as results of analyses which characterize the debris environment for NASA's robotic missions.

Description: The Magnetospheric Multiscale (MMS) formation flying mission, which consists of four spacecraft flying in a tetrahedral formation, has challenging navigation requirements associated with determining and maintaining the relative separations required to meet the science requirements. The baseline navigation concept for MMS is for each spacecraft to independently estimate its position, velocity and clock states using GPS pseudorange data provided by the Goddard Space Flight Center-developed Navigator receiver and maneuver acceleration measurements provided by the spacecraft's attitude control subsystem. State estimation is performed onboard in real-time using the Goddard Enhanced Onboard Navigation System flight software, which is embedded in the Navigator receiver. The current concept of operations for formation maintenance consists of a sequence of two maintenance maneuvers that is performed every 2 weeks. Phase 2b of the MMS mission, in which the spacecraft are in 1.2 x 25 Earth radii orbits with nominal separations at apogee ranging from 30 km to 400 km, has the most challenging navigation requirements because, during this phase, GPS signal acquisition is restricted to less than one day of the 2.8-day orbit. This paper summarizes the results from high-fidelity simulations to determine if the MMS navigation requirements can be met between and immediately following the maintenance maneuver sequence in Phase 2b.

Description: The Lunar Reconnaissance Orbiter (LRO) project had a rapid development schedule starting with project conception in spring of 2004, instrument and launch vehicle selection late in 2005 and then launch in early 2009. The lunar thermal environment is one of the harshest in our solar system with the heavy infrared loading of the moon due to low albedo, lack of lunar atmosphere, and low effective regolith conduction. This set of constraints required a thermal design which maximized performance (minimized radiator area and cold control heater power) and minimized thermal hardware build at the orbiter level (blanketing, and heater service). The orbiter design located most of the avionics on an isothermalized heat pipe panel called the IsoThermal Panel (ITP). The ITP was coupled by dual bore heat pipes to an Optical Solar Reflector (OSR) covered heat pipe radiator. By coupling all of the avionics to one system, the hardware was simplified. The seven instruments were mainly heritage instruments which resulted in their desired radiators being located by their heritage design. This minimized instrument redesigns and therefore allowed them to be delivered earlier, though it resulted in a more complex orbiter level blanket and heater service design. Three of the instruments were mounted on a tight pointing M55J optical bench that needed to be covered in heaters to maintain pointing. Two were mounted to spacecraft controlled radiators. One was mounted to the ITP Dual Bores. The last was mounted directly to the bus structure on the moon facing panel. The propulsion system utilized four-20 pound insertion thrusters and eight-5 pound attitude control thrusters (ACS) in addition to 1000 kg of fuel in two large tanks. The propulsion system had a heater cylinder and a heated mounting deck for the insertion thrusters which coupled most of the propulsion design together simplifying the heater design. The High Gain Antenna System (HGAS) and Solar Array System (SAS) used dual axis actuator gimbal systems. HGAS required additional boom heaters to cool the approximately 10 W of RF losses thru the rotary joints and wave guides from the 40 W Ka system. By design this module needed a fair amount of heater, blanketing, and radiator complexity. The SAS system required a separate cable wrap radiator to help cool the Solar Array harness which dissipated 30 W thru the actuators and cable wraps. This module also was complex.

