Marine Biodiversity Data Flow in the UK

The report provides a review of the current level of exchange in marine life data and its management in the UK taking into account the current structures that are in place between data providers, custodians and managers. In addition, the report makes recommendations on how data flow can be improved over the next few years to achieve greater exchange and interoperability within the marine sector

Solving Problems Caused by Small Micrometeoroid and Orbital Debris Impacts for Space-Walking Astronauts

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Predicting the Consequences of MMOD Penetrations on the International Space Station

The threat from micrometeoroid and orbital debris (MMOD) impacts on space vehicles is often quantified in terms of the probability of no penetration (PNP). However, for large spacecraft, especially those with multiple compartments, a penetration may have a number of possible outcomes. The extent of the damage (diameter of hole, crack length or penetration depth), the location of the damage relative to critical equipment or crew, crew response, and even the time of day of the penetration are among the many factors that can affect the outcome. For the International Space Station (ISS), a Monte-Carlo style software code called Manned Spacecraft Crew Survivability (MSCSurv) is used to predict the probability of several outcomes of an MMOD penetration-broadly classified as loss of crew (LOC), crew evacuation (Evac), loss of escape vehicle (LEV), and nominal end of mission (NEOM). By generating large numbers of MMOD impacts (typically in the billions) and tracking the consequences, MSCSurv allows for the inclusion of a large number of parameters and models as well as enabling the consideration of uncertainties in the models and parameters. MSCSurv builds upon the results from NASA's Bumper software (which provides the probability of penetration and critical input data to MSCSurv) to allow analysts to estimate the probability of LOC, Evac, LEV, and NEOM. This paper briefly describes the overall methodology used by NASA to quantify LOC, Evac, LEV, and NEOM with particular emphasis on describing in broad terms how MSCSurv works and its capabilities and most significant models

Investigation of MMOD Impact on STS-115 Shuttle Payload Bay Door Radiator

The Orbiter radiator system consists of eight individual 4.6 m x 3.2 m panels located with four on each payload bay door. Forward panels #1 and #2 are 2.3 cm thick while the aft panels #3 and #4 have a smaller overall thickness of 1.3 cm. The honeycomb radiator panels consist of 0.028 cm thick Aluminum 2024-T81 facesheets and Al5056-H39 cores. The face-sheets are topped with 0.005 in. (0.127 mm) silver-Teflon tape. The radiators are located on the inside of the shuttle payload bay doors, which are closed during ascent and reentry, limiting damage to the on-orbit portion of the mission. Post-flight inspections at the Kennedy Space Center (KSC) following the STS-115 mission revealed a large micrometeoroid/orbital debris (MMOD) impact near the hinge line on the #4 starboard payload bay door radiator panel. The features of this impact make it the largest ever recorded on an orbiter payload bay door radiator. The general location of the damage site and the adjacent radiator panels can be seen in Figure 2. Initial measurements of the defect indicated that the hole in the facesheet was 0.108 in. (2.74 mm) in diameter. Figure 3 shows an image of the front side damage. Subsequent observations revealed exit damage on the rear facesheet. Impact damage features on the rear facesheet included a 0.03 in. diameter hole (0.76 mm), a approx.0.05 in. tall bulge (approx.1.3 mm), and a larger approx.0.2 in. tall bulge (approx.5.1 mm) that exhibited a crack over 0.27 in. (6.8 mm) long. A large approx.1 in. (25 mm) diameter region of the honeycomb core was also damaged. Refer to Figure 4 for an image of the backside damage to the panel. No damage was found on thermal blankets or payload bay door structure under the radiator panel. Figure 5 shows the front facesheet with the thermal tape removed. Ultrasound examination indicated a maximum facesheet debond extent of approximately 1 in. (25 mm) from the entry hole. X-ray examinations revealed damage to an estimated 31 honeycomb cells with an extent of 0.85 in. x 1.1 in. (21.6 x 27.9 mm). Pieces of the radiator at and surrounding the impact site were recovered during the repair procedures at KSC. They included the thermal tape, front facesheet, honeycomb core, and rear facesheet. These articles were examined at JSC using a scanning electron microscope (SEM) with an energy dispersive x-ray spectrometer (EDS). Figure 6 shows SEM images of the entry hole in the facesheet. The asymmetric height of the lip may be attributed to projectile shape and impact angle. Numerous instances of a glass-fiber organic matrix composite were observed in the facesheet tape sample. The fibers were approximately 10 micrometers in diameter and variable lengths. EDS analysis indicated a composition of Mg, Ca, Al, Si, and O. Figures 7 and 8 present images of the fiber bundles, which were believed to be circuit board material based on similarity in fiber diameter, orientation, consistency, and composition. A test program was initiated in an attempt to simulate the observed damage to the radiator facesheet and honeycomb. Twelve test shots were performed using projectiles cut from a 1.6 mm thick fiberglass circuit board substrate panel. Results from test HITF07017, shown in figures 9 and 10, correlates with the observed impact features reasonably well. The test was performed at 4.14 km/sec with an impact angle of 45 degrees using a cylindrical projectile with a diameter and length of 1.25 mm. The fiberglass circuit board material had a density of 1.65 g/cu cm, giving a projectile mass of 2.53 mg. An analysis was performed using the Bumper code to estimate the probability of impact to the shuttle from a 1.25 mm diameter particle. Table 1 shows a 1.6% chance (impact odds = 1 in 62) of a 1.25 mm or larger MMOD impact on the radiators of the vehicle during a typical ISS mission. There is a 0.4% chance (impact odds = 1 in 260) that a 1.25 mm or larger MMOD particle would impact the RCC wing leading edge and nose cap during a typical miion. Figure 11 illustrates the vulnerable areas of the wing leading edge reinforced carbon-carbon (RCC), an area of the vehicle that is very sensitive to impact damage. The highlighted red, orange, yellow, and light green areas would be expected to experience critical damage if impacted by an OD particle such as the one that hit the RH4 radiator panel on STS-115

