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