Description: The NASA Engineering and Safety Center (NESC) is developing the Max Launch Abort System (MLAS) as a risk-mitigation design should problems arise with the baseline Orion spacecraft launch abort design. The Max in MLAS is dedicated to Max Faget, the renowned NASA spacecraft designer. The MLAS flight test vehicle consists of boost skirt, coast skirt and the MLAS fairing which houses a full scale boilerplate Orion Crew Module (CM). The objective of the flight test is to prove that the CM can be released from the MLAS fairing during pad abort conditions without detrimental recontact between the CM and fairing, achieving performance similar to the Orion launch abort system. The boost and coast skirts provide the necessary thrust and stability to achieve the flight test conditions and are released prior to the test -- much like the Little Joe booster was used in the Apollo Launch Escape System tests. To achieve the test objective, two parachutes are deployed from the fairing to reorient the CM/fairing to a heatshield first orientation. The parachutes then provide the force necessary to reduce the total angle of attack and body angular rates required for safe release of the CM from the fairing. A secondary test objective after CM release from the fairing is to investigate the removal of the CM forward bay cover (FBC) with CM drogue parachutes for the purpose of attempting to synchronously deploying a set of CM main parachutes. Although multiple parachute deployments are used in the MLAS flight test vehicle to complete its objective, there are only two parachute types employed in the flight test. Five of the nine parachutes used for MLAS are 27.6 ft D(sub 0) ribbon parachutes, and the remaining four are standard G-12 cargo parachutes. This paper presents an overview of the 27.6 ft D(sub 0) ribbon parachute system employed on the MLAS flight test vehicle for coast skirt separation, fairing reorientation, and as drogue parachutes for the CM after separation from the fairing. Discussion will include: the process used to select this design, previously proven as a spin/stall recovery parachute; descriptions of all components of the parachute system; the minor modifications necessary to adapt the parachute to the MLAS program; the techniques used to analyze the parachute for the multiple roles it performs; a discussion of the rigging techniques used to interface the parachute system to the vehicle; and a brief description of how the evolution of the program affected parachute usage and analysis. An overview of the Objective system, rationale for the MLAS approach and the future of the program will also be presented. We hope to have flight test results to report at the time of the Conference Presentation.

Description: The Low Cost Rapid Response Spacecraft (LCRRS) is an ongoing research development project at NASA Ames Research Center (ARC), Moffett Field, California. The prototype spacecraft, called Cost Optimized Test for Spacecraft Avionics and Technologies (COTSAT) is the first of what could potentially be a series of rapidly produced low-cost satellites. COTSAT has a target launch date of March 2009 on a SpaceX Falcon 9 launch vehicle. The LCRRS research system design incorporates use of COTS (Commercial Off The Shelf), MOTS (Modified Off The Shelf), and GOTS (Government Off The Shelf) hardware for a remote sensing satellite. The design concept was baselined to support a 0.5 meter Ritchey-Chretien telescope payload. This telescope and camera system is expected to achieve 1.5 meter/pixel resolution. The COTSAT team is investigating the possibility of building a fully functional spacecraft for $500,000 parts and $2,000,000 labor. Cost is dramatically reduced by using a sealed container, housing the bus and payload subsystems. Some electrical and RF designs were improved/upgraded from GeneSat-1 heritage systems. The project began in January 2007 and has yielded two functional test platforms. It is expected that a flight-qualified unit will be finished in December 2008. Flight quality controls are in place on the parts and materials used in this development with the aim of using them to finish a proto-flight satellite. For LEO missions the team is targeting a mission class requiring a minimum of six months lifetime or more. The system architecture incorporates several design features required by high reliability missions. This allows for a true skunk works environment to rapidly progress toward a flight design. Engineering and fabrication is primarily done in-house at NASA Ames with flight certifications on materials. The team currently employs seven Full Time Equivalent employees. The success of COTSATs small team in this effort can be attributed to highly cross trained engineering team. The engineers on the team are capable of functioning in two to three engineering disciplines which allows highly efficient interdisciplinary engineering collaboration. NASA Ames is actively proposing mission concepts to use the COTSAT platform to accomplish science. If the COTSAT team validates this approach, it will allow the possibility for remote sensing missions to produce a high science yield for minimal cost and reduced schedule. Another aim of this approach is to yield an accelerated pathway from a Phase A study to mission launch. Leaders in the aerospace industry have shown interest in this methodology. Several visits and tours have been given in the lab. Although the concept of low-cost development is initially met with skepticism from some within the prohibitive aerospace industry, the project's efforts have been highly praised for the accomplishments met within a limited time and budget. Overall the development has progressed tremendously well and the team is answering critical questions for current and future low-cost small satellite developments. COTSAT subsystems are not limited to a specific weight class and could be adapted to produce smaller platforms and to fit various launch vehicles.

Description: The Orion Crew Exploration Vehicle is America s next generation of human rated spacecraft. The Orion Launch Abort System will take the astronauts away from the exploration vehicle in the event of an aborted launch. The pad abort mode of the Launch Abort System will be flight-tested in 2009 from the White Sands Missile Range in New Mexico. This paper examines some of the efforts currently underway at the NASA Dryden Flight Research Center by the Controls & Dynamics group in preparation for the flight test. The concept of operation for the pad abort flight is presented along with an overview of the guidance, control and navigation systems. Preparations for the flight test, such as hardware testing and development of the real-time displays, are examined. The results from the validation and verification efforts for the aerodynamic and atmospheric models are shown along with Monte Carlo analysis results.