Factors affecting internal standard selection for quantitative elemental bio-imaging of soft tissues by LA-ICP-MS

Element response variations under different laser ablation-inductively coupled plasma-mass spectrometry (LA-ICP-MS) operating conditions were investigated to identify important factors for selecting an internal standard (IS) for quantitative elemental bio-imaging. Analytes covering a range of atomic masses and first ionisation potentials (FIP) were selected to investigate the signal response variation with changes in laser spot diameter, mass bias and cell sampling position. In all cases, an IS improved experimental precision regardless of a close match in element mass or FIP but optimal analyte/IS combinations depended on the difference in masses of the analyte and IS. Particular attention was paid to 13C as this isotope is typically used as an IS in elemental bio-imaging applications. Despite its non-ideal IS characteristics (often different mass and FIP to many analytes), possibility of abundance sensitivity effects and poor signal-to-background ratio, 13C was a suitable IS candidate exhibiting a linear response with respect to the mass ablated, apparent independence from the high abundance of the adjacent 14N mass peak and effective analyte normalisation after background subtraction as long as the 13C signal from the sample was at least 6% of the gross signal. © 2011 The Royal Society of Chemistry

Hypervelocity Impact Performance of Open Cell Foam Core Sandwich Panel Structures

Open cell metallic foam core sandwich panel structures are of interest for application in spacecraft micrometeoroid and orbital debris shields due to their novel form and advantageous structural and thermal performance. Repeated shocking as a result of secondary impacts upon individual foam ligaments during the penetration process acts to raise the thermal state of impacting projectiles ; resulting in fragmentation, melting, and vaporization at lower velocities than with traditional shielding configurations (e.g. Whipple shield). In order to characterize the protective capability of these structures, an extensive experimental campaign was performed by the Johnson Space Center Hypervelocity Impact Technology Facility, the results of which are reported in this paper. Although not capable of competing against the protection levels achievable with leading heavy shields in use on modern high-risk vehicles (i.e. International Space Station modules), metallic foam core sandwich panels are shown to provide a substantial improvement over comparable structural panels and traditional low weight shielding alternatives such as honeycomb sandwich panels and metallic Whipple shields. A ballistic limit equation, generalized in terms of panel geometry, is derived and presented in a form suitable for application in risk assessment codes

Hypervelocity Impact Performance of 3D Printed Aluminum Panels

With the continued development of additive manufacturing methods, control over the shape of ligaments, cell regularity, and macroscopic shape can all be easily tuned. This capability allows for tailoring of component architecture and promotes potential mass savings in a space vehicle structure. Additionally, it allows one the flexibility of combining structural elements such as MMOD protection and vehicle stiffness for launch loads for an overall mass reduction. At NASA JSC this technology is being explored in many different ways with the goal being a multifunctional structural component. For this study, four different types of aluminum panels have been 3D printed for testing, three being of a body centric cubic (BCC) lattice structure core and one being kelvin cell structure core. All samples have a 5.33 cm (0.05) nominally thick aluminum face sheet printed on the front and back side of each panel, with all core materials having a 5.08 cm (2.0) nominal thickness (see Table 1 for test sample summary and Figures 1 2 for sample illustrations). These tests will evaluate the performance of 3D printed aluminum panels under hypervelocity impact (HVI) conditions. The hypervelocity impact tests are being conducted at the JSC White Sands Test Facility (WSTF) Remote Hypervelocity Test Laboratory (RHTL), located in Las Cruces, New Mexico. All tests will be conducted with a 3.4mm Al 2017-T4 sphere at 6.8 km/s impacting at 0 to surface normal (i.e., impacting with no obliquity). Each sample will be trapped between two metal frames, with gasket material residing between the sample and frame, which will be the shipping and testing configuration for all tests. There will be an Al 2017-T4 witness plate staged 5.08 cm (2.0) from each sample to capture signature of debris, if the rear face sheet of the sample were to perforate from the HVI test event

