Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO015 Experiment Name::Free-Flyer Biostack Experiment Investigator(s)::Reitz, Dr. Gunther - Invest. Role: Present|Bucker, Horst - Invest. Role: Original Experiment Description:: Studies on the biological effectiveness of HZE particles (particles of high atomic number Z and high energy) and stars are necessary to confirm ground-based work, as well as to assess the biological effects of HZE particles not currently available from accelerators on Earth. Spaceflight experiments are required to analyze and evaluate the biological effects of the different species of HZE particles prevalent in space. In comparison with the Apollo lunar mission, the dosimetric data calculated for a 6-month flight of LDEF yields an increase in total dose of approximately 360 percent, in HZE particle fluence of approximately 200 percent, and in stars of even 2700 percent. Thus LDEF offers a unique opportunity to gather information on the effects of stars on biological matter. Objective:: The free-flyer biostack experiment is part of a radiobiological space research program that includes experiments in space as well as in accelerators on Earth. The program is specially designed to increase knowledge concerning the importance, effectiveness, and hazards of the structured components of cosmic radiation to man and to any biological specimen in space. Radiobiological and dosimetric data must be collected in space as baseline information to estimate radiation risks to man in future space missions and to establish radiation standards for man in space. Scientific and technical coordination of this program constitutes a significant portion of the experimental approach which needs to be realized within still restrictive constraints of space missions. Up to now, our understanding of the ways in which HZE particles might affect biological matter is based on a few spaceflight experiments from the last Apollo missions (Biostack I and II, Biocore, and Apollo light flash investigations) and the Apollo Soyuz Test Project (Biostack III), and in the limited data available from heavy-ion irradiation from accelerators. In the near future, accelerators capable of accelerating particles up to higher atomic numbers and higher energies will promote increased activity in ground-based studies on biological effects of HZE particles. Comparison of data from such irradiation experiments on Earth with those from an actual spaceflight experiment will show any potential influence of the inevitably attendant spaceflight factors (e.g., weightlessness) on the radiobiological events. Further, the long duration of the LDEF flight will increase the chance of studying the biological effects of the different radiations present in space, especially the effects of single heavy cosmic particles of high energy loss, the combined effects of space radiation, rare components of cosmic radiation, such as iron nuclei or superheavy particles of high energy which are not yet available from ground- based facilities, and other spaceflight factors, such as microgravity, and the actual nature and distribution of the radiation field at the surface or inside spacecraft. Experiment, Approach:: The flight hardware used to achieve this objective consists of biological specimens and nuclear track detectors. Correlation of the biological and physical events was achieved by using a special sandwich construction of visual track detectors and monolayers of biological objects. This arrangement - known as Biostack concept - allows localizing the trajectory of each heavy ion in the biological layer and identifying the site of penetration inside the biological object. The precision obtained for the reconstruction of the geometric relation between particle trajectory and test organism depends on the latter and could be pushed as low as 0.2 um for the smallest object (bacterial spores, approx. diameter 1 um) in previous Biostack experiments. These experiments comprise a wide spectrum of biological objects, such as bacterial spores (bacillus subtilis), plant seeds (arabidopsis thaliana, nicotiana tabaccum, zea mays, and rice), insect eggs, and shrimp eggs (artemia salina, one of the most primitive crustaceans). These species have different organization levels and different radiation sensitivity. They are well known and showed at least one typical genetic or somatic radiation effect. All objects were exposed in resting state and were tested to survive the period of experimental procedure. The fluence of heavy particles and/or nuclear disintegration stars depends on the locations of the experiment on LDEF. Therefore, two experiments locations with different shielding against space were used. The experiment consisted of 20 detector units, 12 units mounted in a 6-in.-deep end corner tray on the Earth-facing end of LDEF, and 8 mounted in one-third of a 6-in.- deep peripheral tray. Each unit weights approximately 2 kg and does not require power. Knowledge of the temperature history around the single units is necessary. For all other orbital parameters needed for the experiment, normal tracking of the spacecraft is sufficient. All experiment data analysis is conducted at various experimenters' laboratories in Europe and the U.S. Associated Tray(s)::Tray Location: C02 - Orientation: 141.9 degrees off ram incidence angle|Tray Location: G02 - Orientation: Earth-facing end Photograph Classification::Flight Associated Photograph(s)::LaRC #: L90-10498 KSC #: None JSC #: S32-89-029|LaRC #: L84-07057 KSC #: KSC-384C-144.08 JSC #: None|LaRC #: L91-11607 KSC #: KSC-390C-1938.09 JSC #: None|LaRC #: L90-10392 KSC #: None JSC #: S32-76-068|LaRC #: L84-07011 KSC #: KSC-384C-14.09 JSC #: None|LaRC #: L92-17835 KSC #: KSC-390C-609.09 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO019 Experiment Name::Influence of Extended Exposure in Space on Mechanical Properties of High-Toughness Graphite-Epoxy Composite Material Investigator(s)::Felbeck, Dr. David K. - Invest. Role: Present|Felbeck, Dr. David K. - Invest. Role: Original Experiment Description:: Graphite-epoxy composites are attractive for structural use in space vehicles because of their high strength and elastic modulus properties. Low fracture toughness has been solved by use of techniques of intermittent interlaminar bonding, which consists of introducing a thin perforated layer of Mylar film between adjacent plies of a cross-ply composite so as to limit the area of inter-ply bonding. In this way, fracture of the composite is diverted when crossing regions have no bonding between plies, with a consequent substantial increase in total area of fracture and an increase in fracture energy, usually with only minor reduction in strength and elastic modulus. However, before this material can be widely adapted for long-term space use, confidence must be gained that its mechanical properties are not degraded by exposure to the space environment. Objective:: The objective of this experiment was to test the effect of extended exposure to a space environment on the mechanical properties of a specially toughened T300/5208 graphite-epoxy composite material. Specimens made by techniques of intermittent interlaminar bonding will be exposed and afterward tested for fracture toughness, tensile strength, and elastic modulus. For each of the two classes of specimens, tensile/modulus and fracture toughness, the cross-play angle and the fraction (percent) of contact between adjacent plies are varied. The interlaminar contact fraction is controlled by the fraction of holes in the Mylar sheet. Experiment, Approach:: The approach of this experiment was to provide a frame on which the specimens could be mounted with their flat sides normal to the LDEF radius, each specimen with an unobstructed exposure of about 2 pi sr. The specimens were mounted so that they neither fracture from high stress nor fail from excessive heating during launch and return. Any damage to the specimens during the orbit period must be considered to be part of the experiment. Since the experiment was passive, nothing was required except the mechanical and thermal anchoring of the test specimens. There were six fracture toughness specimens and nine tensile modulus specimens utilizing one-sixth of a 3-in.-deep peripheral tray. An identical set of specimens produced at the same time were stored in the ground laboratory for final testing at the same time as the orbited specimens. A third set of specimens produced at the same time were tested a short time after fabrication and curing. These three sets of specimens were to be used to determine the effects of time plus orbit environment, time plus ground environment, and time alone on the mechanical properties of interest. After the specimens were tested to fracture, those of special interest were examined by scanning electron microscopy in order to identify any changes in fracture mode or path as a result of exposure to the space environment. Associated Tray(s)::Tray Location: D12 - Orientation: 81.9 degrees off ram incidence angle Photograph Classification::Prelaunch Associated Photograph(s)::LaRC #: L84-07093 KSC #: KSC-384C-221.11 JSC #: None|LaRC #: L90-13490 KSC #: KSC-390C-1069.10 JSC #: None|LaRC #: L90-10502 KSC #: None JSC #: S32-89-052 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO023 Experiment Name::Multiple-Foil Microabrasion Package Investigator(s)::McDonnell, Dr. J.A.M. - Invest. Role: Original|McDonnell, Dr. J.A.M. - Invest. Role: Present|Ashworth, D.G. - Invest. Role: Original|Carey, W.C. - Invest. Role: Original|Flavill, R.P. - Invest. Role: Original|Jennison, R.C - Invest. Role: Original Experiment Description:: Microfoil penetration techniques have been successfully employed as space particulate detectors since the beginning of space exploration. A number of the early Explorer satellites, the Ariel II, and three Pegasus satellites measured meteoroid penetrations in near-Earth space. These in-space penetration measurements, in addition to providing spacecraft design data, were used with existing ground-based radar and visual meteor data to extend size estimates of the near-Earth meteoroid environment to particles as small as approximately 10 ng. They offer high sensitivity of detection and yet are rugged and simple. Other early U.S. and Russian spacecraft used microphone type meteoroid detectors to measure small-particle impact fluxes. The microphone data indicated a small-particle population much greater than that indicated by the penetration measurements. The microphone data were, in fact, interpreted by some to indicate a dust belt around the Earth. Difficulties in simulating meteoroid impacts in the laboratory created a number of uncertainties in interpreting both the microphone and penetration early measurements in terms of the near-Earth meteoroid environment. Data on the near-Earth meteoroid environment have also resulted from analysis of the lunar-material samples obtained during the Apollo Program. The analysis of craters on the lunar material in terms of the meteoroid environment is limited by the facts that the craters occurred over a very long period of time (100,000 to 1,000,000 years) and the exact exposure time is uncertain. Taking advantage of the now recoverable and improved very sensitive thin- foil penetration detectors, this experiment will make a substantial step toward the elimination of a number of the remaining uncertainties in the estimates of the near-Earth micrometeroid environment. In a very cost- effective way, the experiment will provide both design data regarding the erosion of spacecraft by microparticles and data on the near-Earth micrometeroid environment. Objective:: Several objectives of the MAP experiment were: definition of the flux distribution as a function of crater size or perforation thickness, determination of the 3-dimensional flux distribution, characterization of the velocity distribution and angular distribution on a detector surface, discrimination between particle sources (e.g.: Earth orbital and interplanetary, natural or space debris if Earth-orbital, and asteroidal or cometary if natural), and particulate chemistry. The specific scientific objectives of this experiment are to measure the spatial distribution, size, velocity, radiance, and composition of microparticles in near-Earth space. The technological objectives are to measure erosion rates resulting from microparticle impacts and to evaluate thin-foil meteor bumpers. Experiment, Approach:: The Microabrasion Foil Experiment, comprised arrays of frames, each supporting two layers of closely spaced metallic foils and a back-stop plate. The detector design utilizes rolled aluminum foil down to a thickness of 1.5 um. The arrays, deploying aluminum and brass foil ranging from 1.5 to 30 um, were arranged on the North, South, East, West, and Space pointing faces of the Long Duration Exposure Facility (LDEF). The foils were bonded to meshes, in turn bonded to frames, which were bolted to a base plate occupying (for the NSEW faces) one-third of an LDEF tray; the space- point array occupies one-half tray. Bonding to etched grid supports achieves very rugged structures capable of withstanding vibrational levels and atmospheric pressure gradients typical of the LDEF-Shuttle environment. The sensitivity of foil detectors was achieved by the quality of the foil and its thickness. The experiment approach utilizes the well-established technique of thin-foil hypervelocity penetration supported by extensive investigations and calibrated in laboratory simulation to a high precision. Several different classes of impact events were anticipated to encounter such a thin-foil array. This range of classes indicates the potency of this technique compared to simple polished plates or single-foil penetrations. Deployment of such thin-foil detectors around LDEF will measure the spatial anisotropy of the impact flux. Microprobe analysis of penetration and spallation areas after recovery will determine the particle elemental composition. Combined with the identification of hypervelocity impact features by Scanning Electron Microscopy (SEM) post-flight examinations, not only can the highest confidence in a true space impact be established, but parameters of the particle such as mass or velocity can be inferred from the morphology. When, further, a second surface is placed immediately behind this foil, a capture cell is formed. Although marginally penetrating particles cannot be expected to provide ejecta detectable behind the foil, larger particles penetrate and are retained even without a significant loss. Their matter, shocked through impact, is spread out over a cone and condenses on the second surfaces; it is thus readily available for SEM and Energy Dispersive Spectroscopy. With the use of a windowless detector, light elements including carbon may be studied. The experiment was located in one-third of each of four 3-in.-deep trays located at 90 intervals around the LDEF periphery and in about two-thirds of a 3-in.-deep end corner tray on the space-facing end of the LDEF. Associated Tray(s)::Tray Location: E06 - Orientation: 98.1 degrees off ram incidence angle|Tray Location: D12 - Orientation: 81.9 degrees off ram incidence angle|Tray Location: H11 - Orientation: Space-facing end|Tray Location: C09 - Orientation: 8.1 degrees off ram incidence angle; leading edge|Tray Location: C03 - Orientation: 171.9 degrees off ram incidence angle; trailing edge Photograph Classification::Prelaunch Associated Photograph(s)::LaRC #: L84-07015 KSC #: KSC-384C-15.01 JSC #: None|LaRC #: L90-13407 KSC #: KSC-390C-1030.09 JSC #: None|LaRC #: L90-10416 KSC #: None JSC #: S32-77-065|LaRC #: L91-09090 KSC #: KSC-390C-1558.02 JSC #: None|LaRC #: L90-10451 KSC #: None JSC #: S32-78-100|LaRC #: L84-07005 KSC #: KSC-384C-14.03 JSC #: None|LaRC #: L90-13490 KSC #: KSC-390C-1069.10 JSC #: None|LaRC #: L84-07093 KSC #: KSC-384C-221.11 JSC #: None|LaRC #: L90-10502 KSC #: None JSC #: S32-89-052|LaRC #: L90-13439 KSC #: KSC-390C-1033.09 JSC #: None|LaRC #: L84-07068 KSC #: KSC-384C-193.10 JSC #: None|LaRC #: L90-10454 KSC #: None JSC #: S32-82-006|LaRC #: L91-15678 KSC #: KSC-390C-2065.07 JSC #: None|LaRC #: L92-21191 KSC #: None JSC #: S32-S-284|LaRC #: L84-06997 KSC #: KSC-384C-8.04 JSC #: None|LaRC #: L90-10336 KSC #: None JSC #: S32-75-060 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO034 Experiment Name::Atomic-Oxygen-Stimulated Outgassing Investigator(s)::Burris, Dr. Charles - Invest. Role: Present|Linton, Mr. Roger - Invest. Role: Original|Linton, Mr. Roger - Invest. Role: Present|Scott, Dr. Robert - Invest. Role: Present|Scott, Dr. Robert - Invest. Role: Original Experiment Description:: The interaction of atomic oxygen with materials in low Earth orbit (LEO) is, for many materials, a degrading factor leading to erosion, surface modification, and emission of gaseous species. Many materials (e.g., thermal control surfaces) are known to produce outgassed products and possible particulate contamination when exposed to a space environment. While these emitted gaseous products are generally low molecular weight byproducts of oxidation, concerns for the atomic oxygen stimulated emission of optically degrading outgassing products led to the development of AO034. The contamination can produce severe optical damage to the surface. It can cause an increase in surface absorption of incident radiation, thereby altering a thermal control surface, or it can cause off- axis scattering of incident radiation thereby reducing the imaging characteristics of a specular reflector. These concerns were primarily based on observations of the relative degree of contamination for thermal control coatings exposed to varying degrees of atomic oxygen on Skylab. The NASA Marshall Space Flight Center initially became involved in the contamination problem when it investigated the potential contamination problems associated with the Apollo Telescope Mount experiments. Since then, areas of interest which have been investigated are sources, effects, and abatement of contamination; restoration of surfaces; and sizing of micron-size particles using light scatter. Thermal control surface, solid-rocket plume impingements, and returned Skylab specimens have been examined. The mechanisms of contaminants and synergistic effects of the space environment are not fully understood, and the results of these investigations do not show contamination damage of the magnitude observed on the Skylab mission. Objective:: The objective of this experiment is to determine if the impingement of atomic oxygen in low Earth orbit is a major factor in producing optically damaging outgassed products. The expected results are to obtain samples which have been exposed to atomic oxygen for long durations. Analysis of these samples will determine if the impingement of atomic oxygen on the thermal control surfaces stimulates a significant amount of outgassed products. This experiment will give a clearer picture of the contamination problem and will assist in assuring that future Shuttle payloads, such as the Space Telescope and High-Energy Astronomy Observatory, will not experience Skylab contamination levels. Experiment, Approach:: LDEF experiment AO034 was designed to provide evidence of atomic oxygen interaction with selected thermal control coatings leading to the stimulation, emission, and deposition of optically degrading outgassing byproducts on adjacent optical collector mirrors. Selected thermal control surfaces were exposed to the atomic oxygen and the combined space environment in the near-Earth orbit. Passive collecting samples collected any induced outgassing resulting from the oxygen impingement. The optical condition of the passive samples was measured using a ground-based integrating sphere reflectometer and a directional reflectometer. Two packages were required, each occupying identical one-sixth 3-in.