Description: Whipple shields were first proposed as a means of protecting spacecraft from the impact of micrometeoroids in 1947 [1] and are currently in use as micrometeoroid and orbital debris shields on modern spacecraft. In the intervening years, the function of the thin bumper used to shatter or melt threatening particles has been augmented and enhanced by the use of various types and configurations of intermediate layers of various materials. All shield designs serve to minimize the threat of a spall failure or perforation of the main wall of the spacecraft as a result of the impact of the fragments. With increasing use of Whipple shields, various ballistic limit equations (BLEs) for guiding the design and estimating the performance of shield systems have been developed. Perhaps the best known and most used are the "new" modified Cour-Palais (Christiansen) equations [2]. These equations address the three phases of impact: (1) ballistic (〈3 km/s), where the projectile is moving too slowly to fragment and essentially penetrates as an intact projectile; (2) shatter (3 to 7 km/s), where the projectile fragments at impact and forms an expanding cloud of debris fragments; and (3) melt/vaporization (〉7 km/s), where the projectile melts or vaporizes at impact. The performance of Whipple shields and the adequacy of the BLEs have been examined for the first two phases using the results of impact tests obtained from two-stage, light-gas gun test firings. Shield performance and the adequacy of the BLEs has not been evaluated in the melt/vaporization phase until now because of the limitations of launchers used to accelerate projectiles with controlled properties to velocities above 7.5 km/s. A three-stage, light-gas gun, developed at the University of Dayton Research Institute (UDRI) [3], is capable of launching small, aluminum spheres to velocities above 9 km/s. This launcher was used to evaluate the ballistic performance of two Whipple shield systems, various thermal protection system materials, and other spacecraft-related materials to the impact of 1.6-mm- to 2.6-mm-diameter, 2017-T4 aluminum spheres at impact velocities ranging from 8.91 km/s to 9.28 km/s. Test results, details of the shield systems, and nominal ballistic limits for the two Whipple shields are shown in Figures 1 and 2.

Description: The STS-125 mission, launched May 11, 2009, is the final servicing mission to the Hubble Space Telescope. The repair mission's EVA tasks are described, including: installing a new wide field camera; installing the Cosmic Origins Spectrograph; repairing the Space Telescope Imaging Spectrograph; installing a new outer blanket layer; adding a Soft Capture and Rendezvous System for eventual controlled deorbit in about 2014; replacing the 'A' side Science Instrument Command and Data Handling module; repairing the Advanced Camera for surveys; and, replacing the rate sensor unit gyroscopes, fine guidance sensors and 3 batteries. Additionally, the Shuttle crew cabin thermal environment is described. A CFD model of per person CO2 demonstrates a discrepancy between crew breathing volume and general mid-deck levels of CO2. A follow-on CFD analysis of the mid-deck temperature distribution is provided. Procedural and engineering mitigation plans are presented to counteract thermal exposure upon reentry to the Earth atmosphere. Some of the procedures include: full cold soak the night prior to deorbit; modifying deck stowage to reduce interference with air flow; and early securing of avionics post-landing to reduce cabin thermal load prior to hatch opening. Engineering mitigation activities include modifying the location of the aft starboard ICUs, eliminating the X3 stack and eliminating ICU exhaust air directed onto astronauts; improved engineering data of ICU performance; and, verifying the adequacy of mid-deck temperature control using CFD models in addition to lumped parameter models. Post-mitigation CFD models of mid-deck temperature profiles and distribution are provided.