Atmospheric and oceanic impacts of Antarctic glaciation across the Eocene-Oligocene transition

The glaciation of Antarctica at the Eocene–Oligocene transition (approx. 34 million years ago) was a major shift in the Earth’s climate system, but the mechanisms that caused the glaciation, and its effects, remain highly debated. A number of recent studies have used coupled atmosphere–ocean climate models to assess the climatic effects of Antarctic glacial inception, with often contrasting results. Here, using the HadCM3L model, we show that the global atmosphere and ocean response to growth of the Antarctic ice sheet is sensitive to subtle variations in palaeogeography, using two reconstructions representing Eocene and Oligocene geological stages. The earlier stage (Eocene; Priabonian), which has a relatively constricted Tasman Seaway, shows a major increase in sea surface temperature over the Pacific sector of the Southern Ocean in response to the ice sheet. This response does not occur for the later stage (Oligocene; Rupelian), which has a more open Tasman Seaway. This difference in temperature response is attributed to reorganization of ocean currents between the stages. Following ice sheet expansion in the earlier stage, the large Ross Sea gyre circulation decreases in size. Stronger zonal flow through the Tasman Seaway allows salinities to increase in the Ross Sea, deep-water formation initiates and multiple feedbacks then occur amplifying the temperature response. This is potentially a model-dependent result, but it highlights the sensitive nature of model simulations to subtle variations in palaeogeography, and highlights the need for coupled ice sheet–climate simulations to properly represent and investigate feedback processes acting on these time scales

Failure Mechanisms of Ni-H2 and Li-Ion Batteries Under Hypervelocity Impacts

Lithium-Ion (Li-Ion) batteries have yielded significant performance advantages for many industries, including the aerospace industry, and have been selected to replace nickel hydrogen (Ni-H2) batteries for the International Space Station (ISS) program to meet the energy storage demands. As the ISS uses its vast solar arrays to generate its power, the solar arrays meet their sunlit power demands and supply excess power to battery packs for power delivery on the sun obscured phase of the approximate 90 minute low Earth orbit. These large battery packs are located on the exterior of the ISS, and as such, the battery packs are exposed to external environment threats like naturally occurring meteoroids and artificial orbital debris (MMOD). While the risks from these solid particle environments has been known and addressed to an acceptable risk of failure through shield design, it is not possible to completely eliminate the risk of loss of these assets on orbit due to MMOD, and as such, failure consequences to the ISS have been considered

Multi-Shock Shield Performance at 15 MJ for Catalogued Debris

While orbital debris of ten centimeters or more are tracked and catalogued, the difficulty of finding and accurately accounting for forces acting on the objects near the ten centimeter threshold results in both uncertainty of their presence and location. These challenges result in difficult decisions for operators balancing potential costly operational approaches with system loss risk. In this paper, the assessment of the feasibility of protecting a spacecraft from this catalogued debris is described using numerical simulations and a test of a multi-shock shield system against a cylindrical projectile impacting normal to the surface with approximately 15 MJ of kinetic energy. The hypervelocity impact test has been conducted at the Arnold Engineering Development Complex (AEDC) with a 598 g projectile at 6.905 km/s on a NASA supplied multi-shock shield. The projectile used is a hollow aluminum and nylon cylinder with an outside diameter of 8.6 cm and length of 10.3 cm. Figure 1 illustrates the multi-shock shield test article, which consisted of five separate bumpers, four of which are fiberglass fabric and one of steel mesh, and two rear walls, each consisting of Kevlar fabric. The overall length of the test article was 2.65 m. The test article was a 5X scaled-up version of a smaller multi-shock shield previously tested using a 1.4 cm diameter aluminum projectile for an inflatable module project. The distances represented by S1 and S1/2 in the figure are 61 cm and 30.5 cm, respectively. Prior to the impact test, hydrodynamic simulations indicated that some enhancement to the standard multi-shock system is needed to address the effects of the cylindrical shape of the projectile. Based on the simulations, a steel mesh bumper has been added to the shield configuration to enhance the fragmentation of the projectile. The AEDC test occurred as planned, and the modified NASA multi-shock shield successfully stopped 598 g projectile using 85.6 kg/m(exp 2). The fifth bumper layer remained in tact, although it was torn free from its support structure and thrown into the first rear wall. The outer Kevlar layer of the first rear wall tore likely from the impact of the fifth bumper's support structure, but the back of the rear wall was intact. No damage occurred to the second rear wall, or to the witness plate behind the target