-deep tray modules. One package was positioned on the leading edge (Row 9) and, for reference, one on the trailing (Row 3), wake, edge of the LDEF. Duplicate specimens in each module were mounted under windows and opaque covers to provide some measure of effects of individual environmental factors. The thermal control surfaces used on Skylab, as well as newly developed surfaces, were contained in the packages. Optical mirrors were adjacent to the thermal coatings for deposition of outgassing products. The atomic oxygen impinged on the thermal control surfaces contained in the leading-edge container. Passive specular collecting samples were positioned to collect any condensable outgassed products produced as a result of the oxygen impingement. The thermal control surfaces, as well as the collecting samples, were exposed to the available ultraviolet radiation so that environment synergistic effects could be observed. Ultraviolet grade windows and metal covers provide for additional assessment of the effects of the various environmental factors. Preexposure and postexposure analysis include total hemispherical and bidirectional reflectance measurements. The range of these reflectance measurements are 2.5 to 2500 angstroms. Postexposure analysis of the leading- edge samples exposed to the atomic oxygen is compared with the control samples positioned on the trailing edge and shielded from the atomic oxygen. Associated Tray(s)::Tray Location: C09 - Orientation: 8.1 degrees off ram incidence angle; leading edge|Tray Location: C03 - Orientation: 171.9 degrees off ram incidence angle; trailing edge Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L91-09090 KSC #: KSC-390C-1558.02 JSC #: None|LaRC #: L90-10416 KSC #: None JSC #: S32-77-065|LaRC #: L84-07015 KSC #: KSC-384C-15.01 JSC #: None|LaRC #: L90-10451 KSC #: None JSC #: S32-78-100|LaRC #: L90-13407 KSC #: KSC-390C-1030.09 JSC #: None|LaRC #: L84-07005 KSC #: KSC-384C-14.03 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO044 Experiment Name::Holographic Data Storage Crystals for LDEF Investigator(s)::Callen, Dr. Russell - Invest. Role: Original|Callen, Dr. Russell - Invest. Role: Present|Gaylord, Dr. Thomas - Invest. Role: Present|Gaylord, Dr. Thomas - Invest. Role: Original Experiment Description:: Lithium niobate is a significant electro-optic material, with potential applications in ultra- high capacity data storage and processing systems. Lithium niobate is the material of choice for many integrated optical devices and holographic mass memory systems. A compact high-bit-capacity recorder (on the order of 100,000,000,000 bits) and memory system is sought. Electro-optic holographic recording systems are being developed and appear extremely promising. Objective:: The objective of this experiment is to test the spaceworthiness of electro-optic crystals for use in ultrahigh-capacity space data storage and retrieval systems. Experiment, Approach:: To record the volume holograms in the lithium niobate, two plane waves, produced by the same laser, are interfered within the crystal. The interference maxima and minima produce a corresponding refractive index variation in the photorefractive crystal, which produces the volume hologram. By passing one of these beams, the object beam through a data page mask prior to incidence on the crystal, digital data in page oriented format may be stored as a volume hologram. The data page may be displayed by subsequent illumination with the reference beam alone. By rotation of the crystal with respect to the laser beam, multiple pages of data can be stored. The experiment approach was to passively expose four holographic data storage crystals, each 10 by 10 by 2 mm in size and with the optic axis lying in the plane of the surface, to the space environment. Three of the iron-dope lithium niobate crystals contained recorded holograms and one is unrecorded (control sample). The crystals were specified to be iron-doped to 0.005 mole percent iron in the melt. Crystal 1 is heat treated for maximum sensitivity and is blank. Crystal 2 contains a plane wave hologram written with a helium-neon laser (lambda = 632.8 nm), Crystal 3 contains a plane waves hologram written with an argon laser (lambda = 514.5 nm), and crystal 4 contains a spoke pattern hologram written with an argon laser (lambda = 514.5 nm). The crystals containing the holograms were fixed in an atmosphere of lithium carbonate to extend the lifetime of the holograms. This spectrum of crystals assures determination of the most suitable crystal treatment for space use. A glass control sample was also be flown. In crystals 2, 3 and 4, the data was protected by charge neutrality of combined ion and electron patterns, and the holograms should be directly recoverable upon reexposure to uniform illumination. Two control crystals remained on the ground, one containing a helium-neon laser hologram and one containing an argon laser hologram. The crystals for this experiment were included with the various electro- optical components of LDEF experiment S0050, Investigation of the Effects of Long-Duration Exposure on Active Optical System Components, and were located in the same experiment tray. Associated Tray(s)::Tray Location: E05 - Orientation: 128.1 degrees off ram incidence angle Photograph Classification::Prelaunch Associated Photograph(s)::LaRC #: L89-04387 KSC #: KSC-384C-255.06 JSC #: None|LaRC #: L91-10508 KSC #: KSC-390C-2114.03 JSC #: None|LaRC #: L90-13450 KSC #: KSC-390C-1035.05 JSC #: None|LaRC #: L90-10413 KSC #: None JSC #: S32-77-050 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO038 Experiment Name::Interstellar Gas Experiment Investigator(s)::Geiss, Johannes - Invest. Role: Original|Lind, Dr. Don L. - Invest. Role: Present|Buhler, Dr. Fritz - Invest. Role: Original|Buhler, Dr. Fritz - Invest. Role: Present|Lind, Dr. Don L. - Invest. Role: Original Experiment Description:: In the vicinity of the solar system, interstellar particles are mostly individual neutral atoms. Because of their motion relative to the sun, a portion of this flux can penetrate into the solar system as far as the region of the inner planets. The presence of these particles near the Earth was first confirmed by the OGO-5 spacecraft in 1969 and they have provided us with considerable insight as to the nature of the nearby interstellar medium. In the past, the observed regularities in the abundance of elements and their isotopes, upon which the theory of nucleosynthesis rests, have been obtained primarily from solar system abundances, in particular meteoritic, solar, terrestrial, and solarwind data. However, this sample represents only a tiny fraction of the material of the universe. Thus, even a small sample of extra-solar-system isotopes will give significant insight into the various element building processes that have occurred in the original nucleosynthesis and those which have occurred in our galaxy. Isotopic analysis of the noble-gas component of the interstellar gas will provide a significant new data source and will complement other promising techniques, such as millimeter wave, cosmic-ray, and nuclear-gamma-ray astronomy. The thin foil detection technique for these interstellar particles which we propose for the Interstellar Gas Experiment (IGE) was first employed on the Apollo missions to the moon to measure the isotopic ratios of the solar wind. Later, on the Skylab mission this same technique was used to measure the isotopes of precipitating magnetospheric particles. The technique has also been utilized on a sounding rocket to analyze auroral particles. Thus, considerable experience has been accumulated in this method of collecting extraterrestrial particle samples. Objective:: The primary objective of this experiment was to collect interstellar gas atoms in metal foils at several locations around the Earth's orbit. These particles arrive in the vicinity of the Earth as the neutral interstellar wind penetrates the heliosphere and enters the region of the inner planets. The flow pattern of the interstellar wind is controlled mainly by the gravitational attraction of the Sun, and its density is reduced through ionization by solar photons and by charge exchange with the solar-wind particles. The flux of the interstellar atoms that survive the upstream journey in the solar wind is increased due to gravitational focusing as they pass beyond the Sun. The angular distribution of these particles is also significantly modified in the gravitational focusing process. Thus the density and angular distribution of the interstellar gas flux vary considerably at different points about the Earth's orbit. In addition to these variations, the velocity of these particles as they impact the spacecraft changes over a wide range as the orbital motion of the Earth moves seasonally first upstream, then cross stream, and finally downstream in the interstellar wind. These seasonal variations constitute the signature by which the interstellar particles can be identified. By collecting these particles at several locations in the Earth's orbit, it is possible not only to achieve an in situ detection of the interstellar gas for the first time, but also to study the dynamics of the interstellar wind as it flows through the heliosphere and interacts with the solar photon flux and the solar wind. In addition, because the dynamics of the interstellar wind depend on its density and velocity before entering the heliosphere, it is possible to investigate these characteristics of the inter- stellar medium outside the region of the solar system. The objectives of this experiment were to collect and isotopically analyze interstellar gas atoms around the orbit of the Earth for the purpose of obtaining new data relevant to understanding nucleosynthesis, and to study the dynamics of the interstellar wind inside the heliosphere and the isotopic composition of the interstellar medium outside the heliosphere. Another primary goal of the experiment was to detect and, if possible, to isotopically analyze the noble gas component of the local interstellar medium and to investigate the noble gas isotopic ratios of this interstellar sample of matter which originated outside the solar system. Experiment, Approach:: The Interstellar Gas Experiment (IGE) exposed thin metallic foils aboard the LDEF spacecraft in low Earth orbit (LEO) in order to collect neutral interstellar gas particles which penetrate the solar system due to their motion relative to the sun. The experiment hardware acted as a set of simple cameras with high-purity 15 micron thick copper-beryllium collecting foils with a beryllium-oxide surface layer serving as the film. In the IGE, the foils were located at the bottom of a collector - a rectangular box which established the field of view for the foil and the orientation of this field of view on the celestial sphere. IGE consisted of seven such collectors, each viewing a different direction relative to the LDEF spacecraft. The experiment housing was to: mount and thermally control the foils, establish the viewing angles and viewing direction, provide baffling to reject ambient neutral particles, provide a voltage grid to reject ionospheric charged particles, sequence collecting foils, control exposure times, and protect the foils from contamination during the deployment and retrieval of the LDEF. By mechanical penetration the atoms are embedded in the collecting foils along with precipitating magnetospheric ions and, possibly, with ambient atmospheric atoms. The impact velocities of the particles are sufficient to embed them into the surface of the collecting foils. Since the collected particle sample is extremely minute, special mass spectrometer techniques are required for their measurement. After being returned to Earth, the entrapped particle were to be liberated by heating the foils. The gases are passed through a chemical getter. This removes all chemical elements except the noble gases. These atoms can be analyzed by mass spectroscopy to determine the relative abundance of the different isotopes of helium and neon. An attempt was planned to detect argon. At present, the noble gases are the only species for which this method is sufficiently sensitive. In the analysis of these particles, not only can the amounts of the various noble gas isotopes be measured, but additional information can be obtained by heating the collecting foils in increments. At the first relatively low temperature (450 degrees), the least tightly bound particles are released. At higher temperature steps, the particles which had penetrated farther into the foil are released. Thus we can determine a rough approximation of the impact velocity for the various portions of the collected sample. The experiment used four trays, two 12-in.-deep peripheral trays and two 12-in.-deep end center trays, on the space-facing end of the LDEF. One of the peripheral trays contained only one camera and the rest of the trays contained two cameras. Power requirements were supplied by LiSO2 batteries. Associated Tray(s)::Tray Location: H06 - Orientation: Space-facing end|Tray Location: E12 - Orientation: 81.9 degrees off ram incidence angle|Tray Location: H09 - Orientation: Space-facing end|Tray Location: F06 - Orientation: 98.1 degrees off ram incidence angle Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L90-13488 KSC #: KSC-390C-1069.09 JSC #: None|LaRC #: L90-10431 KSC #: None JSC #: S32-78-031|LaRC #: L89-04418 KSC #: KSC-384C-538.06 JSC #: None|LaRC #: L89-04424 KSC #: KSC-384C-538.12 JSC #: None|LaRC #: L90-10453 KSC #: None JSC #: S32-82-002|LaRC #: L90-13438 KSC #: KSC-390C-1033.08 JSC #: None|LaRC #: L89-04416 KSC #: KSC-384C-538.04 JSC #: None|LaRC #: L90-10369 KSC #: None JSC #: S32-75-063|LaRC #: L92-21197 KSC #: None JSC #: S32-S-290|LaRC #: L89-04420 KSC #: KSC-384C-538.08 JSC #: None|LaRC #: L91-11828 KSC #: KSC-390C-1641.03 JSC #: None|LaRC #: L91-11881 KSC #: KSC-390C-1636.10 JSC #: None|LaRC #: L90-10367 KSC #: None JSC #: S32-75-061|LaRC #: L92-21193 KSC #: None JSC #: S32-S-286 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO054 Experiment Name::Space Plasma High-Voltage Drainage Experiment Investigator(s)::Yaung, Dr. Jim - Invest. Role: Present|Komatsu, G. - Invest. Role: Original|Taylor, W.L. - Invest. Role: Original Experiment Description:: Thin dielectric films are frequently employed as the coating materials for solar arrays and in thermal control applications. These films are subject to electric stress as a result of either voluntary or involuntary actions. For solararray applications, the presence of array voltages causes electric stress across the dielectric film and to the space plasma, with resultant current drainage from the plasma to the array cells. As array voltages are raised, electric stress and current drainage levels also rise and may impact array operation and efficiency. For both thermal control coating materials and array coating materials on spacecraft immersed in energetic particle environments in space, involuntary charge buildup occurs and results in both transient and steady-state current drainages that may impact spacecraft operation. TRW's SP-HVDE on LDEF provides the first in-flight investigation of current through biased dielectric films. Objective:: The experiment was designed to investigate the long term leakage of dielectric materials under high voltage bias. The objectives of this experiment were to place large numbers of dielectric samples under electric stress and exposure to the LEO space environment; to determine the in-space current drainage behavior of positively and negatively charged dielectrics; to recover, inspect, and further test these samples in laboratory facilities; and finally to specify allowable electric stress levels for these materials as applied to solar-array and thermal control coatings for prolonged exposure in space. Other benefits such as material degradations caused by the LEO environment and functional degradation of high voltage systems can be investigated. These findings, in turn, will pace the design of encapsulated, lightweight, high-voltage solar arrays as well as the development of coating materials for spacecraft operation in energetic charged-particle environments such as that experienced at geosynchronous altitudes during magnetic substorms. Experiment, Approach:: The Space Plasma - High Voltage Drainage Experiment (SP-HVDE) provided a unique opportunity to study long term space environment effects on materials because it was comprised of two identical 3-in.-deep peripheral experimental trays; one tray located on the ram-facing side (D10), and the other on the wake-facing side (B04) of the Long Duration Exposure Facility (LDEF). This configuration allowed for the comparison of identical materials exposed to two distinctly different environments and the determination of charged-particle drainage as a function of plasma density. The plasma environment contributes the major source of the leakage currents. The cosmic rays cause much less impact to the experiment since they tend to penetrate the samples. The electromagnetic radiation environment is also of minor interest, for it induces insignificant leakage current through photoabsorption process in our experimental samples. The drainage current behavior of thin dielectric (insulating) films in space is determined by placing the forward (exposed) face of the film in contact with the space plasma while applying a bias voltage to a conducting layer on the rear (nonexposed) face. Current flow from the rear-face conduction film through the dielectric to the charged-particle environment of space occurs as a result of the bias potential. The completion of the current loop occurs when charged particles are collected from the space plasma by the frame of the LDEF and then delivered to the ground return of the bias voltage supply. The bias potential is developed by a self-contained battery and power processor unit. Each dielectric sample has an associated battery and power processing unit, except for the spectator samples, which were not electrically stressed in flight and hence allowed a determination of the effects of merely being present on the LDEF. The dielectric sample power processor was equipped with two coulometers. These Plessey coulometers record currents flowing in either direction by either deplating or plating the electrode. The coulometer was used as either a leakage current time integrator or a bias voltage time integrator. The first of these was in series with the bias voltage lead and determined the integral of the drainage current during the flight. The second coulometer, which had a high-value in-series resistor, was placed between the bias potential and LDEF ground to determine the time-integrated applied-bias voltage. An average front-to-back resistance of the dielectric sample was determined from the measured time integrals of drainage current and bias potential, and the bulk resistivity of the dielectric material under applied stress and in the space environment was determined from known surface area and film thickness. If this bulk resistivity remains greater than certain minimal limits for a given bias potential, then use of the material for high-voltage solar arrays would be permitted for this series of specified environmental and electrical conditions. Deterioration of dielectric properties under continued stress would prohibit use in the high-voltage arrays and would present significant long-term equilibration data for spacecraft coating materials subjected involuntarily to charge and voltage buildup because of energetic charged-particle deposition. Associated Tray(s)::Tray Location: D10 - Orientation: 21.9 degrees off ram incidence angle|Tray Location: B04 - Orientation: 158.1 degrees off ram incidence angle Photograph Classification::Flight Associated Photograph(s)::LaRC #: L90-10407 KSC #: None JSC #: S32-77-027|LaRC #: L92-17647 KSC #: KSC-390C-730.08 JSC #: None|LaRC #: L84-07086 KSC #: KSC-384C-221.04 JSC #: None|LaRC #: L90-10439 KSC #: None JSC #: S32-78-062|LaRC #: L90-13393 KSC #: KSC-390C-1029.07 JSC #: None|LaRC #: L84-07065 KSC #: KSC-384C-193.07 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO056 Experiment Name::Exposure to Space Radiation of High-Performance Infrared Multilayer Filters and Materials Technology Experiments Investigator(s)::Lipscombe, Derek - Invest. Role: Original|Whatley, A. - Invest. Role: Original|Hawkins, Dr. Gary J. - Invest. Role: Present|Hunneman, Dr. Roger - Invest. Role: Original|Hunneman, Dr. Roger - Invest. Role: Present|Seeley, Dr. John - Invest. Role: Original|Seeley, Dr. John - Invest. Role: Present Experiment Description:: Infrared multilayer interface filters have been used extensively in satellite radiometers for several decades. Filters manufactured by the University of Reading have been used in Nimbus 5, 6, and 7, TIROS N, and the Pioneer Venus orbiter. Filters for those types of spaceflight radiometer instruments measure IR emissions from the major atmospheric gases, pollutants, and aerosols, and from the sea to land surface. Progress in the general science of radiometry from space has been rapid, having evolved to the extent that global change is considered observable; currently, radiometers are being planned to fly in the EOS/ERS/POEM satellites capable of monitoring parameters such as structure, atmosphere dynamics and composition in situ. Further, EOS accuracy is intended to be sufficiently good for valid deduction of temperature gradient, mixing/transport of the gases, surface flux, etc. As with the earlier radiometers, spectral bands are defined by optical filters for the selection of the particular molecular species of interest. The spectral width and resolution of filters, together with the stability and durability of their materials and processes, therefore remains especially significant. The ability of the filters to withstand the space environment in these applications is critical; if degradation takes place, the effects would range from worsening of signal-to-noise performance to complete system failure. This LDEF experiment enabled the filters, for the first time, to be subjected to authoritative spectral measurements following space exposure to ascertain their suitability for spacecraft use and to permit an understanding of degradation mechanisms. The University of Reading experiment was intended to expose infrared multilayer interference filters of novel design, high performance, construction and manufacture to space radiation. The spectral behavior of these filters when in space was unconfirmed previously, but crucial to filter performance. High-performance, in this context, relates to the precision and discrimination with which the emission spectra of important atmo- spheric gases can be resolved. Additionally, the understanding of the effects of prolonged space exposure on spacecraft materials, surface finishes, and adhesive systems is of great interest to the spacecraft designer. Thus, a series of materials technology experiments were included with the experiment on infrared multilayer filters. Objective:: The objective of the multilayer filters experiment was to expose high-performance infrared multilayer filters to the space environment and recover them for subsequent analysis and comparison with laboratory control samples. Semiconductors such as PbTe, Si, and Ge were examined to see if excess free carriers were generated by exposure, and for evidence of surface contamination or degradation and/or decomposition. ZnS and other dielectrics were examined for evidence of bulk degradation, such as enhanced absorption, color center excess, and Reststrahl abnormalities. The objectives of the materials technology experiments were to evaluate the degradation of spacecraft surface finishes, the outgassing of spacecraft surface finishes, the effect of thermal paints on carbon-fiber reinforced plastic (CFRP) sheet and the thermal differentiation of expansion between base material and thermal coating, the strength of adhesive-bonded joints (lap shear), the effect of CFRP strength stiffness and interlaminar strength, the dimensions of CFRP curvature, and the effect on bond strength between CFRP-aluminum alloy skins and honeycomb core. Experiment, Approach:: The experiment utilized one-sixth of a 3-in.deep peripheral tray and one-fourth of a 3-in.