Description: An aeroelastic analysis of the behavior of an entry vehicle utilizing an attached inflatable aerodynamic decelerator during supersonic flight is presented. The analysis consists of a planar, four degree of freedom simulation. The aeroshell and the IAD are assumed to be separate, rigid bodies connected with a spring-damper at an interface point constraining the relative motion of the two bodies. Aerodynamic forces and moments are modeled using modified Newtonian aerodynamics. The analysis includes the contribution of static aerodynamic forces and moments as well as pitch damping. Two cases are considered in the analysis: constant velocity flight and planar free flight. For the constant velocity and free flight cases with neutral pitch damping, configurations with highly-stiff interfaces exhibit statically stable but dynamically unstable aeroshell angle of attack. Moderately stiff interfaces exhibit static and dynamic stability of aeroshell angle of attack due to damping induced by the pitch angle rate lag between the aeroshell and IAD. For the free-flight case, low values of both the interface stiffness and damping cause divergence of the aeroshell angle of attack due to the offset of the IAD drag force with respect to the aeroshell center of mass. The presence of dynamic aerodynamic moments was found to influence the stability characteristics of the vehicle. The effect of gravity on the aeroshell angle of attack stability characteristics was determined to be negligible for the cases investigated.

Description: The supersonic aerodynamic and aeroelastic characteristics of a tension cone inflatable aerodynamic decelerator were investigated by wind tunnel testing. Two sets of tests were conducted: one using rigid models and another using textile models. Tests using rigid models were conducted over a Mach number range from 1.65 to 4.5 at angles of attack from -12 to 20 degrees. The axial, normal, and pitching moment coefficients were found to be insensitive to Mach number over the tested range. The axial force coefficient was nearly constant (C(sub A) = 1.45 +/- 0.05) with respect to angle of attack. Both the normal and pitching moment coefficients were nearly linear with respect to angle of attack. The pitching moment coefficient showed the model to be statically stable about the reference point. Schlieren images and video showed a detached bow shock with no evidence of large regions of separated flow and/or embedded shocks at all Mach numbers investigated. Qualitatively similar static aerodynamic coefficient and flow visualization results were obtained using textile models at a Mach number of 2.5. Using inflatable textile models the torus pressure required to maintain the model in the fully-inflated configuration was determined. This pressure was found to be sensitive to details in the structural configuration of the inflatable models. Additional tests included surface pressure measurements on rigid models and deployment and inflation tests with inflatable models.

Description: Deployable aeroshells offer the promise of achieving larger aeroshell surface areas for entry vehicles than otherwise attainable without deployment. With the larger surface area comes the ability to decelerate high-mass entry vehicles at relatively low ballistic coefficients. However, for an aeroshell to perform even at the low ballistic coefficients attainable with deployable aeroshells, a flexible thermal protection system (TPS) is required that is capable of surviving reasonably high heat flux and durable enough to survive the rigors of construction handling, high density packing, deployment, aerodynamic loading and aerothermal heating. The Program for the Advancement of Inflatable Decelerators for Atmospheric Entry (PAIDAE) is tasked with developing the technologies required to increase the technology readiness level (TRL) of inflatable deployable aeroshells, and one of several of the technologies PAIDAE is developing for use on inflatable aeroshells is flexible TPS. Several flexible TPS layups were designed, based on commercially available materials, and tested in NASA Langley Research Center's 8 Foot High Temperature Tunnel (8ft HTT). The TPS layups were designed for, and tested at three different conditions that are representative of conditions seen in entry simulation analyses of inflatable aeroshell concepts. Two conditions were produced in a single run with a sting-mounted dual wedge test fixture. The dual wedge test fixture had one row of sample mounting locations (forward) at about half the running length of the top surface of the wedge. At about two thirds of the running length of the wedge, a second test surface drafted up at five degrees relative to the first test surface established the remaining running length of the wedge test fixture. A second row of sample mounting locations (aft) was positioned in the middle of the running length of the second test surface. Once the desired flow conditions were established in the test section the dual wedge test fixture, oriented at 5 degrees angle of attack down, was injected into the flow. In this configuration the aft sample mounting location was subjected to roughly twice the heat flux and surface pressure of the forward mounting location. The tunnel was run at two different conditions for the test series: 1) 'Low Pressure', and 2) 'High Pressure'. At 'Low Pressure' conditions the TPS layups were tested at 6W/cm2 and 11W/cm2 while at 'High Pressure' conditions the TPS layups were tested at 11W/cm2 and 20W/cm2. This paper details the test configuration of the TPS samples in the 8Ft HTT, the sample holder assembly, TPS sample layup construction, sample instrumentation, results from this testing, as well as lessons learned.