-deep end center tray on the Earth-facing end of the LDEF. The experiment comprised 46 individual components divided between an Earth facing tray (G12) and leading edge tray (B08). Each was designed to maximize the exposure at full aperture to the complete range of space radiations and temperature excursions received by LDEF. Each of the components was housed in aluminum holders designed for mechanical stability and thermal contact, to ensure uniform temperature across the exposed aperture. Samples were retained in their holders with disc springs and circlips located behind a chromic-anodized BS.L93 aluminum backing piece. Thermal contact was ensured by Pb washers located on either side of the substrate. The holders were retained in the base plate using a disc spring and circlip located in a recess on the outside of the holder. A small pressure release hole was located in the center of each backing piece to minimize the pressure differential across the substrate and to prevent flexing or deformation. The samples were subdivided into three categories: uncoated crystals and materials, soft-multilayer coatings and substrate materials, and hard-multilayer coatings and substrate materials. Equivalent samples were placed in both experiment trays to assess performance differences between the two LDEF exposure sites. The following uncoated crystals and materials were selected for their long-wave (Reststrahl) absorption properties, and as typical substrate materials: calcium fluoride, magnesium fluoride, germanium, silicon, cadmium telluride, sapphire, Y-cut quartz, Z-cut quartz, KRS-5, KRS-6. Soft materials comprised principally KRS-5 (TlBrI)-based multilayers deposited on KRS-5 or KRS-6 (TlClBr) substrates. These were designed to utilize, by multilayer interference, longwavelength Reststrahl blocking properties. All of them were originally fabricated for the GALILEO Jupiter probe and exposed with an aperture diameter of 13.0 mm. Hard materials comprised spectral filters from atmospheric-sensing, weather forecasting, research, and planetary satellites, viz: NIMBUS 4, 5, 6, 7, ITOS, TIROS N, PIONEER, and GALILEO; each was exposed with an aperture diameter of 21.0 mm. Three material combinations were employed in the low-contrast quarter-wave blocking stacks, viz: KRS-5 with a choice of As2S3 / CdTe / ZnS or ZnSe as alternate layer materials. Each combination was highly transparent through the NIR/IR, showing periodic fringes, but ending at longwavelength (approx. 40 um) in a deep Reststrahl absorption. Eight generic designs were selected for LDEF. They represented a cross-section of the different types used in radiometer instrumentation: Low index-ratio quarter-wave stack (for blocking - 6 off), Single spacer bandpass (4 off), Double half-wave narrow band design (3 off), Blocked and unblocked triple half-wave bandpass designs (8 off), Unblocked split-spacer Fabry-Perot band- pass designs (6 off), Twin-peaked bandpass design (1 off), Optimized Tschebychev low pass edge filter extraction (1 off), and single layer, single wavelength antireflection coatings (2 off). The samples were measured on an infrared spectrometer with particular reference to any critical parts of their spectrum (e.g., peak transmission and center wave number of bandpass filters, edge position and steepness in edge filter, and transmission in longwave filters) prior to assembly into the experimental structure. At the same time, the samples were visually inspected and photo- graphed, and possible other testing (for example, adhesion tests) was undertaken. Upon retrieval, the samples were visually inspected prior to shipment back to the laboratory for postflight tests. Associated Tray(s)::Tray Location: G12 - Orientation: Earth-facing end|Tray Location: B08 - Orientation: 38.1 degrees off ram incidence angle Photograph Classification::Prelaunch Associated Photograph(s)::LaRC #: L84-07116 KSC #: KSC-384C-294.07 JSC #: None|LaRC #: L90-13421 KSC #: KSC-390C-1031.11 JSC #: None|LaRC #: L90-10380 KSC #: None JSC #: S32-76-026|LaRC #: L90-10388 KSC #: None JSC #: S32-76-054|LaRC #: L84-07165 KSC #: KSC-384C-317.08 JSC #: None|LaRC #: L91-11718 KSC #: KSC-390C-1997.02 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO076 Experiment Name::Cascade Variable-Conductance Heat Pipe Investigator(s)::Grote, Mr. Michael - Invest. Role: Original|Grote, Mr. Michael - Invest. Role: Present|Calhoun, Leslie - Invest. Role: Original Experiment Description:: A number of spacecraft applications could benefit from a precise temperature control system which requires zero electrical power. The dry-reservoir variable-conductance heat pipe (VCHP) system could provide this capability, for long term spaceflight. Objective:: The objective of this experiment was to verify the capability of a cascade variable-conductance heat pipe (CVCHP) system to provide precise temperature control of long-life spacecraft without the need for a feedback heater or other power sources for temperature adjustment under conditions of widely varying power input and ambient environment. Experiment, Approach:: The experiment approach was consistent with the LDEF capabilities (i.e., relatively long duration, zero gravity environment, and minimal electrical power and data system capability). Solar energy was the heat source and space the heat sink for thermally loading two series-connected variable-conductance heat pipes. Electronics and power supply equipment requirements were minimal. The experiment power data system (EPDS) in LDEF experiment S1001 (Low-Temperature Heat Pipe Experiment Package (HEPP) for LDEF) was used for data recording. A 7.5-V lithium battery sup- plied the power for thermistor-type temperature sensors for monitoring system performance, and a 28-V lithium battery supplies power for valve actuation. Two external-surface subpanels were employed which were thermally coupled to opposite ends of two series-connected variable-conductance heat pipes. One panel was designed as a heat absorber through application of a high absorptivity/emissivity surface coating, and the second was a radiator with a low absorptivity/emissivity surface coating. Multilayer insulation and fiberglass structural attachments were used to thermally isolate the experiment from the LDEF tray structure and interior. Each of the heat pipe evaporators was maintained within preselected temperature ranges by sizing the collector, radiator, and insulation and by servicing the noncondensible dry gas reservoirs, thus demonstrating passive variable-conductance heat pipe operation. The CVCHP experiment configuration used two gas-loaded, dry-reservoir VCHP's in series. The coarse-control heat pipe temperature was controlled to plus or minus 3 degrees C and was used as a sink for the fine-control heat pipe. The dry gas reservoir temperatures were controlled by locating the reservoirs next to the heat pipe evaporators. The principal concern was drift in the set point temperature due to many heat load and/or environment temperature cycles. The cyclic operation moved vapor into or out of the noncondensible gas reservoir, which changed the set point temperature by introducing a varying working-fluid partial pressure into the gas reservoir. The capillary tube was located between the heat pipe and the reservoir to prevent working fluid from entering the reservoir. The diameter and length of the capillary were selected to satisfy two criteria. First, the capillary must provide sufficient volume to accommodate the entire volume of gas displaced when the vapor front moves from its minimum to its maximum position. Second, the capillary length must be sufficient to prevent diffusion into the reservoir for the mission lifetime. The heat pipe wick design had a single-pedestal artery with seven tubelets enclosed in a sheath of two layers of 400-mesh screen. The sheath provided the high capillary pumping pressure and the tubelets provide high permeability to reduce the axial pressure drop. The wall wick has a 200-mesh out layer for low pressure drop and a middle layer of 400 mesh for capillary pumping. Ammonia was used as the working fluid for both heat pipes. The coarse-control VCHP had a reservoir-to-condensor volume ratio of 20, which yields a control band of plus or minus 3 degrees C. A larger 90-to 1 volume ratio was used on the fine-control heat pipe to attain a control of plus or minus 0.3 degrees C. The fine-control VCHP has 8.5 m of capillary, and the course-control VCHP had a larger 13.4-m section because of its larger condenser volume. Experiment operation began when the battery circuit was initiated at LDEF deployment from the Shuttle. When the temperature of the fine-coarse heat exchanger fell below -7 degrees C, a valve opened to permit initiation of the coarse-control VCHP and allow the heat pipes to prime prior to beginning VCHP operation. This temperature-controlled initiation ensured that heat pipe temperatures were below their operating set point so that vapor was not forced into the reservoirs. Twenty-five hours later, another valve opens to initiate the fine-control VCHP operation. This delayed opening permitted data to be taken on a stabilized coarse-control VCHP before fine-control VCHP operation. Data were collected at least twice daily during the LDEF mission and were stored on magnetic tape for subsequent retrieval and playback. Associated Tray(s)::Tray Location: F09 - Orientation: 8.1 degrees off ram incidence angle; leading edge Photograph Classification::Prelaunch Associated Photograph(s)::LaRC #: L84-07310 KSC #: KSC-383C-4418.02 JSC #: None|LaRC #: L90-10448 KSC #: None JSC #: S32-78-087|LaRC #: L90-01826 KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO114 Experiment Name::Interaction of Atomic Oxygen With Solid Surfaces at Orbital Altitudes Investigator(s)::Gregory, Dr. John C. - Invest. Role: Original|Gregory, Dr. John C. - Invest. Role: Present|Peters, Dr. Palmer - Invest. Role: Original|Peters, Dr. Palmer - Invest. Role: Present Experiment Description:: Atomic oxygen and nitrogen are known to be extremely reactive when impinging on solid surfaces. Chemical changes can occur which alter optical and electrical properties and in some cases even remove layers of material. If the atoms impinge with the kinetic energy of orbital velocity (approximately 5 eV for atomic oxygen), the possibility of physical sputtering exists. There is, however, no experimental evidence for this because laboratory beams of sufficient flux at these energies are extremely difficult to produce. The mechanisms for these interactions are not entirely. This experiment was designed to expose a wide variety of surfaces to the intense atom flux in orbit in order to determine the gross nature of the effects. As a platform for this type of experiment, the LDEF is particularly well suited compared with other spacecraft designed to point or remain fixed in space. The attitude stabilization mode of LDEF resulted in the same surface always being presented to the ambient atmospheric flux along the velocity factor, while the opposite surface of the vehicle remained in a hard vacuum. Images produced by pinhole cameras using film sensitive to atomic oxygen provide information on the ratio of spacecraft orbital velocity to the most probable thermal speed of oxygen atoms. Alternatively, information on the spacecraft attitude relative to the orbital velocity can be obtained, provided that corrections are properly made for thermal spreading and a co-rotating atmosphere. Objective:: The objectives of this experiment were to advance the knowledge of atom-surface interactions in the experimentally difficult energy range near 5 eV, to enable surface experiments to be designed for future Shuttle-era programs with greater chance of success, and to provide engineers and scientists in other areas with foreknowledge of the effects of the oxygen atom beam on critical surfaces. Experiment, Approach:: The basic approach to this experiment was to expose a wide variety of material surfaces to the atomic flux in orbit. The experiment was passive and depended on preflight and postflight measurements of the test surfaces in the laboratory. The experiment also included a reflectometer device to measure atomic beam reflection angles and thus momentum accommodations, and a unique passive spacecraft attitude sensor. Exposure of Samples: Samples consisting of solid disks or thin film coatings on substrate disks were mounted in a panel. The face of this panel was flown on LDEF normal to the incident stream of oxygen atoms. Each disk had part of its front surface masked so exposure to the atomic-oxygen reaction was limited to selected areas, the shadowed areas being used as control surfaces in the measurements. By masking half of each sample, it is possible to measure changes in roughness, erosion depths, materials growth and changes in film thickness. A typical sample was an optically flat quartz disk overcoated with a film of the material of interest. These included Ag, Au, Pt, Nb, Ni, Al, C, Si, Ge, LiF, and a few engineering materials. Some materials for which the expected removal rate is high, such as carbon, were solid disks rather than thin films. The experiment consisted of two flight units. Each unit occupied one-sixth of a 3-in.-deep peripheral tray with one unit located on the leading edge of LDEF and the other unit on the trailing edge. The samples on the trailing-edge unit were not subjected to atomic-oxygen impingement and served as control samples. A third set of control samples were kept in the laboratory to aid in postflight analysis. To estimate the effects of contamination encountered during ascent, deployment, and descent of the Shuttle, a few samples were also contained in an experiment exposure control canister (EECC) located with LDEF experiment S0010, Exposure of Spacecraft Coatings. The effect of atomic oxygen on materials is highly variable. No method of measuring the surface is optimum for all materials. Postflight measurement techniques include step-height measurement by interferometry and surface profilometry, optical densitometry, electrical resistivity, and depth profile of chemical composition by Auger electron spectroscopy. One of the most effective techniques has been to utilize the measurement of etched step at interfaces between exposed and unexposed, or masked, areas by stylus profilometry. Stylus profilometers typically measure the vertical displacement of a stylus (usually a fine pointed diamond) as it is scanned horizontally across the surface. Highly magnified vertical displacements are plotted against horizontal positions greatly exaggerating surface detail. The technique has the ability to measure a wide range of etch steps, from below 1 nm to 1 mm. Reflected Atoms: Angles of reflection of the hyperthermal oxygen atom beam are related to the extent of momentum accommodation, of which little is known at these energies. A strong forward lobe in the distribution of atomic oxygen would be detectable by sensor surfaces arrayed in the reflected beam, rather like the film in a cylindrical X-ray diffraction camera. To examine the momentum accom- modation aspect, three cylindrical reflectometers were also included with this experiment. Slits in the panels and cylinders permit a beam of atomic oxygen to impinge on a given sample at a selected angle of incidence. For the first flight of the reflectometers, the objects were limited to testing the concept of the silver film detectors and distinguishing between specular and cosine law reflection at the same surface. Materials chosen included LiF, stainless steel, and aluminum. Spacecraft Attitude Sensor: A unique passive spacecraft attitude sensor was incorporated into each unit of this experiment to serve as a means of determining the orientation of LDEF with respect to its velocity vector. The sensor was designed to measure the angular offset of LDEF from its nominal flight attitude. The atomic oxygen sensitive pinhole camera used the fact that oxygen atoms dominate the atmosphere in low Earth orbits, and formation of a nearly collimated beam of oxygen atoms passing through a pinhole in a satellite front surface occurs as a result of the orbital velocity being greater than the most probable Maxwell-Boltzmann speed of the oxygen atoms. Thus, the range of incidence angles of atoms to satellite surfaces is very limited. Any variation in the attitude of the LDEF's velocity vector relative to the atmosphere would cause the spot to wander, producing a nonspherical, larger than normal, spot compared to that produced by thermal spreading of the beam. A thin film of material (silver in this case), which is sensitive to atomic oxygen, then forms an image on the impact spot. Silver was used because it discolors from formation of oxide. The pinhole camera consisted of a 0.3 mm thick stainless steel hemisphere 3.25 cm radius, polished on the concave surface and coated with vacuum-evaporated silver. The pinhole was positioned at the center of the silvered hemisphere. As a subexperiment provided by Dr. Gerald J. Fishman of NASA Marshall Space Flight Center, a number of activation metal samples were included with the other samples previously mentioned. After exposure to the space environment, these samples become slightly more radioactive due to ambient proton and neutron irradiation. Upon recovery, the radioactivity was carefully analyzed by the NASA MSFC Space Science Laboratory to provide measurements of the average proton and neutron fluence during the LDEF mission. Associated Tray(s)::Tray Location: C09 - Orientation: 8.1 degrees off ram incidence angle; leading edge|Tray Location: C03 - Orientation: 171.9 degrees off ram incidence angle; trailing edge Photograph Classification::Flight Associated Photograph(s)::LaRC #: L90-10416 KSC #: None JSC #: S32-77-065|LaRC #: L91-09090 KSC #: KSC-390C-1558.02 JSC #: None|LaRC #: L84-07015 KSC #: KSC-384C-15.01 JSC #: None|LaRC #: L90-13407 KSC #: KSC-390C-1030.09 JSC #: None|LaRC #: L84-07005 KSC #: KSC-384C-14.03 JSC #: None|LaRC #: L90-10451 KSC #: None JSC #: S32-78-100 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO133 Experiment Name::Effect of Space Environment on Space-Based Radar Phased-Array Antenna Investigator(s)::Whiteside, Mr. James - Invest. Role: Original|Whiteside, Mr. James - Invest. Role: Present|DeIasi, Richard - Invest. Role: Original|Heuer, Ronald - Invest. Role: Original|Kesselman, Martin - Invest. Role: Original|Kuehne, Fredrick J. - Invest. Role: Original|Rossi, Martin - Invest. Role: Original Experiment Description:: Large space structures of low areal density are currently being developed for near-term applications such as space-based radar (SBR) and the Earth Observation System (EOS). The practical implementation of these structures depends largely on identifying low-cost, low-density, high-strength-to-weight materials that are not degraded by the low Earth orbit (LEO) and geosynchronous Earth orbit (GEO) environments. Because of the necessity for low weight and density, candidate materials include polymeric materials. Polymeric materials satisfy many of these requirements; however, the long-term stability of these materials exposed to the space environment is a major concern. The nature of the chemical bonds causes these materials to be susceptible to some degree of degradation from either ultraviolet or charged-particle (particularly high-energy electron) components of the space environment. In addition, for materials required to retain stiffness and dimensional stability, thermal excursions become an important factor because of creep at elevated temperatures. Atomic oxygen erosion is not a major concern in the SBR application environment, but it certainly becomes an issue at the lower altitudes flown by LDEF. Based on the performance of numerous polymeric materials following accelerated laboratory testing, Kapton polyimide film was selected as the baseline material for the Grumman SBR concept. To gain the requisite confidence for long-term service durability, it was desirable to subject material specimens as well as a portion of the SBR antenna directly to the combined space environ- ment and compare property degradation to that caused by laboratory simulation. Objective:: The overall objective of this program was to evaluate the effect of the space environment on polymeric materials currently being considered for the Grumman SBR Phased-Array Antenna. Degradation mechanism caused by thermal cycling, ultraviolet and charged-particle irradiation, applied load, and high-voltage plasma interaction were evaluated. Of primary interest was a study of the degradation of the polyimide film Kapton (DuPont trademark), the material considered for use in the antenna plane. The second objective was to investigate the interaction between high-voltage electrodes and typical spacecraft contaminants in simulation of discharge triggering across differen- tially charged dielectric surface (spacecraft charging conditions). Experiment, Approach:: The experiment occupied a 6-in.-deep end corner tray (H07) located on the space end of LDEF and consisted of both passive and active parts. The passive part addressed the effect of environment and stress on dimensional stability of spliced and continuous Kapton, both plain and reinforced. Flight and ground-based test methodologies to measure the time-dependent deformation of large space structure materials under applied stress have been developed. Deflections on the order of 0.001 - 0.0001 in./in. were measured on 10-in.-long specimens. The specimen array contained eight 1-in.-wide specimens and sixteen 0.5-in.