Description: Airborne Systems North America (formally Irvin Aerospace Inc) has developed an Airbag Landing System for the Orion Crew Module of the Crew Exploration Vehicle. This work is in support of the NASA Langley Research Center Landing System Advanced Development Project. Orion is part of the Constellation Program to send human explorers back to the moon, and then onwards to Mars and other destinations in the Solar System. A component of the Vision for Space Exploration, Orion is being developed to also enable access to space following the retirement of the Space Shuttle in the next decade. This paper documents the development of a conceptual design, fabrication of prototype assemblies, component level testing and two generations of airbag landing system testing. The airbag system has been designed and analyzed using the transient dynamic finite element code LS-DYNA(RegisteredTradeMark). The landing system consists of six airbag assemblies; each assembly comprising a primary impact venting airbag and a non-venting anti-bottoming airbag. The anti-bottoming airbag provides ground clearance following the initial impact attenuation sequence. Incorporated into each primary impact airbag is an active vent that allows the entrapped gas to exit the control volume. The size of the vent is tailored to control the flow-rate of the exiting gas. An internal shaping structure is utilized to control the shape of the primary or main airbags prior to ground impact; this significantly improves stroke efficiency and performance.

Description: This slide presentation reviews the study done on alternate attitudes for the Orion project's crew exploration vehicle. The analysis focused on the thermal performance of the vehicle with the alternate attitudes. The pressure vessel heater power, other vehicle heaters and radiator sink temperatures were included in the analysis.

Description: The Orion spacecraft will replace the Space Shuttle Orbiter for American and international partner access to the International Space Station by 2015 and, afterwards, for access to the moon for initial sorties and later for extend outpost visits as part of the Constellation Exploration Initiative. This work describes some of the efforts being undertaken to ensure that Orion design will meet or exceed the stringent MicroMeteoroid and Orbital Debris (MMOD) requirements set out by NASA when exposed to the environments encountered with these missions. This paper will provide a brief overview of the approaches being used to provide MMOD protection to the Orion vehicle and to assess the spacecraft for compliance to the Constellation Program s MMOD requirements.

Description: Determination that equipment can operate in and survive exposure to the humidity environments unique to human rated spacecraft presents widely varying challenges. Equipment may need to operate in habitable volumes where the atmosphere contains perspiration, exhalation, and residual moisture. Equipment located outside the pressurized volumes may be exposed to repetitive diurnal cycles that may result in moisture absorption and/or condensation. Equipment may be thermally affected by conduction to coldplate or structure, by forced or ambient air convection (hot/cold or wet/dry), or by radiation to space through windows or hatches. The equipment s on/off state also contributes to the equipment s susceptibility to humidity. Like-equipment is sometimes used in more than one location and under varying operational modes. Due to these challenges, developing a test scenario that bounds all physical, environmental and operational modes for both pressurized and unpressurized volumes requires an integrated assessment to determine the "worst-case combined conditions." Such an assessment was performed for the Constellation program, considering all of the aforementioned variables; and a test profile was developed based on approximately 300 variable combinations. The test profile has been vetted by several subject matter experts and partially validated by testing. Final testing to determine the efficacy of the test profile on actual space hardware is in the planning stages. When validation is completed, the test profile will be formally incorporated into NASA document CxP 30036, "Constellation Environmental Qualification and Acceptance Testing Requirements (CEQATR)."

Description: The carbon wall oxidation technique coupled with a CFD technique was employed to study the flow in the expanding crack channel caused by the oxidation of the channel carbon walls. The recessing 3D surface morphing procedure was developed and tested in comparison with the arcjet experimental results. The multi-block structured adaptive meshing was used to model the computational domain changes due to the wall recession. Wall regression rates for a reinforced carbon-carbon (RCC) samples, that were tested in a high enthalpy arcjet environment, were computationally obtained and used to assess the channel expansion. The test geometry and flow conditions render the flow regime as the transitional to continuum, therefore Navier-Stokes gas dynamic approach with the temperature jump and velocity slip correction to the boundary conditions was used. The modeled mechanism for wall material loss was atomic oxygen reaction with bare carbon. The predicted channel growth was found to agree with arcjet observations. Local gas flow field results were found to affect the oxidation rate in a manner that cannot be predicted by previous mass loss correlations. The method holds promise for future modeling of materials gas-dynamic interactions for hypersonic flight.