-wide specimens exposed directly to the space environment and a like number of shadowed specimens. The shadowed specimens underwent limited thermal excursion and were exposed to minimal solar and charged-particle radiation and were protected from direct atomic oxygen and micrometeoroid and debris impact. Each specimen contained a bonded splice located so that both spliced and continuous Kapton regions could be tested after exposure. Roughly one-half of the area was occupied by tension assemblies containing panels of Kapton and glass/Kapton maintained at one of four stress levels (30, 150, 300, and 450 psi) which were selected based on the anticipated SBR antenna plane average sustained and peak local stresses. The maximum stress was selected to accelerate the extent of creep. Stress levels were maintained by springs of known force constant. It was of interest to determine the mechanical integrity, dimensional stability, and extent of physical property degradation these panels suffered from long-duration exposure to the space environment. The remainder of the tray contained a portion of the SBR antenna plane and active electronic components. The active part of the experiment addressed the issue of the interaction between high voltage and low Earth orbit plasma. This part of the experiment was designed to investigate discharge triggering across high-voltage electrodes, and determine the effects of the subsequent high-voltage discharge events on the Kapton of the antenna plane. A 14-by 28-in. section of the Grumman SBR antenna (two Kapton antenna planes and a perforated aluminum ground plane) was selected as the test specimen. The electronic components included a high-voltage power supply, timing control circuitry, and a memory system. The electrodes provided by copper dipole ele- ments deposited on the Kapton plane were held at 1 and 2 kV. A counting circuit and recently available Grumman-designed microprocessor memory using EEPROM's (a nonvolatile, electrically erasable, programmable read only memory) were used to record the number of electrical discharges. The experiment timer-sequencer delayed the application of high voltage for 16 days and powers up the memory subsystem every 20 minutes for 0.6 second. Special testing circuitry was included to assure proper circuit operation, and mission time scale speedup capability was included to permit simulation of the entire flight. Associated Tray(s)::Tray Location: H07 - Orientation: Space-facing end Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L92-21195 KSC #: None JSC #: S32-S-288|LaRC #: L90-10368 KSC #: None JSC #: S32-75-062|LaRC #: L89-04390 KSC #: KSC-384C-255.09 JSC #: None|LaRC #: L91-15659 KSC #: KSC-390C-1990.01 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO134 Experiment Name::Space Exposure of Composite Materials for Large Space Structures Investigator(s)::Slemp, Mr. Wayne - Invest. Role: Original|Slemp, Mr. Wayne - Invest. Role: Original|Slemp, Mr. Wayne - Invest. Role: Present|Slemp, Mr. Wayne - Invest. Role: Present Experiment Description:: As space systems become larger and more complex, they require much longer lifetimes in space to be economically feasible. Currently these mission lifetimes are projected to be 10 to 20 years for antenna systems and up to 30 years for a solar-powered satellite system. This requires the structural materials to perform for much longer lifetimes than those required for current spacecraft. It can be assumed that electrical or electronic systems may be replaced or repaired, but the structure should generally be maintenance-free for the duration of these missions. Resin matrix composite materials offer unique advantages over conventional metallic materials for large space system applications due to their superior strength and stiffness-to-weight ratios and their low coefficient of thermal expansion. The major problem in utilizing composites for long-term space structure applications is the absence of data on the effects of space radiation on the mechanical and thermophysical properties of these materials. Although ground laboratory testing programs are in progress, these programs are substantially impaired by lack of information on the effects of space radiation on the properties of these materials. Without a space-flight-generated data base, it is difficult to project the useful life of these materials. The same is true of other classes of materials such as polymeric films. Objective:: The objective of this experiment was to evaluate the effects of the near-Earth orbital environment on the physical and chemical properties of laminated continuous-filament composites and composite resin films for use in large space structures and advanced spacecraft. Also, the experiment objective was to evaluate the response of resin matrix composite materials to extended LEO exposure, which includes atomic oxygen, ultraviolet and particulate radiation, meteoroid and debris, vacuum and temperature cycling. Experiment, Approach:: The experiment was passive and occupied about one-half of a 6-in.-deep peripheral tray. Some materials in this study were flown inside an Experiment Exposure Control Canister (EECC), which was closed when LDEF was launched. The EECC was programmed to open one month after deployment and to close 10 months later. Specimens of similar materials were placed outside the canister. Specimens of composite materials and polymeric and resin films were arranged above and below the experiment mounting plate to enable both exposure and nonexposure to sunlight. This provided a comparison of the effects of ultraviolet plus vacuum plus thermal cycling and those of vacuum plus thermal cycling on these materials. The experiment tray was thermally isolated from the LDEF structure to allow the material specimens to experience a wide range of thermal cycles. Tensile and compression specimens were used to evaluate the laminated composite materials. A number of the specimens were precut and ready for testing after space exposure, whereas other specimens were prepared from larger samples. Both 0.005-in. and 0.003-in.-ply thicknesses of prepreg (resin-impregnated material) were used. The tensile specimens were fabri- cated with (+45 /-45) degrees layup. The effects of flight exposure were evaluated by determining the stress-strain and ultimate tensile strength before and after flight exposure. Metal matrix composites were also included to evaluate the changes in coefficient of thermal expansion. Polymeric and resin films (e.g., Mylar, Kapton, P-1700 polysulfone, and FEP Teflon(R)) were used to provide additional data on the behavior of polymers in space. Data were obtained to determine the thermal stability, glass transition temperature, dynamic modulus, and loss modulus using a thermogravimetric analyzer, a thermomechanical analyzer, and a dynamic mechanical analyzer. An IR scan and an elemental analysis were also performed before and after flight exposure. A series of laboratory tests were conducted on all materials to simulate LDEF space flight conditions. The laboratory test program included one set of specimens exposed to UV radiation at one solar constant in 0.0000001 torr vacuum for 1000 hours and another set of specimens exposed to vacuum for the duration of LDEF flight exposure. These specimens provide data to help isolate the effects of ultraviolet light, vacuum, and time on the flight specimens. Associated Tray(s)::Tray Location: B09 - Orientation: 8.1 degrees off ram incidence angle; leading edge Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L91-02536 KSC #: KSC-390C-1913.09 JSC #: None|LaRC #: L84-07027 KSC #: KSC-384C-59.01 JSC #: None|LaRC #: L90-10374 KSC #: None JSC #: S32-76-001 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO135 Experiment Name::Effect of Space Exposure on Pyroelectric Infrared Detectors Investigator(s)::Robertson, Dr. James - Invest. Role: Present|Clark, Mr. Ivan - Invest. Role: Present|Robertson, Dr. James - Invest. Role: Original|Clark, Mr. Ivan - Invest. Role: Original|Crouch, Roger - Invest. Role: Original Experiment Description:: NASA's commitment to air pollution monitoring and thermal mapping of the Earth, which includes the remote sensing of aerosols and limb scanning infrared radiometer projects, requires photodetection in the 6 to 20 micron region of the spectrum. The Hg-Cd-Te detectors that are used in these wavelengths must be cooled to 50 to 80 K. The cryogenic systems required to achieve these temperatures are large, complex, and expensive. Pyroelectric detectors are one of many different types of infrared radiation detectors. The pyroelectric detectors are of interest for long-term space use because they do not require cooling during operation. Also, they can detect at very long wavelengths and they have a relatively flat spectral response. Pyroelectric detectors can detect radiation in the 1 - 100 micron region while operating at room temperature. This make the pyroelectric detector a prime candidate to fill NASA's thermal infrared detector requirements. A disadvantage is that the radiation must be chopped in order to be detected by a pyroelectric detector. Objective:: The objective of this experiment was to determine the effects of long-duration space exposure and launch environment on the performance of commercially-available pyroelectric detectors. This information is valuable to potential users of pyroelectrics for predicting performance degradation, setting exposure limits, or determining shielding requirements. Experiment, Approach:: The approach was to measure important detector parameters on a number of detectors before and after flight on the LDEF and try to determine the amount and cause of the degradation. Performance parameters to be measured were responsivity, detectivity, and spectral response. Material properties to be measured were pyroelectric coefficient and dielectric-loss tangent. The detector parameters to be measured before and after flight were signal strength, noise, and detectivity, calculated from signal and noise data. After the detectors were returned to the laboratory, all tests and measurements were repeated to determine the amount and type of damage suffered during launch and exposure. The experiment was passive; no data were taken during flight. The detectors for this experiment were included with the various electro-optical components of experiment S0050, Investigation of the Effects of Long Duration Exposure on Active Optical System Components, and were located in the same experiment tray. Commercially available detectors were purchased for the experiment. A total of twenty detectors were flown on LDEF and another seven were kept on the ground as controls. There were three different types of detector based on the type of pyroelectric material used: lithium-tantalate (LTO), strontium-barium-niobate (SBN), and triglycine-sulfide (TGS). These three detector types were treated as three separate experiments. The detectors were mounted on tray E05, which was a slightly-trailing-side location. The tray was covered with a perforated aluminum plate for thermal control. The plate blocked 50 percent of incident radiation. Four of the twenty flight detectors were further shielded by a solid aluminum plate which blocked their view of space but left them exposed to space vacuum. Associated Tray(s)::Tray Location: E05 - Orientation: 128.1 degrees off ram incidence angle Photograph Classification::Prelaunch Associated Photograph(s)::LaRC #: L89-04387 KSC #: KSC-384C-255.06 JSC #: None|LaRC #: L90-10413 KSC #: None JSC #: S32-77-050|LaRC #: L91-10508 KSC #: KSC-390C-2114.03 JSC #: None|LaRC #: L90-13450 KSC #: KSC-390C-1035.05 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-01 Experiment Name::Study of Meteoroid Impact Craters on Various Materials Investigator(s)::Mandeville, Dr. Jean-Claude - Invest. Role: Original|Mandeville, Dr. Jean-Claude - Invest. Role: Present|Mandeville, Dr. Jean-Claude - Invest. Role: Present|Mandeville, Dr. Jean-Claude - Invest. Role: Original Experiment Description:: Interplanetary space contains solid objects whose size distribution continuously covers the interval from submicron-sized particles to km-sized asteroids or comets. Some meteoroids originate from comets (mainly dust ejected at perihelion), some originate from collisions within the asteroid belt. A majority of particles are likely to come from comets but data from the Infrared Astronomy Satellite (IRAS) indicates that asteroids could be a source larger than expected. In addition to natural particles, a significant and growing number of particles has been added by human activity in near Earth space. In the vicinity of Earth, gravitational perturbations and the influence of the atmosphere greatly affect the distribution of the particles. Present knowledge of the occurrence and physical properties is based primarily on Earthbound observation of meteors, comets, zodiacal light, data from infrared satellites (IRAS) as well as on board measured flux by instrumented spacecraft (Pegasus, Vega, Giotto, Space Shuttle and the MIR Soviet Space Station ), study of lunar samples and dust collection in the upper atmosphere. The spatial density (number per unit volume) of meteoroids varies as a function of distance from the sun, distance from a planet, ecliptic latitude and longitude. The lifetime of interplanetary dust is dynamically limited, gravitational and solar radiation pressure (Poynting Roberston effect) gradually reducing the size of the orbit after typically 10,000 years; the lifetime of particles is also controlled by collision processes. Submicron particles will be blown out of the planetary system by solar radiation pressure (beta meteoroids). In the vicinity of Earth, gravitational perturbations and the influence of the atmosphere greatly affect the distribution of the particles. In-situ detection and collection of dust by experiments flown on LDEF are expected to improve our current understanding of this aspect of the space environment. Interplanetary dust particles (micrometeoroids) were expected to form well-defined craters upon impacting exposed material in space. Studying the frequency and features of these craters will provide data on the mass-flux distribution of micrometeoroids and, to a lesser extent, on the velocity, magnitude and direction. Limited crater studies have been done in the past with materials retrieved after exposure in space on Surveyor 3, Apollo 4 and 11, Gemini 10 and 11, and Skylab. However, little had been learned regarding the composition of impacting particles. This experiment focused on the determination of the composition of meteoroid material residues inside craters. The determination of the chemical composition of the impacting particles is a critical issue. In general they are physically destroyed and mixed with target material in the process of crater formation. Even though little or no pristine material may be left for chemical analysis, particularly in metals such as tungsten or gold, it is possible to collect quite sufficient projectile residue material for analysis. Based on laboratory experiments, such residues may be reduced to a probable initial composition. Objective:: This experiment studied impact craters produced by micrometeoroids on selected materials (metals and glasses in the form of thick targets) to obtain valuable technological and scientific data. Specifically, the studies focused on determining micrometeoroid composition and mass-flux distribution. Analyses were also made on the distribution of impact velocity vectors. Experiment, Approach:: A variety of sensor and collecting devices have made possible the study of impact processes on materials of technological interest. Preliminary examination of hypervelocity impact features give valuable information on size distribution and nature of interplanetary dust particles in low Earth orbit, within the 0.5-300 micron size range. High-velocity impact effects on various materials have been studied extensively in the laboratory. It is, however, impossible to obtain velocities higher than 7 to 8 km/sec with relatively large particles (less than 1 ng). The LDEF provided a unique opportunity to expose large-area targets for an extensive period of time and to recover then for subsequent analysis. Selected materials that act as impact detectors were exposed to space. This passive method consisted of thick targets (compared to the dimension of expected particles) of pure metals and glass. The collecting area (approx. 750 cm sq) is expected to record, with a high probability, impacts of micrometeoroids with a mass in the range of 10(-14) to 10(-7) g (corresponding to 0.2 to 35 micron particle diameters). The experiment was accommodated on an aluminum mounting plate (420 by 400 mm) located in one-sixth of a 12-in.-deep peripheral tray that contained nine other experiments from France. Two types of samples were used. Type I were metallic surfaces (100 by 100 mm) bolted to an aluminum mounting plate (6 samples). Type II were glass samples (25mm in diameter) bolted to an aluminum mounting plate (27 samples). A set of similar samples remained in the laboratory for subsequent comparison with space-exposed samples. A set of samples were also retained to perform tests with a hypervelocity accelerator. The first task after experiment retrieval was a careful scanning of exposed material to search for impact microcraters. Usually, high-velocity craters produced in metals and in brittle materials have a distinctive morphology and can be distinguished easily from other surface features. An optical microscope was used to identify craters larger than about 10 microns. Crater size distribution from these thick target experiments enables, with the aid of laboratory calibration by solid particle accelerators, the evaluation of the incident microparticle flux in the near Earth environment. Information on the velocity, particle density and incident direction are generally difficult to decode; however this could be partially determined by studying the geometry of impact craters. Measurements made on craters included diameter and depth measurements that, by comparison with experimental results, give the mass and density of impacting particles. Velocity can be estimated from the morphology of craters produced on brittle materials. Impact direction can be evaluated by the shape of the craters. There is a close relationship between the circularity of craters and the angle of impact. The relationship between flux and mass will be derived from the areal density of microcraters. For craters showing evidence of remnants of the projectile, chemical analysis was made with Xray microprobe and ion or Auger microprobe. If possible, atomic absorption spectrophotometry or neutron activation analysis was used. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-02 Experiment Name::Attempt at Dust Debris Collection with Stacked Detectors Investigator(s)::Mandeville, Dr. Jean-Claude - Invest. Role: Original|Mandeville, Dr. Jean-Claude - Invest. Role: Present|Mandeville, Dr. Jean-Claude - Invest. Role: Present|Mandeville, Dr. Jean-Claude - Invest. Role: Original Experiment Description:: Since the beginning of space exploration, a significant amount of data have been gathered on micrometeoroids. Since the first NASA U-2 flight collection in 1974, the collection and analysis of IDP orbiting around the Earth have been greatly enhanced. This enhancement occurred especially by the analyses of hard collectors that were exposed in low Earth orbits, before the impacting grains could have been processed by their entry in the Earth's atmosphere. Flux-mass relationships, velocities, and orbits of the particles have been established with meteoroid impact and penetration detectors on satellites and space probes. Also, studies of impact craters on lunar samples and a few retrieved sample materials exposed to space have added data on micrometeoroids. However, these techniques have limitations that prevent the study of undisturbed particles. The primary interest in the analysis of IDPs arises from the possibility that an unknown fraction of these particles could be of cometary origin and thus contain information on the early history of the solar system. In addition, asteroidal and interstellar particles may also be present. Cometary materials is likely to be the most primitive material accessible for analysis. It is thought that grains once present in the cometary nuclei and now present as individual grains in interplanetary space are the best candidates for having remnant properties that were acquired before and/or during condensation in the protosolar nebula. The smaller size fraction (grains less than 10 microns in diameter) are assumed to be enriched in grains of cometary origin. Collected IDPs have been subjected to various kinds of irradiations, inside the past and present solar system. Benit and Bibring have theorized that these different irradiations of grains could result in different physical, chemical and isotopical properties. In particular, carbonaceous material present in some grains could have been synthesized during early periods of intense solar irradiation. Manmade orbital debris is also present and many of these particles had velocities similar to some IDPs. Debris particles are recognizable by their compositional signature (Ti or Zn of paint flakes, aluminum oxide spheres or lack of a chondritic composition, etc). To study undisturbed particles, cosmic-dust collectors have been flown on balloons and rockets, and more recently on high-altitude aircraft. These techniques for dust collection in the atmosphere are limited because of short exposure times and uncertainty in the discrimination between cosmic-dust particles and terrestrial contaminants. Thus, among all the spacecrafts returns, LDEF was the first one designed to study the effects of space environment and to determine particle flux and orbital parameters. The FRECOPA experiments, in particular, were devoted to the study of dust particles and contained two entirely passive experiments flown for the detection of micropar- ticles: AO138-01 and AO138-02. Objective:: The primary objective of this experiment was to gain information on the micron-size fraction of IDPs by hard capture into a high purity aluminum surface. A secondary objective was to investigate the feasibility of multilayer thin-film detectors acting as energy sorters to collect micrometeoroids, if not in their original shape, at least as fragments suit- able for chemical analysis. It is expected that this kind of particle collector will help in solving one of the most puzzling topics in cosmic-dust studies, the mineralogical and chemical composition of the particles. Even though the impacted particles are mostly destroyed (some intact grains survive moderate to low velocity impacts), meaningful information on composition, flux, and particle size can still be obtained. Moreover, the light elements, particularly the biogenic elements, C, H and N, and possibly intact carbonaceous compounds, can be suitably analyzed to characterize possible cometary particles. This is a matter of great interest in the study of the origin and evolution of the solar system. Experiment, Approach:: The multiple foil penetration and collection experiment samples were mounted on a plate inside one of the three FRECOPA boxes in a 12-in.