Description: Efforts to form this Alliance began in 2008 to showcase the impact testing capabilities within the southern United States. Impact testing customers can utilize SITA partner capabilities to provide supporting data during all program phases-materials/component/ flight hardware design, development, and qualification. This approach would allow programs to reduce risk by providing low cost testing during early development to flush out possible problems before moving on to larger scale1 higher cost testing. Various SITA partners would participate in impact testing depending on program phase-materials characterization, component/subsystem characterization, full-scale system testing for qualification. SITA partners would collaborate with the customer to develop an integrated test approach during early program phases. Modeling and analysis validation can start with small-scale testing to ensure a level of confidence for the next step large or full-scale conclusive test shots. Impact Testing Facility (ITF) was established and began its research in spacecraft debris shielding in the early 1960's and played a malor role in the International Space Station debris shield development. As a result of return to flight testing after the loss of STS-107 (Columbia) MSFC ITF realized the need to expand their capabilities beyond meteoroid and space debris impact testing. MSFC partnered with the Department of Defense and academic institutions as collaborative efforts to gain and share knowledge that would benefit the Space Agency as well as the DoD. MSFC ITF current capabilities include: Hypervelocity impact testing, ballistic impact testing, and environmental impact testing.

Description: The Floating Potential Measurement Unit (FPMU) has detected high negative amplitude rapid charging events (RCEs) on the International Space Station (ISS) at the morning terminator. These events are larger and more rapid than the ISS morning charging events first seen by the Floating Potential Probe (FPP) on ISS in 2001. In this paper, we describe a theory for the RCEs that further elucidates the nature of spacecraft charging in low Earth orbit (LEO) in a non-equilibrium situation. The model accounts for all essential aspects of the newly discovered phenomenon, and is amenable to testing on-orbit. Predictions of the model for the amplitude of the ISS RCEs for the full set of ISS solar arrays and for the coming solar cycle are given, and the results of modeling by the Environments WorkBench (EWB) are compared to the observed events to show that the phenomenon can be explained by solar array driven charging. The situation is unique because the coverglasses have not yet reached equilibrium with the surrounding plasma during the RCEs. Finally, a prescription for further use of the ISS for investigating fundamental plasma physics in LEO is given. Already, plasma and charging monitoring instruments on ISS have taught us much about spacecraft interactions with the dense LEO plasma, and we expect they will continue to yield more valuable science when the Japanese Experiment Module (JEM) is in place.

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: Analyzing the remains of Space Shuttle Columbia has proven technically beneficial years after the vehicle breakup. This investigation focused on charred deposits on fragments of Columbia overhead windowpanes. Results were unexpected relative to the engineering understanding of material performance in a reentry environment. The TEM analysis demonstrated that the oxides of aluminum and titanium mixed with silicon oxides to preserve a history of thermal conditions to which portions of the vehicle were exposed. The presence of Ti during the beginning of the deposition process, along with the thermodynamic phase precipitation upon cool down, indicate that temperatures well above the Ti melt point were experienced. The stratified observations implied that additional exothermic reaction, expectedly metal combustion of a Ti structure, had to be present for oxide formation. Results are significant for aerospace vehicles where thermal protection system (TPS) breaches cause substructures to be in direct path with the reentry plasma. 1

Description: The JSC Flight Safety Office has developed this compilation of historical information on spacecraft crew hatches to assist the Safety Tech Authority in the evaluation and analysis of worldwide spacecraft crew hatch design and performance. The document is prepared by SAIC s Gary Johnson, former NASA JSC S&MA Associate Director for Technical. Mr. Johnson s previous experience brings expert knowledge to assess the relevancy of data presented. He has experience with six (6) of the NASA spacecraft programs that are covered in this document: Apollo; Skylab; Apollo Soyuz Test Project (ASTP), Space Shuttle, ISS and the Shuttle/Mir Program. Mr. Johnson is also intimately familiar with the JSC Design and Procedures Standard, JPR 8080.5, having been one of its original developers. The observations and findings are presented first by country and organized within each country section by program in chronological order of emergence. A host of reference sources used to augment the personal observations and comments of the author are named within the text and/or listed in the reference section of this document. Careful attention to the selection and inclusion of photos, drawings and diagrams is used to give visual association and clarity to the topic areas examined.