-deep tray that con- tained nine other experiments from France. The tray was located on the face of LDEF directly opposed to the velocity vector (west facing direction) in location B03 according to the LDEF description. Its position was assumed to be exposed mostly to grains of extraterrestrial origin. The experiment included 31 samples with a sampling surface area of 240 cm sq. The FRECOPA box provided maximum protection for the fragile thin metal films before and after space exposure. All samples were mounted within aluminum frames (40 by 40 mm or 30 by 30 mm) which held the thick target and the thin foils in front. The canisters were opened a short time after LDEF deployment and closed nine months later. The experiment consisted of targets made of one or two thin metal foils placed in front of a thicker plate. The maximum sample thickness of 125 microns was chosen to prevent foil perforation by particles with a mass in the range of 0.01 ng. The behavior of hypervelocity particle impacts on thin foils was extensively studied in the laboratory and these data provided a basis for interpretation. A particle penetrating the foil undergoes either a deceleration or fragmentation, depending on impact velocity, density of the target or projectile, and thickness to diameter ratio. Foil thicknesses chosen for this experiment range from 0.75 to 5 microns of aluminum andwere expected to decelerate particles with diameters between 1 and 10 microns without complete destruction. One or more thin metallic foils were set in front of the main target in order to produce size selective detectors. Separation distance between foils was 1mm, enough to have eventual fragments dispersed over a large area. Measurements after flight exposure were similar to those described for Experiment AO138-01. Emphasis was on the study of thin-film behavior during the cratering process and on the chemical analysis of projectile remnants. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-03 Experiment Name::Thin Metal Film and Multilayers Experiments Investigator(s)::Delaboudiniere, Dr. Jean Pierre - Invest. Role: Original|Carabetian, Dr. C. - Invest. Role: Present|Delaboudiniere, Dr. Jean Pierre - Invest. Role: Present|Hochedez, Dr. J.F. - Invest. Role: Present|Berset, J.M. - Invest. Role: Original Experiment Description:: It is well known that ultraviolet (UV) and extreme ultraviolet (EUV) experiments suffer degradation during space missions of even 1 month duration. It is believed that the degradation is due mainly to condensation of outgassing products, followed by solar-induced polymerization. However, penetrating charged particles are also known to produce volume effects. On the other hand, degradation may start immediately after manufacturing of the component due to oxidation, moisture, or chemical corrosion by atmospheric constituents such as CO2 and SO2. Finally, when the filters are used as windows for gas absorption cells or gas filters, or when they define the instrumental bandwidth by themselves (as in photometers and colorimeters), the effects of mechanical degradation by thermal cycling and/or dust impact may be dramatic. In preparation for the SOHO mission planned by NASA/ESA for launch in 1995, test samples of optical components to be used by the Extreme Ultraviolet Imaging Telescope were placed on LDEF. These components included thin film filters used for visible light rejection, and a new type of optical reflectors developed for the EUV since 1975. These reflectors consisted of a periodic stack of multilayered thin films, deposited on glass substrates. They operated by building up reflectivity from constructive interference of individual beams reflected at the interface between successive highly and weakly absorbing materials. The layer thicknesses are a fraction of the wavelength of the light beam reflected so that the period of the structure is lambda/2 (lambda being the operating wavelength). samples of such mirrors were located on the LDEF in order to evaluate the effect of thermal cycling, and surface contamination in low Earth orbit. Objective:: The objectives of this experiment were to investigate the sources of degradation of both state-of-the-art and newly developed components and to test the usefulness of the concept of storing experiment samples in dry nitrogen under launch and space vacuum conditions during reentry mission phases. Experiment, Approach:: A complement of EUV optical components, including mirrors and thin film filters, were flown as part of LDEF AO138-03. The experimental approach was to passively expose EUV thin films and UV filters to the space environment for postflight measurement and comparison with preflight measurements. The most original amongst these components were multilayered interference reflectors for the 10-40 nm wavelength range. The experiment was located in one of the three FRECOPA boxes in a 12-in.-deep peripheral tray that contained nine other experiments from France. The FRECOPA box provided protection for the experiment samples during the launch and reentry phases of the LDEF mission. All samples were manufactured at LPSP according to carefully controlled techniques and were separated into two exactly similar lots, one of which served as control samples and were stored under vacuum conditions in the laboratory. The samples were produced by electron beam vacuum deposition at Institut d'Optique (IOTA). Their reflectivity was measured at the synchrotron radiation source ACO of LURE (Universitie d'Orsay). Two lots of samples were placed in one of the FRECOPA containers for launch on board LDEF. This container was vacuum tight and filled with dry nitrogen at 10 mbar pressure several months before launch on the Space Shuttle. The pressure within the container was monitored until very close to the launch date, so that by extrap- olation it was certain that the samples had been exposed only to dry nitrogen before deployment. Spare components in a spare container were kept on ground in the same high vacuum conditions. The flight container was opened several days after the LDEF separated from the Shuttle. After seven months in orbit the container closed again under high vacuum conditions. The components which were exposed to the low Earth orbit space environment on board LDEF consisted of two identical lots. One was mounted in the FRECOPA box to see solar illumination and the other was protected from solar illumination. Direct impingement of atmospheric particles (atomic oxygen) was only be possible for the lot exposed to the sun. However the FRECOPA container was placed in the wake of the LDEF spacecraft so that bombardment by atomic oxygen was minimal, as is expected for the SOHO spacecraft, at the L1 Lagrangian point between the earth and the sun (1.5 million kilometers away from the Earth). After the flight samples were retrieved, their optical properties and the optical properties of the control lot were remeasured. Each half-lot of the flight samples contained 12 EUV free-standing single-layer thin films (set S(1)), 24 EUV single-layer thin films (set S(2)), 12 multilayers deposited on glass substrates (set S(3)), and 4 UV crystals (set S(4)). The filters in sets S(1) and S(2) were thin (1500 to 3000 Angstroms) films of selected metals. Under such thickness, good optical transmission is obtained in wavelength bands approximately 100 Angstroms wide. Selected materials that provide bands in the extreme ultraviolet included Al, Al + C, Sn, and In. The metallic multilayers in set S(3) were new optical components for the EUV region. Interference effects within a stack of alternatively absorbing and transparent materials of appropriate thickness were used to increase the reflecting efficiency within a narrow wavelength range. The number of periods was on the order of 10 to 40. Layers of Si/W and C/W were scheduled to be included if available. The UV crystal filters in set S(4) were relatively thick (2 mm) crystal windows of LiF and MgF2 and were of general use in the far UV range. For the preflight test program, fabrication controls and preliminary EUV bandpass measurements were made with a grazing incidence monochromator. The full bandwidth of free-standing filters and multilayers were measured with a synchrotron light source in several wavelength intervals from 40 to 2000 Angstroms, with emphasis on the interval from 100 to 1000 Angstroms. Optical constants were measured from deposits on glass substrates. Preflight measurements were repeated for the postflight test program, and surface physical and chemical analyses were made using the samples deposited on glass substrates and some of the free-standing films in cases where drifts have been observed. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-04 Experiment Name::Vacuum-Deposited Optical Coatings Experiment Investigator(s)::Malherbe, A. - Invest. Role: Original|Charlier, Dr. Jean - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present Experiment Description:: In the past, the Matra Optical Division has developed a wide range of optical components manufactured by vacuum deposition, such as metallic and multidielectric reflective coatings in the UV range, metal-dielectric interference filters in the UV (down to 1100 Angstroms) and IR ranges, narrow bandpass filters in the near-UV and visible ranges, selective metallic mirrors in the range from 1500 to 2500 Angstroms, antireflective and reflective IR coatings, beam splitters in the visible and IR ranges, and optical surface reflection (OSR) coatings. Many of these components, some of which were the first of this type to be manufactured in the world (e.g., interference filters at Lyman alpha), have been incorporated into scientific and technical experiments flown on balloons and rockets as well as on Symphonie, Meteosat, OTS, D2-B, TIROS N, and others. These components appear to have operated successfully in flight, but detailed information concerning their long-term behavior is not available. Objective:: The objective of this experiment was to analyze the stability of vacuum-deposited optical coatings exposed to the space environment. Experiment, Approach:: The experimental approach was to passively expose samples of the optical coatings of interest. The AO138-04 FRECOPA experiment consisted of 20 sorts of optical components and coatings subjected to space exposure. They cover a large range of the UV to IR spectrum: filters, mirrors, dichroics, beam splitters and anti-reflection coatings made of several different materials as layers and substrates. All layers were deposited by evaporation under vacuum. Five samples of each component or coating were manufactured. Two flight samples were mounted in the canister, one directly exposed to space, the other looking backwards. In a similar arrangement two spare samples were stored on ground and kept under vacuum during the LDEF flight. Finally, one set of reference samples was stored in a sealed canister under dry nitrogen atmosphere. Preflight and postflight optical measurements, including visual and microscopic inspections, were compared to determine the effects of space environment exposure. The experiment was located with nine other experiments from France in a 12-in.-deep peripheral tray. The optical coating samples was located in one of the three FRECOPA boxes located in the tray. The FRECOPA box provided protection from contamination for the samples during the launch and reentry phases of the LDEF mission. Preflight and postflight measurements included visual and microscopic examination and spectrophotometric analysis. For samples that show changes, microphysical analyses were performed by experienced laboratories. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-05 Experiment Name::Ruled and Holographic Gratings Experiment Investigator(s)::Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Bonnemason, Dr. Francis - Invest. Role: Present|Moreau, Gilbert - Invest. Role: Original Experiment Description:: In the past, several ruled and holographic gratings from Jobin-Yvon with various coatings were successfully flown on rocket experiments from LPSP and other organizations as well as satellites D2-A, D2-B, OSO-I, and some from the U.S. Future utilizations of such gratings have been considered for various projects by France, Germany, Belgium, and other countries. The technique used to replicate gratings can also be used to obtain a wide range of lightweight optical components, including sophisticated aspherical, highly polished mirrors. Objective:: The objective of this experiment was to test the behavior of ruled and holographic gratings with various coatings after extended exposure to the space environments. Specific objectives included examining the coatings for possible changes and differentiating between the influences of vacuum and solar illu- mination. Experiment, Approach:: The experiment approach to study the optical behavior of passively exposed diffraction gratings with coatings depending on space vacuum and long environmental exposure. To separate these environmental parameters influence from senescence on ground, three sets of samples were manufactured. Preflight and postflight examination to characterize the optical quality of the grat- ings included measurement of wavefront flatness, reflection efficiency, and stray-light level. These optical performances tests for the before and after flight measurements were operated with the same instrumentation. By test results comparison and analysis we can appreciate the behavior of each type of gratings and coatings. The following is a list of the samples used in this experiment with a description of each. Type G is a replica of a grooved grating with 1200 grooves/mm blazed at 2500 Angstroms (aluminum coated blank pyrex). Type H is an original holographic grating. It has 3600 grooves/mm at the 50-150 nm spectral range and is platinum-coated. Type HU is an ion-etched blazed grating with 1200 grooves/mm blazed at 2500 Angstroms (aluminum-coated blank pyrex). Type W is a control mirror on blank pyrex. One half is coated with aluminum and the other half with platinum. Samples were located on both sides of the mounting plate within the FRECOPA box and have the shape of a parallelepiped with dimensions 34 by 34 by 10 mm. Witness mirrors were joined to the experiment to distinguish coatings behavior. A set of control samples were also stored in the laboratory, under air-nitrogen for the flight duration, for comparison with those retrieved from space. Preflight and postflight measurements to be made included the following parameters. 1. Wave surface flatness.-This separates changes introduced by groove distortion and blank distortion and was measured on an order of zero for type H, G, and W, and on an order of one for types G, H, and HU. The measurement were made by photography using a Michelson interferometer. 2. Reflection efficiency.-This was measured with a photogoniometer from 2200 to 6000 Angstroms for types G, HU, and W and with a vacuum photogoniometer from 584 to 1216 Angstroms for types H and W. 3. Stray light level.- For types G and HU, measurements were made using a continuum spectrum (deuterium lamp) near 2000 Angstroms on a monochromator with a liquid filter. For type HU, measurements were made using a laser line (6328 Angstroms) on a monochromator. These measurements help define the limits of utilization for each type of grating. Comparison of the different measurements before and after space exposure help define the space environment elements that cause degradation of grating optical qualities and the grating components which are damaged by those elements. The selected materials and manufacturing processes were representative of standard grating production for loaded spectroscopic applications: one set stored on ground, one loaded space environment and sun illumination shaded, one loaded space environment and sun illuminated during a pre-defined duration (10 months). The experiment was located with nine other experiments from France in a 12-in.-deep peripheral tray (B03). The grating samples were located in one of the three FRECOPA boxes located in the tray. The FRECOPA box will provide protection for the samples from contamination during the launch and reentry phases of the LDEF mission. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-06 Experiment Name::Thermal Control Coatings Experiment Investigator(s)::Guillaumon, Dr. Jean-Claude - Invest. Role: Original|Scialdone, Dr. John J. - Invest. Role: Present|Paillous, Dr. Alain - Invest. Role: Present|Paillous, Dr. Alain - Invest. Role: Original|Guillaumon, Dr. Jean-Claude - Invest. Role: Present Experiment Description:: In order to assess the degradation of themo-optical properties of coating used on satellites, a space environment simulation is needed. To develop accurate simulation techniques, space data are needed. Spacecraft materials are exposed to the various components of the natural space environment: vacuum, atomic oxygen, UV sun radiation, ionizing particle fluxes, micrometeoroids. Other environment elements which are due to spacecraft operations have also to be considered: thermal cycling on board, contamination by molecular products and dust, debris, etc. All these components, acting alone or synergisticially, are deleterious in their effects on materials. Therefore the degradation which was experienced in space has to be predicted by space environment simulation tests to be carried out on ground. These tests must be simple (addressing only the most dangerous environment elements) and they must give reliable data in accelerated conditions. It is difficult to perform such a task in the laboratory because the simulation involves good vacuum, temperature programming, and irradiation by ultraviolet light and particles. In most cases, caution must be used in interpreting the results because it is impossible to obtain light sources with a spectrum similar to that of the Sun and also because accelerated tests are generally used. A comparison of degradation obtained in the laboratory with degradations obtained in space would be very valuable. Objective:: The purpose of AO138-06 was to evaluate the degradation experienced in space by a number of materials on board spacecraft placed into low Earth orbit (LEO) by the Space Shuttle and to validate simulation techniques used at DERTS, especially those employing UV-irradiations with in situ and ex situ measurements. The main objectives were to expose material surfaces to either the whole environment for a satellite placed and retrieved in LEO by a Space Shuttle in all mission phases (handling before launch, ascent, maneuvers in orbit, external Shuttle environment, docking, retrieving, descent, transportation to KSC, etc.) or only the satellite environment during its autonomous free flight. Other objectives were to retrieve the samples for laboratory analysis and to verify the validity of space environment simulation performed in the laboratory in order to measure the stability of the thermo-optical properties of thermal control coatings; to compare the behavior in space of some materials for which the available ultraviolet solar simulation is inadequate (especially in the far ultraviolet); and lastly, to evaluate possible contamination effects due to the use of the Shuttle. Experiment, Approach:: The experimental approach was to passively expose samples of the thermal coatings of interest. These coatings include black paint, aluminum paint, white paint, conductive paint, a solar absorber, an optical surface reflector, second-surface mirrors, metal coatings, polymeric films, composite materials, and silica fabrics. Preflight and postflight measurements of thermo-optical properties were compared to determine the effects of space environment exposure. The experiment was located with nine other experiments from France in a 12-in.-deep peripheral trailing edge tray (B03). The thermal coating samples were housed in one of the three FRECOPA vacuum-tight boxes located in the tray. The FRECOPA box protected the samples from contamination during the launch and reentry phases of the LDEF mission. Samples were independently maintained in sample holders that allowed their front face to receive maximum solar illumination when the FRECOPA box was open. Thirty samples were tested. Twenty-nine samples were 3/4 in. by 3/4 in. and one sample (Optical Solar Reflector) is 1 1/2 in. by 1 1/2 in. Sample thickness is less than 1/8 in. The experiment was designed in order to allow exposure of a part of the samples to the whole spacraft environment in all mission phases. Twenty-four samples were directly exposed to space for the total flight duration (preflight handling, shuttle bay environment, separation from shuttle, shuttle environment, LEO environment, docking, descent, transfer to KSC) and these samples were laid directly on the FRECOPA tray surface on two separate sample holders. The other part of samples was protected from the external environment of LDEF for all mission phases, except free flight, by the means of a vacuum-tight FRECOPA canister in which they were stored. The maximum temperature during space exposure was recorded by passive temperature indicators fixed to the sample mounting plate. Additionally, the ionizing radiation dose was measured by a passive LiF dosimeter. The FRECOPA box was clock operated; it opened 10 days after separation from Shuttle and closed 10 months later. An O-ring butyl seal kept the samples under vacuum until optical measurements were completed in the laboratory. The entire closed box containing samples under vacuum was placed in a vacuum chamber, where the optical reflectance spectrum of each sample was recorded using an integrating sphere. For that purpose the FRECOPA canister was designed to be opened and closed in a vacuum chamber, thus avoiding, at maxium, air recoveries. An additional set of samples was maintained in the laboratory for comparison with samples subjected to space exposure. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-07 Experiment Name::Optical Fibers and Components Experiment Investigator(s)::Bourrieau, Dr. Jacques - Invest. Role: Present|Bourrieau, Dr. Jacques - Invest. Role: Original Experiment Description:: Fiber optics are becoming important components in communication systems, optoelectronic circuits, and data links. Space applications of optical fibers have these advantages: weight and size reduction, data transmission rate increase (10 to 100 Mbits), and reduction of electromagnetic susceptibility and power requirements. High sensitivity to ionizing radiation, however, may be a restriction for optical fiber use on satellites. An increasing number of laboratories are carrying out irradiation tests on these components using neutrons, gamma rays, and X-rays. Radiation damage on optical materials, however, is strongly linked to the test conditions (temperature, dose rate, energy, and nature of the incident particles), and laboratory tests are not as representative as actual space environment exposure. Objective:: The main objective of this experiment was the comparison of fiber optics permanent damages induced by ionizing radiation after a long exposure in space and after laboratory tests. Specific objectives were to validate irradiation tests performed with radioactive sources (Sr-90 - Y-90), to verify computer codes used for the fluence and dose profiles, to determine the performance of fiber optics waveguides in a low-altitude orbit (doses between 10 and 100 Gy were expected), and to determine the origin of transmission losses (e.g., color centers and index variations) in the material. Another objective of the AO138-07 experiment was a total dose measurement with thermoluminescent detectors (TLD 100) Experiment, Approach:: The experimental approach was to passively expose two optic-fiber waveguides (one step index and one graded index) of some 60 cm in length with connectors. Preflight and postflight measurements of optical properties were compared to determine the effects of space environment exposure. The experiment was located with nine other experiments from France in a 12-in.-deep peripheral tray. The optical fibers were located in one of the three FRECOPA boxes in the tray. The FRECOPA box provided protection for the optical-fiber waveguides from contamination during launch and reentry phases of the LDEF mission. The in-flight absorbed dose profile was measured with five thermoluminescent dosimeters shielded by various aluminum thicknesses. Irradiation of the selected optic fiber waveguides was carried out in the laboratory with an Sr-90 - Y-90 beta ray source in order to simulate the in-flight dose. Before and after flight and laboratory simulation, measurements of the optic-fiber waveguide light transmission were made with a spectrophotometer. Four sets of samples were manufactured, one for ground test and the others for control, flight, and space flight sets. Two identical packages, both of them including five lithium fluoride TLDs inside various aluminum shields, were exposed to the space environment in order to obtain the absorbed dose profile. Without data transmission, only passive dosimeters were used. The radiations encountered by the LDEF was expected to be mainly be trapped particles. Due to the orbit inclination the magnetospheric shielding is very effective for the solar protons and the GCR (galactic cosmic rays). Regarding the heavy ions from solar events their state of charge is not well known and consequently the magnetospheric absorption cannot be defined. Thus, significant exposure comes from trapped protons and electrons and the absorbed dose induced by GCR and solar cosmic rays (SCR) are weak. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-08 Experiment Name::Effect of Space Exposure of Some Epoxy Matrix Composites on Their Thermal Expansion and Mechanical Properties Investigator(s)::Soulat, Dr. Guy - Invest. Role: Present|Elberg, Robert - Invest. Role: Original Experiment Description:: Carbon and Kevlar fiber-reinforced plastic composites are being used increasingly in space structures (launch vehicles, spacecraft, payload elements such as antennas and optical benches). This extensive interest in composites is due to their mechanical properties (i.e., high strength and stiffness associated with a low density) and also to the near-zero value, positive or negative, of their coefficient of thermal expansion. This latter characteristic is due to the design of the composite (i.e., choice of fibers, fiber arrangement, and resin content). A particular point that Matra wished to examine was the effect of space environment on the thermal expansion stability of such products. The thermal expansion stability is a sensitive parameter that assures the desired performance of optical instruments such as telescopes and optical benches. This performance is related to short-term stability, which is assured simultaneously by the thermal control system and the low and stable coefficient of thermal expansion of the structure. Long-term dimensional changes that may occur (for instance, moisture desorption) are often compensated by a refocusing mechanism. Objective:: This experiment had three objectives. The first and main objective was to detect a possible variation in the coefficient of thermal expansion of com- posite samples during a 1-year exposure to the near-Earth orbital environment. A second objective was to detect a possible change in the mechanical integrity of composite products, both simple elements and honeycomb sandwich assemblies. A third objective was to compare the behavior of two epoxy resins commonly used in space structural production. Experiment, Approach:: The experimental approach was to passively expose samples of epoxy matrix composite materials to the space environment and to compare preflight and postflight measurements of mechanical properties. The experiment was located in one of the three FRECOPA (French cooperative payload) boxes in a 12-in.-deep peripheral tray that contained nine other experiments from France. The FRECOPA box protected the samples from contamination during the launch and reentry phases of the LDEF mission. Two identical samples of each type were to be flown with and four different configurations of samples. Two common characteristics of configurations A, B, and C were the length (4 in.), and the existence at each end of the samples of three protrusions. These were the reference points for the measurement of the coefficient of thermal expansion (CTE) The coefficients of thermal expansion were measured on Earth before and after space exposure. This measurement is based on a laser interferometry method working in a vacuum. The method used consisted of forming a fringe pattern generated by two almost-parallel reflecting surfaces fastened to the test sample at the level of the protrusions and illuminated by a stabilized mono- mode He-Ne laser beam. Length variation due to temperature variation results in a fringe motion, which is measured. CTE measurements were made at ambient temperature with a temperature variation of about 20 degrees C. The accuracy for a sample length of 4 in. is about +/- 0.0000001 per degree C. The test program consisted of the following measurements: Weight measurements of all the samples were made before launch (samples previously dried and outgassed) and after launch. Any weight variation is useful information to establish the efficiency of the drying and outgassing process for composite materials. The coefficient of thermal expansion was measured on samples 1-1', 2-2', 3-3', 4-4', and 5-5' before and after launch. Any effect on this parameter after a 1-year exposure at ultrahigh vacuum was detected. Micrographic inspections were conducted on cuts made on the various samples at the level of the composite materials and more specifically at the level of the honeycomb-face-sheet assembly on samples 2-2' and 6-6'. These photographs were compared to similar views of identical materials kept on Earth for reference. Mechanical tests for interlaminar shear strength, flexural strength, and flatwise tensile strength were made on elements cut from the samples after return to Earth. These test results were compared to results from similar samples that were not submitted to the space environment. Although vacuum chamber test of epoxy matrix composites show good mechanical and thermal stability, the results of this experiment are anticipated to increase confidence in the performance of the tested composite specimens, particularly relative to the thermal stability. Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-09 Experiment Name::The Effect of the Space Environment on Composite Materials Investigator(s)::Parcelier, Mr. Michel - Invest. Role: Original|Parcelier, Mr. Michel - Invest. Role: Present Experiment Description:: It is the duty of a satellite manufacturer to verify the characteristics of the products it designs and manufactures. The space industry has complied with this rule through laboratory simulations and evaluations. The launching of the LDEF by the Space Shuttle provided an opportunity to observe the actual behav- ior of materials under exposure to the space environment and has made it possible to correlate with artificial aging tests. Objective:: The objective was to test graphite fiber-reinforced epoxy matrix composite materials, laminates, thermal coatings, and adhesives to determine their actual useful lifetime. These experiments also make it possible to integrate the histories of the thermal and mechanical characteristics into models of the composite structures. Experiment, Approach:: The experiment was passive and was located in one of the FRECOPA boxes in a 12-in.-deep peripheral tray with nine other experiments from France. The FRECOPA box provided protection for the samples from contamination during the launch and reentry phases of the LDEF mission. The experiment included four themes of study: thermal coatings, adhesives, dimensional stability, and mechanical characteristics. The various materials were arranged in six levels with the FRECOPA box, so only the first level was subjected to direct solar radiation. Each level consisted of plates from which test specimens were cut after the mission. The arrangement of specimens chosen for this experiment had an influence on the data; due to the small surface area exposed to the direct space environment, and in order to have a greater number of specimens, six levels were arranged. The thermal coatings were put in the upper level in order to be directly exposed to space environment; in the lower levels, composite materials and adhesives were subjected to vacuum and thermal cycling. A large selection of materials with a limited number of specimens were used. In addition, the requirement on weight for this experiment (less than or equal to 1 Kg) led to minimizing the weight of the fixture: pre-cut specimens were replaced by small panels that require a lighter fixture to meet LDEF vibration specification. Thermal Coatings: Optical solar reflectors and second-surface mirrors were laid on aluminum and carbon supports. The tests evaluated the level of degradation, the mass of contaminants received, and the alteration of the thermo-optical properties. Adhesives: The tests measured the shear in a joint bonded with 408 (room-temperature curing) and 312L (high-temperature curing) adhesives. The purpose was to observe the effects of the thermal stresses due to the assembly of materials having different expansion coefficients (carbon and aluminum) and the thermal cycling in the low Earth orbit of the LDEF. Dimensional Stability: Tests were carried out to verify the predicted thermoelastic deformations in sandwich structures that have withstood the space environment. Expansion tests were made on a sandwich test specimen painted white (located on the upper level) and shaped like a satellite antenna, and also on the constituent parts taken singly. These parts were GY70/87 (0 degree and 90 degree orientation) and 312L. The same experiment were carried out on a sandwich test specimen cocured with 914 and 319L. Mechanical Characteristics The degradation of the mechanical properties (tensile, flexural, and interlaminar tests) were evaluated on the following materials: GY 70/87 (tape), GY70/914 (tape), T300/V108 (tape), and G837/V 108 (fabric). Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO138-10 Experiment Name::Microwelding of Various Metallic Materials Under Ultravacuum Investigator(s)::Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present|Assie, Mr. Jean-Pierre - Invest. Role: Original|Conde, Mr. Eric - Invest. Role: Present|Assie, Mr. Jean-Pierre - Invest. Role: Present|Alet, Dr. Irenee - Invest. Role: Present Experiment Description:: In the space vacuum environment, the spacecraft mechanisms are liable to sustain damaging effects from microwelds due to molecular diffusion of the spacecraft constituent metals. The microwelding of two contacting metal parts of the same mechanism are liable to sustain irreversible effects, detrimental to satellite missions. Such microwelds result in a continuing increase in the friction factors and are even liable to jam the mechanisms altogether. Spacecraft mechanisms include: pointing mechanisms (antennas, instruments), focusing mechanisms (optical components), deployment mechanisms (antennas, solar arrays), hinges and drive mechanisms (mechanical parts of solar arrays), and slip rings (electrical parts of solar arrays). Objective:: The objective of this experiment was to study the metal surfaces representative of the mechanism constituent metals (treated or untreated, lubricated or unlubricated) for microwelds after an extended stay in the space environment. From this, material choice can be optimized to avoid microweld induced damage to spacecraft mechanisms and electrical parts, and the weight of spaceborne harness and wiring may be decreased. Experiment, Approach:: The experimental approach was to passively expose inert metal specimens to the space vacuum and to conduct end-of-mission verification of the significance of microwelds between various pairs of metal washers. The test approach was to statically compare ground samples in terms of aspect, adhesion, microwelds. The following parameters were tested on the experiment materials: metal, surface, treatment, lubrication, and pressure. Also part of the experimental approach was the simulation on various metallic washer couples of different contacts used in static or single-hot space mechanisms and slip rings. The experiment will lead to an evaluation of the compatibility of aluminum conductors with conventional conductors and contacts under different production processes (mechanical or magnetostrictive crimping) and storage (laboratory or long-duration space vacuum). The test approach includes comparative, pre- and post-flight measurement of on-ground and flight samples. The experiment were located in one of the FRECOPA boxes in a 12-in.-deep peripheral tray, B03, that contains nine other experiments from France Associated Tray(s)::Tray Location: N. A. - Orientation: N. A. Photograph Classification::None Associated Photograph(s)::LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO139A Experiment Name::Growth of Crystals From Solutions in Low Gravity Investigator(s)::Nielsen, Dr. Kjeld Flemming - Invest. Role: Original|Nielsen, Dr. Kjeld Flemming - Invest. Role: Present|Lind, David - Invest. Role: Original|Lind, David - Invest. Role: Present Experiment Description:: These crystal growth experiments are an extension of preliminary experiments performed during the Apollo-Soyuz Test Project flight and similar experiments for a Spacelab flight. Following the crystal growth experiments with unidirectional conductors on the European Spacelab by K.F. Nielsen and the Apollo-Soyuz investigations on room temperature solute diffusion crystal growth by M.D. Lind, it was decided by both investigators to share what was later named experiment AO139A on the LDEF. The crystal growth method used consisted of allowing two or more reactant solutions to diffuse slowly toward each other into a region of pure solvent, in which they react chemically to form single crystals of a desired substance. This method depends on suppression of convection and sedimentation. Suppression of convection should also eliminate the microscopic compositional fluctuations often caused by time-dependent convection in crystal growth systems. In many cases, convection and sedimentation can be completely eliminated only under the conditions of continuous low gravity attained during orbital space flight. Ideally, this type of crystal growth process requires that a low level of gravity (less than 0.0004 g) be maintained for a period of weeks or months. The LDEF flights were the only space flights planned that fully satisfy this requirement. Objective:: The objective of these experiments was to develop a novel solute diffusion method for growing single crystals. Crystals to be investigated were PbS, CaCO3, and TTF-TCNQ. Each of these materials had current research and technological importance. PbS is a semiconductor, and CaCO3 has useful optical properties; both would have many applications if it could be synthesized as large, highly perfect single crystals. The important property of TTF-TCNQ is its one- dimensional electrical conductivity. The conductivity is strongly dependent on crystal perfection; crystal growth in low gravity is expected to yield larger and more perfect crystals, which may have unique electrical properties. The experiments were expected to yield crystals of each of the materials which are superior in size, structural perfection, and compositional homogeneity to those heretofore obtainable. The availability of such crystals should make possible better determinations of their physical properties and, perhaps, investigations of possible device applications. The experiments should also contribute to a better understanding of the theory and mechanisms of crystal growth. Experiment, Approach:: Each experimenter had two growth containers, and all four were kept at constant temperature (35 degrees C) in a thermostated drum. The growth containers were manufactured at the experimenters' individual laboratories, while the enclosure and the insulation equipment were manufactured at Rockwell. The four LDEF growth reactors were approximately the same size. A smaller drum containing the data and electronic equipment was manufactured at Terma Electronics, Ltd., in Denmark. The experiments utilized specially designed reactors with three or more compartments separated by valves to keep the reactant solutions and solvent separated until the apparatus reaches low gravity. A mechanism opened the valves automatically to initiate the diffusion and growth processes. The reactant reservoirs were large enough to take advantage of the time provided by the LDEF flight. An array of several reactors were mounted in a 12-in.-deep end center tray located on the Earth-facing end of the LDEF. Several reactors operating simultaneously to allow experimentation with more than one crystal growth system and/or variations of conditions for each. The reactors were enclosed in a vacuum-tight container and were surrounded by thermal insulation. The temperature was regulated and any departures from the desired temperature were recorded. Power requirements were provided by LiSO2 batteries. Associated Tray(s)::Tray Location: G06 - Orientation: Earth-facing end Photograph Classification::Flight Associated Photograph(s)::LaRC #: L90-10391 KSC #: None JSC #: S32-76-066|LaRC #: L89-04391 KSC #: KSC-384C-256.02 JSC #: None|LaRC #: L91-11642 KSC #: KSC-390C-1704.05 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO147 Experiment Name::Passive Exposure of Earth Radiation Budget Experiment Components Investigator(s)::Griffin, Francis - Invest. Role: Original|Mooney, Thomas A. - Invest. Role: Present|Smajkiewicz, Ali - Invest. Role: Present|Hickey, Mr. John R. - Invest. Role: Present|Hickey, Mr. John R. - Invest. Role: Present Experiment Description:: The PEERBEC experiment of the LDEF mission was composed of sensors and components associated with the measurement of the Earth Radiation Budget (ERB) from satellites. These components include the flight spare sensors from the ERB experiment which operated on Nimbus 6 and Nimbus 7 satellites. Earth Radiation Budget (ERB) experiments require accuracies on the order of fractional percentages in the measurement of solar and Earth flux radiation. In order to assure that these high-accuracy devices are indeed measuring real variations and are not responding to changes induced by the space environment, it is desirable to test such devices radiometrically after exposure to the best approximation of the orbital environment. The reason for proposing this experiment was to help explain the results which were obtained by the ERB of Nimbus 6. The Nimbus ERB experiments contained 10 solar radiation measurement sensors (channels) and 4 earth-flux measurement sensors. The solar sensors were mounted on the leading edge of Nimbus and the earth-flux sensors were looking down. Channel 10S of Nimbus 6 (a UV sensing channel) was replaced by a cavity radiometer for total solar radiation measure- ments for the Nimbus 7 Mission. Nimbus 6 was launched June, 1975, during the decline of solar cycle 20. After initial measurements, the solar channels having filters showed a slow degradation. Nimbus 7 was launched in October 1978, as the peak of solar cycle 21 was approaching. After allowing for a period of protection from outgassing contamination, the ERB solar channels were opened and started measurements on November 16, 1978. After initial measurements, the solar sensing channels having filters or windows, exhibited a major degradation. With the increase in solar activity of solar cycle 21 there was a major recovery of the signal for these channels. When the recovery was experienced for Nimbus 7, it was found that the Nimbus 6 sensors (then more than 4 years in orbit) had also recovered. After the initial events the response of the various channels differed. Channel 6S and 9S measured levels higher than their initial readings, while the other 2 interference filter channels (7S and 8S) showed steep declines. The channels with fused silica windows acted differently. While channel 2S showed a decline with only partial recovery, the normally shuttered channel 1S showed very little change until later in the mission when its shutter was left open. As solar cycle 22 approached maximum most of the channels exhibited some measure of recovery. Much information on the contamination of the ERB filtered channels could be gained from the Nimbus data. Nimbus 6 apparently suffered degradation due to its own outgassing and then was cleaned once by AO. The Nimbus 7 record, shows a much greater initial degradation due to contamination and a remarkable recovery due to AO cleaning. But the degradation of some channels, with only partial recovery due to AO cleaning at the next solar maximum, requires further explanation. Objective:: Since the Earth Radiation Budget experiment was operational on Nimbus 6 and Nimbus 7, and since in-flight calibration is difficult for the solar and Earth flux channels, the objective of this experiment was to expose ERB channel components to the space environment and then retrieve them and resubmit them to radiometric calibration after exposure. Subsequently, corrections may be applied to ERB results and information obtained to aid in the selection of components for future operational solar and Earth radiation budget experiments. Experiment, Approach:: Passive exposure of solar and Earth flux channel components of the ERB radiometer is the basis of the approach. The solar sensors were mounted in a Nimbus ERB solar array block.The 10 solar sensors were mounted in LDEF tray B08 along with 10 (non-ERB) interference filters supplied by Barr Associates. Since the B08 tray is 30 degrees off the ram direction, the solar block is tilted to face forward. The Barr Associates filters were mounted flat to the front plate, The 4 earth-flux sensors were mounted in LDEF tray G12 on the earth facing end. A cavity radiometer, similar to channel 10C of Nimbus 7 but not the same field of view, is included as part of the Advanced Photovoltaic Experiment (APEX) which is mounted in LDEF tray E09. One reason for employing cavity radiometers was that the effect of degradation should be much less than on a flat plate sensor. The cavity sensor was on the leading edge. The cavity radiometer and the filter based spectral radiometer were intended as calibration reference instruments for the solar cell measurements. While PEERBEC is a passive experiment, APEX is active. The Eppley Laboratory is involved with the design and fabrication of both the PEERBEC and APEX experiments. These flight spare ERB sensors underwent the same pre-flight processing as the ERB flight sensors according to the procedures of the Nimbus project. 1) All thermopiles and other painted surfaces underwent vacuum bakeout, 2) All sensor modules underwent thermal vacuum testing, 3) All interference filters were burned in (in vacuum) using a xenon arc solar simulator at 1 solar constant level; channels 6, 7, and 8 for 500 equivalent UV solar hours and channels 9 and 10 for 800 equivalent UV solar hours, 4) The Suprasil W fused silica windows and hemispheres were not burned in, 5) The 4 mm thickness of front windows of broad band channels and front substrates of the interference filters was chosen to act as a particle blocker, and 6) The rear windows, 2 mm thick fused silica, were IR blockers mounted behind broad band glass filters and interference filters. Three Earth flux channel types of ERB were mounted in one-fourth of a 3-in.-deep end center tray on the Earth-viewing end of the LDEF. Prior to delivery, these channels underwent complete radiometric and spectrophotometric examination. These tests were repeated after retrieval to evaluate changes in orbit. The solar-channel components were mounted in one-sixth of a 3-in.-deep peripheral tray near the leading edge of the LDEF (in the direction of the velocity vector) to view the Sun in the manner most like the ERB experiment on Nimbus. Solar-channel components to be tested included thermopiles, interference filters, and fused silica optical windows. Additionally, some state-of-the-art vacuum-deposited interference filters were included to examine space environment effects on these components. The two thermopiles had different black paint on the receivers. The cavity unit was similar to that proposed for future solar-constant measurement missions. Vacuum bakeout was performed on all elements as prescribed and performed for Nimbus prior to delivery. Radiometric testing and weighing were performed on the thermopile and cavity devices. Spectrophotometric testing were performed on filters and fused silica components. Associated Tray(s)::Tray Location: G12 - Orientation: Earth-facing end|Tray Location: B08 - Orientation: 38.1 degrees off ram incidence angle Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L90-13421 KSC #: KSC-390C-1031.11 JSC #: None|LaRC #: L84-07116 KSC #: KSC-384C-294.07 JSC #: None|LaRC #: L90-10388 KSC #: None JSC #: S32-76-054|LaRC #: L90-10380 KSC #: None JSC #: S32-76-026|LaRC #: L91-11718 KSC #: KSC-390C-1997.02 JSC #: None|LaRC #: L84-07165 KSC #: KSC-384C-317.08 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO171 Experiment Name::Solar-Array-Materials Passive LDEF Experiment Investigator(s)::Stella, Dr. Paul - Invest. Role: Original|Whitaker, Dr. Ann F. - Invest. Role: Present|Stella, Dr. Paul - Invest. Role: Present|Young, Mr. Leighton - Invest. Role: Original|Young, Mr. Leighton - Invest. Role: Present|Whitaker, Dr. Ann F. - Invest. Role: Original|Whitaker, Dr. Ann F. - Invest. Role: Present|Whitaker, Dr. Ann F. - Invest. Role: Original|Brandhorst, Henry - Invest. Role: Original|Forestieri, A.F. - Invest. Role: Original|Bass, James - Invest. Role: Original|Brandhorst, Henry - Invest. Role: Original|Gaddy, Edward - Invest. Role: Original|Forestieri, A.F. - Invest. Role: Original|Smith, Charles - Invest. Role: Original Experiment Description:: The long-duration functional lifetime requirements on lightweight high-performance solar arrays demand careful selection of array materials. The space environment, however, is a hostile environment to many materials, and some of the problems are well documented. A thermal-vacuum environment can affect materials by accelerating the outgassing of volatile species. The condensation of these outgassed products on array cover slips leads to reduced solar-cell electrical output, a situation that is especially critical at high astronomical units (AUs). Outgassing can reduce mechanical strength in materials, which will affect the integrity of the array substrate, hinges, and deployment mechanisms and create electrical problems through insulation breakdown. A further effect of outgassing is the degradation of thermal control and reflector surfaces. Some extended performance arrays that have been studied but never flown utilize deployable concentrators whose reflectance is especially important at large AUs. Protons, electrons, atomic oxygen, and UV irradiation contribute to surface damage in these array materials. Thin-film materials can be embrittled and thermal control surfaces can become discolored by this irradiation. Severe mission environments, coupled with the lack of knowledge of space environment materials degradation rates, require the generation of irradiation and outgassing engineering data for use in the design phase of flight solar arrays. Objective:: The objective of this experiment was to evaluate the synergistic effects of the space environment on various solar-array materials, including solar cells, cover slips with antireflectance (AR) coatings, adhesives, encapsulants, reflector materials, substrate strength materials, mast and harness materi- als, structural composites, and thermal control treatments. Electrical, mechanical and optical property changes induced by the combined space environments were studied. Experiment, Approach:: Experiment AO171, the Solar Array Materials Passive LDEF Experiment, includes contributions from NASA MSFC, NASA LeRC, NASA GSFC ,and JPL. The overall design and integration was done by NASA MSFC. It is located at A08, a near ram position. The passive experiment consisted of an arrangement of material specimens mounted in a 3-in.-deep peripheral tray. SAMPLE contained a complement of specimens that addressed primarily solar array materials, although other materials of interest to the aerospace industry were included. With the exception of experiment solar cell and solar cell modules, all test specimens were weighed before flight, thus allowing an accurate determination of mass loss as a result of space exposure. Since almost all of the test specimens were thermal vacuum-baked before flight, the mass loss sustained can be attributed principally to atomic oxygen attack. The effects of the space environment on the specimens were determined by comparison of preflight and postflight measurements of mechanical, electrical, and optical properties. The JPL portion of the experiment consisted of an aluminum test plate with thirty individual thin silicon solar cell/cover samples. The cells were silicon devices fabricated by Solarex Corporation. Silver-plated Invar tabs were welded to each cell to facilitate pre and post flight electrical performance measurements. Each cell and tab assembly is bonded to a slightly oversize sheet of Kapton insulation bonded to the aluminum plate. The bonding materials were standard spacetype silicone RTVs. The covers were attached to the front surface of the cell and consist of conventional cerium doped microsheet platelets and potential candidate materials, such as FEP Teflon, silicone RTVs, glass resins, polyimides, and a silicone-polyimide copolymer encapsulant. Associated Tray(s)::Tray Location: A08 - Orientation: 38.1 degrees off ram incidence angle Photograph Classification::Flight Associated Photograph(s)::LaRC #: L90-10381 KSC #: None JSC #: S32-76-027|LaRC #: L90-13422 KSC #: KSC-390C-1031.12 JSC #: None|LaRC #: L89-04422 KSC #: KSC-384C-538.10 JSC #: None|LaRC #: L90-10403 KSC #: None JSC #: S32-77-017 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO172 Experiment Name::Effects of Solar Radiation on Glasses Investigator(s)::Whitaker, Dr. Ann F. - Invest. Role: Present|Whitaker, Dr. Ann F. - Invest. Role: Original|Kinser, Dr. Donald - Invest. Role: Present|Whitaker, Dr. Ann F. - Invest. Role: Present|Kinser, Dr. Donald - Invest. Role: Original|Whitaker, Dr. Ann F. - Invest. Role: Original|Mendenhall, Dr. Marcus - Invest. Role: Present|Nichols, Mr. Ronald - Invest. Role: Original|Nichols, Mr. Ronald - Invest. Role: Present|Wiedlocher, Mr. David - Invest. Role: Present|Tucker, Dr. Dennis - Invest. Role: Present|Weller, Prof. Robert - Invest. Role: Present Experiment Description:: Damage of glass in space systems by space debris and micrometeoroids is of interest due to the susceptibility of glass to catastrophic fracture under impact load. The Gemini window impact and simulated meteoroid impact failures in Orbiter windows indicate possible catastrophic effects of impact events in space. Lunar soil samples collected during the Apollo 11 mission reveal micrometeoroid impacts in glass spheres formed from the ejecta of larger meteoroid impacts. Impacts in such glasses exhibit central melt regions sur- rounded by fracture zones and spall areas. The influence of debris or micrometeorite impacts on mechanical properties of glass determine, in part, their sensitivity to the space environment. The deterioration of glass when subjected to solar radiation has been scientifically observed. Since glass is in a metastable state, this is not an unexpected event; the glass would achieve a lower state of energy if its atoms were rearranged in a long-range repetitious lattice structure. Changes in the properties of a glass are commonly associated with exposure to solar radiation. Because of insufficient test data for glasses exposed to actual space radiation, the materials engineer must attempt to extrapolate from data for artificial solar radiation exposure in order to select glasses for use in hardware that were exposed to the space environment for long periods of time. This limitation severely degrades the confidence level for the performance of glasses utilized in space. Objective:: The objective of this experiment was to determine the effects of solar radiation and other aspects of the space environment on glasses in space flight by exposing glass specimens to the space environment and analyzing the optical, mechanical, and chemical property changes that occur. The property changes of samples receiving differing cumulative solar radiation exposure were compared. Experiment, Approach:: This experiment was conducted by passively exposing glass samples to the space environment. Samples were exposed on LDEF in two locations: 1) the negative velocity vector, and 2) approximately perpendicular to the velocity vector facing Earth. Glass samples occupying one-sixth of a 3-in.-deep tray were located near the trailing edge of the LDEF so that they were exposed to a maximum amount of incident solar radiation. This location contained 68 cylindrical disc samples 1.25 in. in diameter. Another group of 52 samples occupying one-fourth of a 3-in. deep tray was located on the Earth-facing end of the LDEF and will receive minimum exposure to solar radiation. The properties of each sample were measured prior and subsequent to exposure to the space environment. Several samples of each glass composition selected for flight were evaluated to allow a statistical analysis of the data obtained. Compositions include aluminosilicates, fused silica, titanium silicate, lead silicates, borosilicates, soda potash lime, potash borosilicate, and soda lime silica glasses. Associated Tray(s)::Tray Location: G12 - Orientation: Earth-facing end|Tray Location: D02 - Orientation: 141.9 degrees off ram incidence angle Photograph Classification::Flight Associated Photograph(s)::LaRC #: L90-10388 KSC #: None JSC #: S32-76-054|LaRC #: L84-07165 KSC #: KSC-384C-317.08 JSC #: None|LaRC #: L91-11718 KSC #: KSC-390C-1997.02 JSC #: None|LaRC #: L84-07158 KSC #: KSC-384C-317.01 JSC #: None|LaRC #: L92-17662 KSC #: KSC-390C-731.11 JSC #: None|LaRC #: L90-10495 KSC #: None JSC #: S32-89-013 Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::S1005 Experiment Name::Transverse Flat-Plate Heat Pipe Experiment Investigator(s)::Shular, Mr. Dave - Invest. Role: Present|Owen, Mr. James - Invest. Role: Present|Owen, Mr. James - Invest. Role: Original|Edelstein, Fred - Invest. Role: Original Experiment Description:: The use of heat pipes as a thermal control system in industrial, military, and space applications has increased in recent decades. Applications cover a wide range of temperatures from cryogenic and ambient to high temperature systems, and employ working fluids such as methanol and ammonia to liquid metals. Heat pipes are employed in everyday uses such as circuit board cooling and pipeline de-icing systems, as well as tactical missiles, battlefield communications systems, and spacecraft electronics. The operational performance of heat pipes has also increased with expanded applications. Early generation heat pipes could transport 50-100 watts of power only a few feet at ambient temperatures. Currently, artery wick heat pipes can transport 1000 watts of power up to distances of 10 feet or more with temperature drops in the range of a few degrees. Heat pipes are also incorporating variable conductance features and diode functions to maintain heat source temperatures at a constant level while the heat input increases up to 200 percent or more. The application of heat pipes as a thermal control system offers many advantages over a conventional pumped fluid loop or cold plate. Due to the nature of their design, heat pipes have an inherent reliability since they possess no moving parts. Since the enthalpy of vaporization for a fluid is generally quite large, heat pipes have the capability of transporting heat energy over relatively large distances with very little temperature drop. Also, since heat pipes are driven by the pressure differences between the fluid and the vapor, no external power is required. These features, along with the variable conductance capability to provide temperature control over wide variations in source heat loads, make heat pipes a high attractive option for waste heat management of on-orbit systems and spacecraft. For a number of years, NASA Marshall Space Flight Center has actively pursued the practical application of heat pipe technology to actual thermal control hardware. A number of heat pipe concepts have been developed into breadboard hardware and extensively tested under thermal vacuum condition to verify performance. For example, programs have been successfully completed which demonstrated a deployable heat pipe radiator, transverse heat pipes, an isothermal heat pipe plate, and a total heat pipe thermal control system. In addition to these hardware programs, thermal investigations of future vehicles, such as the space station, strongly indicate the advantages of heat pipe thermal control systems. With this overwhelming support favoring heat pipe thermal control systems, future payload currently in the early design phase still revert to flight-proven thermal control techniques. This experiment offers a unique opportunity to provide flight demonstration of currently available heat pipe thermal control technology to remove the stigma from its general acceptance for space applications. A transverse heat pipe is a variable-conductance heat pipe (VCHP) which can handle relatively large thermal loads. It was developed to circumvent the gas bubble artery blockage problem associated with conventional artery wick designs which limited their capacity to small loads in the VCHP mode. In the basic design of a transverse heat pipe, liquid flows in a direction transverse or perpendicular to the vapor flow. Temperature control is achieved by using conventional noncondensible-gas techniques. The concept of this investigation was to utilize current basic heat pipe technology to design and fabricate a heat pipe thermal control module experiment, demonstrate the hardware capability and performance in the Shuttle flight environment, and verify the ground verses flight data correlation. It was anticipated that the self-regulated transverse flat-plate heat pipe would maintain the temperature control areas of the experiment within the tolerance specified with varying heat inputs independent of LDEF orientation. Objective:: The purpose of this experiment was to evaluate the zero-g performance of a number of transverse flat-plate heat pipe modules. Performance included the systems heat transport capability of the pipes, the temperature drop between evaporator and condensor, and the ability to maintain temperature over varying duty cycles and environments. Additionally, performance degradation was evaluated over the length of the LDEF mission. This information is necessary if heat pipes are to be considered for system designs where they offer benefits not available with other thermal control techniques. The objective of this experiment was to demonstrate that a self-regulated heat pipe thermal control system can maintain hardware tem- peratures within tolerance of +/- 9 degrees F or less with varying heat inputs, independent of vehicle/payload orientation. Experiment, Approach:: The Transverse Flat Plate Heat Pipe is a thermal control device that serves the dual function of temperature control and mounting base for electronic equipment. In its ultimate application, the heat pipe would be a lightweight structural member that could be configured in a platform or enclosure, and provide temperature control for large space structures, flight experiments, and equipment. Grumman Aerospace Corporation, the manufacturer of the heat pipe modules for LDEF experiment S1005, was granted a patent for the transverse concept. Originally configured in a cylindrical geometry, it was developed to reduce the problem of gas bubble artery blockage found in conventional artery wick designs. This blockage by generated non-condensible gases limits the operating capacity in the variable conductance mode to small loads. In the transverse flat plate design, liquid is returned to the evaporator from the condensor in a direction perpendicular (i.e. transverse) to the vapor flow. Variable conductance temperature control was achieved by using non-condensible gas (nitrogen) which was contained in a reservoir separate from the heat pipe module. Three transverse flat-plate heat pipe modules were installed in a 12-in.-deep peripheral tray. Heat was supplied to the evaporator side of the module by a battery power supply that simulated various watt density equipment heat dissipators. This heat from the condensor region of the heat pipe modules was radiated to space from the outboard-facing radiator surface of the modules. Pretimed heater duty cycles provided load inputs at discrete mission times. Thermocouple data recording the performance of the heat pipes was stored on magnetic tape for analysis after retrieval of the experiment. The entire experiment was self-contained with respect to power supply, data storage, and on-orbit cycling. An experiment power and data system (EPDS) was used for the data recording on magnetic tape and LiSO2 batteries provided EPDS power. Heat was supplied to the evaporator side of the test modules through electrical strip foil heaters which simulate heat rejection by electronic equipment and were bonded directly to the interior surface of the modules. The batteries and other components were not utilized to supply heat to the modules because it was desired to be able to accurately control the module environment and vary the heat loads to allow a more detailed verification of the experiment capacity. Heater power was provided by 28-V lithium monofluorographite batteries; the particular types were Eagle Picher MAP 9036. Three identical experiment ON times were planned during the mission. Each ON time lasted approximately 13 hours (8.6 orbits), with heat input to the modules. Each ON period was subdivided into two 4.3-orbit heat input periods to verify proper operation of each module. The initial ON occurred approximately 1 month after launch, the second on time 67.5 days later, and the third 67.5 days after that. The three identical ON periods at different times of the mission should allow identification of any performance changes during orbital lifetime. Associated Tray(s)::Tray Location: B10 - Orientation: 21.9 degrees off ram incidence angle Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L92-17793 KSC #: KSC-390C-612.02 JSC #: None|LaRC #: L90-10444 KSC #: None JSC #: S32-78-076|LaRC #: None KSC #: None JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO175 Experiment Name::Evaluation of Long-Duration Exposure to the Natural Space Environment on Graphite-Polyimide and Graphite-Epoxy Mechanical Properties Investigator(s)::P.O.well, Howard - Invest. Role: Original|Welch, Douglas - Invest. Role: Original|Vyhnal, Dr. Richard - Invest. Role: Present Experiment Description:: Graphite-polyimide and graphite-epoxy are two composite materials being used in current spacecraft construction, and both materials are being considered for more extensive use in future lightweight space-oriented structural compo- nents. The accumulation of operational data on the effects of long-duration exposure of these two materials to the multiple environmental elements of space is needed to properly evaluate them for both current and future applications. In particular, data are needed on the mechanical properties of these materials. Objective:: The primary objective of this experiment was to accumulate the needed opera- tional data associated with the exposure of graphite-polyimide and graphite-epoxy material to the environments of space. Secondary objectives for testing the graphite-polyimide materials were to evaluate laminar microcracking and crack propagation and to eliminate any concerns associated with unknowns. A specific objective of testing the graphite-epoxy material was to validate the mechanical-property knock-down factors that were applied to the design and analysis of the Space Shuttle payload bay door. The assessment of the degree of matrix cracking and crack propagation phenomena resulting from differential expansion of unlike materials coupled with large thermal excursion, and the deletion of unknowns resulting from simultaneous application of multiple environmental factors relative to the payload bay door composite and adhesive system, were secondary objectives in the graphite-epoxy tests. Experiment, Approach:: LDEF Experiment AO175 involves the non-instrumented exposure of seven carbon-fiber-reinforced resin-matrix advanced composite panels contained in two trays - A07 and A01. These two trays were located, respectively, on the leading and trailing faces of LDEF. The experiment was mounted in two 3-in.-deep peripheral trays. Graphite-polyimide specimens occupied 1 1/3 trays and the graphite-epoxy specimens occupied two-thirds of a tray. The seven panels consisted of six flat laminates of the following composite material systems: carbon/epoxy (T300/934), carbon/bismaleimide (T300/F178), and carbon/polyimide (C6000/LARC-160 and C6000/PMR-15), plus one bonded honeycomb sandwich panel (T300/934 facesheets and Nomex core) patterned after the Space Shuttle payload bay door construction. These material systems were selected to represent a range of available matrix resins which, by virtue of their differing polymer chemistry, could conceivably exhibit differing susceptibility to the low Earth orbit (LEO) environment. The experiment approach required two matched sets of specimens with traceable records that were maintained for materials processing and specimen quality. After fabrication, one set of each test specimen was sectioned and structurally tested to serve as a data baseline. After the LDEF flight, the other set of specimens underwent extensive measurements of mechanical properties for comparison with the original data baseline. Structural testing of the graphite-polyimide specimens provided strength and elastic data in tension, compression, and shear. Transverse tension microcracking and crack propagation were evaluated by photomicroscopy. Structural testing of the graphite-epoxy specimens included verification of laminate, core, adhesive, and fatigue properties as applied to the design and analysis of the payload bay door. Microcracking and crack propagation will also be analyzed by photomicroscopy. Associated Tray(s)::Tray Location: A07 - Orientation: 68.1 degrees off ram incidence angle|Tray Location: A01 - Orientation: 111.9 degrees off ram incidence angle Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L92-17828 KSC #: KSC-390C-609.02 JSC #: None|LaRC #: L84-07083 KSC #: KSC-384C-221.01B JSC #: None|LaRC #: L90-10430 KSC #: None JSC #: S32-78-029|LaRC #: L92-17551 KSC #: KSC-390C-725.04 JSC #: None|LaRC #: L90-10398 KSC #: None JSC #: S32-76-095|LaRC #: L84-07071 KSC #: KSC-384C-209.01 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::S1006 Experiment Name::Balloon Material Degradation Investigator(s)::Strganac, Dr. Thomas - Invest. Role: Present|Allen, David - Invest. Role: Original Experiment Description:: It has been observed that exposure of polymers to the low Earth orbit (LEO) environment can result in significant degradation due primarily to atomic oxygen attack. LEO lies between 200 and 500 km above the Earth's surface and has an atmosphere which consisted predominately of atomic oxygen. Spacecraft at this particular orbit travel at a rate of 8 km/s which has the effect of providing the atomic oxygen with a translational energy of approximately 5 eV as it strikes ram facing (direction of travel) surfaces. Under these condition many polymers are degraded with resulting mass loss. With the increasing importance of polymers in orbiting spacecraft it has become imperative to determine how, and to what extent, the properties of polymers are affected by this type of an environment. Knowing this polymers can subsequently be developed or selected which are suitable for LEO applications. This degradation is synergistically increased in the presence of high levels of ultraviolet (UV) radiation. To date, the large body of knowledge associated with the use of polymers in LEO has probed the chemical mechanisms associated with atomic oxygen attack and UV exposure. Although it is recognized that the presence of atomic oxygen and UV radiation alters the chemical integrity of polymers, it was not know to what extent these chemical alterations affect the mechanical behavior of the material. Data on the effects of the space environment on very thin polymeric films are needed to properly support NASA programs involving the flight of extremely high altitude scientific balloons. In particular, significant scientific benefit will be derived from the development of a long-duration balloon platform capable of carrying payloads on the order of 250 kg to altitudes greater than 40 km for periods in excess of 60 days. The National Scientific Balloon Facility has actively pursued this program in the past. However, the engineering of these large systems could be significantly accelerated if data regarding degradation and/or alteration of various material properties could be obtained and compared to laboratory simulations of the space environment. LDEF was essentially a free-flying cylindrical structure developed by NASA to accommodate 57 totally self-contained experiments on trays mounted on the exterior of the structure. The extended duration of the LDEF mission significantly enhanced the opportunities to characterize the morphological and mechanical properties of exposed polymers. The balloon materials exposure experiment by Texas A&M, S1006, consisted of 38 polymer film specimens and 24 fibrous cord specimens. The location of these samples (polyethylene, polyester, and nylon films in addition to Kevlar fibers) on the LDEF was fortuitous because they were exposed to the smallest amount of direct sunlight of any other experiment row aboard the satellite. In addition, the samples were placed such that the normal from the surface of the samples was nearly perpendicular to the ram direction of the atomic oxygen flux, thereby limiting the atomic oxygen fluence of the specimens. The finding of this research contributes to the predictive models of material constitutive characteristics. Objective:: The objective of this experiment was to assess the effects of long-term exposure of candidate balloon films, tapes, and lines to the hostile environment above the Earth's atmosphere. Degradation of mechanical and radiometric properties were observed by a series of tests on exposed materials. Experiment, Approach:: The experiment was passive and tested candidate balloon films, tapes, and lines. The experiment occupied one-third of a 3-in.-deep peripheral tray. Two additional identical sets of material were prepared. The first set was tested immediately and the second was held in a controlled environment until the recovery of the samples placed in orbit. Tests were then performed on this second set to determine any effects of aging. The specimens that were recovered from the LDEF were also tested and the effects of long-duration exposure noted. In addition to these specimens, another set of specimens were exposed to the Texas A & M University accelerated exposure facility and the results were compared with those specimens exposed in situ. The balloon materials exposure experiment consisted of 38 thin film polymer samples and 24 fibrous cord samples. The thin-film polymer samples ranged from 0.35 mil to 1.8 mil in thickness and were primarily constructed of polyester and polyethylene. Some of the thin-film specimens were reinforced with nylon, Kevlar, or polyester fibers. These polymeric materials were intended for use in long-duration scientific balloons and, except for the laminates and composite films, were manufactured as a flown film. This manufacturing process is known to introduce a biaxial orientation to the film resulting in anisotopic mechanical properties. Hence, pairs of specimens with mutually perpendicular orientation were included in the experiment package to account for directionality. Each non-reinforced specimen was six inches long and one inch wide. Each reinforced film was 6 in. long and 1.5 in. wide. Subsequent to exposure, two types of tests were performed on each specimen. The films were subjected to uniaxial state of stress at room temperature at -80 degrees C with a constant strain rate of 0.2 percent per minute. The load and deformation were recorded as a function of time and stress-strain diagrams were prepared. Five specimens of each film type were used to insure repeatability. After detailed elastic data were taken, the film was loaded to failure. It is anticipated that significant chemical changes would occur which affect the effective absorptivity and emissivity properties of the film. Therefore, care was taken not to clean or otherwise alter the surface of the exposed films. Associated Tray(s)::Tray Location: E06 - Orientation: 98.1 degrees off ram incidence angle Photograph Classification::Postflight Associated Photograph(s)::LaRC #: L90-13439 KSC #: KSC-390C-1033.09 JSC #: None|LaRC #: L90-10454 KSC #: None JSC #: S32-82-006|LaRC #: L84-07068 KSC #: KSC-384C-193.10 JSC #: None Spacecraft flown on::LDEF-1 - Long Duration Exposure Facility Experiment's Assigned Number::AO178 Experiment Name::A High-Resolution Study of Ultra-heavy Cosmic-Ray Nuclei Investigator(s)::Domingo, Vicente - Invest. Role: Original|O'Ceallaigh, Cormac - Invest. Role: Original|O'Sullivan, Dr. Denis - Invest. Role: Original|Wenzel, Dr. Klaus-Peter - Invest. Role: Present|Wenzel, Dr. Klaus-Peter - Invest. Role: Original|Thompson, Dr. Alex - Invest. Role: Present|Thompson, Dr. Alex - Invest. Role: Original|O'Sullivan, Dr. Denis - Invest. Role: Present Experiment Description:: Prior to LDEF there were only two spacecraft which carried experiments dedicated to the investigation of ultra-heavy nuclei. HEAO-3 and Ariel 6 were launched in 1979 and employed electronic detectors of geometric factor 5 square meter sr and 2 square meter sr respectively. The combined sample from both missions with Z greater than or equal to 65 comprises approximately 300 events, and the entire sample of actinides (Z greater than or equal to 88) is only 3. The experiment on LDEF which was dedicated to the study of ultra-heavy (UH) nuclei consisted of an extensive array of primarily Lexan polycarbonate solid state nuclear track detectors, of geometric factor 30 square meter sr, which were mounted with cylindrical aluminum pressure vessels in 16 LDEF experiment trays. The measurement of the charge spectrum of ultraheavy cosmic-ray nuclei is of vital importance in many areas of astrophysics. Such measurements are relevant to the study of the origin and age of cosmic rays, acceleration and propagation mechanisms, the nucleosynthesis of the heavy elements in our galaxy, and the search for superheavy nuclei (Z greater than 110), and may even lead to improved nuclear mass formulas and beta decay rate formulas for the heaviest nuclides. The central theme of this experiment was the utilization of the LDEF's large area-time factor to obtain a large and uniform sample of ultraheavy cosmic-ray nuclei in the region Z greater than or equal to 30. Objective:: The main objective of the experiment was a detailed study of the charge spectra of ultra-heavy cosmic-ray nuclei from zinc (Z = 30) to uranium (Z = 92) and beyond using solid-state track detectors. Special emphasis was placed on the relative abundances in the region Z greater than or equal to 65, which is thought to be dominated by r-process nucleosynthesis. Subsidiary objectives included the study of the cosmic-ray transiron spectrum and a search for the postulated long-lived superheavy (SH) nuclei (Z greater than or equal to110), such as SH-294 (Z= 110), in the contemporary cosmic radiation. The motivation behind the search for superheavy nuclei is based on predicted half-lives that are short compared to the age of the Earth but long compared to the age of cosmic rays. The detection of such nuclei would have far-reaching consequences for nuclear structure theory. The sample of ultraheavy nuclei obtained in this experiment will provide unique opportunities for many tests concerning element nucleosynthesis, cosmic-ray acceleration, and cosmic-ray propagation. For example, if the r-process domination for Z greater than 65 is confirmed, a reliable source spectrum will provide details of the nuclear environment, such as temperature and time scale. This information is of great importance to both astrophysics and nuclear physics. The relative abundances of cosmic-ray nuclei in the region Z greater than 82 will lead to a determination of the age of cosmic rays directly from the decay of a primary component, in contrast to estimates based on, for example, Be-10 or Cl-36. The LDEF exposure was anticipated to provide a sufficiently large sample of actinides to achieve this objective. Injection of cosmic-ray particles into the acceleration process may depend upon atomic properties of the elements, such as the first ionization potential. This experiment will help to establish the existence of such mechanisms through their modulation of the ultraheavy charge spectrum. The cosmic-ray charge spectrum in the region (30 greater than or equal to Z greater than or equal to110), based on the statistics available to date, appears to be generally consistent with solar system material. It is hoped that a radical improvement is achieved with the LDEF exposure. For example, it is very important to establish the roles played by the helium-burning slow-neutron-capture process in massive stars and the explosive carbon-burning process during supernova explosions. Since the LDEF orbit inclination was low, the geomagnetic cutoff prevented direct measurement of the ultraheavy energy spectrum. However, it is hoped that the slope of the energy spectrum can be determined from an analysis of the abundance distributions about the major axis of the LDEF (east-west effect). To summarize, cosmic rays constitute a unique sample of material from distant parts of our galaxy which still bears the imprint of the source region. The ultraheavy cosmic-ray composition will provide a great deal of information about the evolution of matter in the universe. This question is closely related to understanding the origins of the elements in the solar system. Experiment, Approach:: The experimental approach centered on the use of solid-state track detectors to identify charged cosmic-ray particles. The basic detector component is a thin sheet of plastic (typically 250 um thick). Since the primary objective of the UHCRE experiment was to study ultra-heavy cosmic ray nuclei of Z less than 65 and the geomagnetic cut-off is approximately 1 GeV/N, the main detector material chosen was Lexan polycarbonate. The determination of charge and velocity depends on the mechanism by which cosmic-ray nuclei the penetrate the plastic sheets produce radiation damage along the particle trajectories. After exposure and recovery, the detector sheets were chemically processed to reveal the tracks produce by the passage of heavily ionizing particles. The effective amplification of a particle's radiation damage trail results from preferential chemical etching along its trajectory. The rate of etching is a unique function of the particle's ionization. The flux of ultraheavy cosmic-ray nuclei is extremely small (on the order of 1 per meter per day). Consequently, the basic experimental approach entailed a large area-time factor exposure coupled with stability in a general radiation environment and the ability to discriminate against the overwhelming background of lower charge components. The main plastic used for the track detectors in this experiment was Lexan polycarbonate (following the discovery by the Dublin group of Taffak polycarbonate as a track detector for heavy cosmic rays, this detector material has also been incorporated), which has a registration threshold given by Z/beta approx. equal to 56. It is this threshold property that enables the Lexan to isolate individual ultraheavy nuclei in a very high flux environment of lower charge nuclei. The Ultra Heavy Cosmic Ray Experiment is the largest array of cosmic ray particle detectors to be flown in space. The nuclear track detectors, with lead foil energy degraders, were assembled in stacks that were mounted in aluminum cylinders designed to fit into 12-in.-deep peripheral trays. The experiment was based on a modular array of 192 solid state nuclear track detector stacks, mounted in sets of four in 48 pressure vessels (at 1 atm of dry air). Three cylinders, each containing four stacks, were placed parallel to the x-axis of each tray. The cylinders were approximately 46 in. long and approximately 10 in. in diameter, and have a wall thickness of approximately 0.5 g/sq cm. All cylinders, except one, were pressurized to 1.0 bar with a dry oxygen-nitrogen-helium mixture in the ratio of 20:70:10. The stacks had a thickness of approximately 4 g/sq cm and were mounted parallel to the tray base and placed symmetrically about the main axis of each cylinder using an Eccofoam matrix. Each stack contains mainly Lexan polycarbonate layers (approximately 70 plates) interleaved with several sheets of lead. The lead sheets act both as electron strippers and velocity degraders and were chosen because of their low cross section for nuclear interactions. The trays were thermally decoupled from the LDEF frame and will carry thermal covers flush with their outer rims. The LDEF UHCRE employed sixteen side viewing trays. Associated Tray(s)::Tray Location: A10 - Orientation: 21.9 degrees off ram incidence angle|Tray Location: B05 - Orientation: 128.1 degrees off ram incidence angle|Tray Location: B07 - Orientation: 68.1 degrees off ram incidence angle|Tray Location: C05 - Orientation: 128.1 degrees off ram incidence angle|Tray Location: C06 - Orientation: 98.1 degrees off ram incidence angle|Tray Location: C08 - Orientation: 38.1 degrees off ram incidence angle|Tray Location: A02 - Orientation: 141.9 degrees off ram incidence angle|Tray Location: A04 - Orientation: 158.1 degrees off ram incidence angle|Tray Location: C11 - Orientation: 51.9 degrees off ram incidence angle|Tray Location: D01 - Orientation: 111.9 degrees off ram incidence angle|Tray Location: F04 - Orientation: 158.1 degrees off ram incidence angle|Tray Location: D05 - Orientation: 128.1 degrees off ram incidence angle|Tray Location: D07 - Orientation: 68.1 degrees off ram incidence angle|Tray Location: D11 - Orientation: 51.9 degrees off ram incidence angle|Tray Location: E02 - Orientation: 141.9 degrees off ram incidence angle|Tray Location: E10 - Orientation: 21.9 degrees off ram incidence angle Photograph Classification::Prelaunch Associated Photograph(s)::LaRC #: L84-07162 KSC #: KSC-384C-317.05 JSC #: None|LaRC #: L90-10414 KSC #: None JSC #: S32-77-055|LaRC #: L90-13475 KSC #: KSC-390C-1066.06 JSC #: None|LaRC #: L90-13349 KSC #: KSC-390C-832.02 JSC #: None|LaRC #: L90-10406 KSC #: None JSC #: S32-78-023