SP-473 LDEF

 

Materials, Coatings, and Thermal Systems (1/2)

 


[8] Growth of Crystals From Solutions in Low Gravity (A0139A)

M. David Lind
Rockwell International Science Center
Thousand Oaks, California
 
Kjeld F. Nielsen
Technical University of Denmark
Lyugby, Denmark

 

Background

These crystal growth experiments are an extension of preliminary experiments performed during the Apollo-Soyuz Test Project flight and similar experiments being developed for a Spacelab flight. The crystal growth method to be used consists 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 flights. Ideally, this type of crystal growth process requires that a low level of gravity (less than 10-4g) be maintained for a period of weeks or months. At present the LDEF flights are the only space flights planned which can fully satisfy this requirement.

 

Objective

The objective of these experiments is to develop a novel solute diffusion method for growing single crystals. Crystals to be investigated are PbS, CaC03, and TTF-TCNQ. Each of these materials has current research and technological importance. PbS is a semiconductor, and CaCO3 has useful optical properties; both would have many applications if they 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 are 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 [9] should make possible better determinations of their physical properties and. perhaps, investigations of possible device applications. The experiment; should also contribute to a better understanding of the theory and mechanism, of crystal growth.

 

Approach

The experiments will utilize specially designed reactors (fig. 5) with three or more compartments separated by valves to keep the reactant solutions and solvent separated until the apparatus reaches low gravity. There will be a mechanism for opening the valves automatically to initiate the diffusion and growth processes. The reactant reservoirs will be large enough to take advantage of the time provided by the LDEF flight. An array of several reactors will be mounted in a 12-in.-deep end center tray located on the Earth-facing end of the LDEF. Several reactors operating simultaneously will allow experimentation with more than one crystal growth system and/or variations of conditions for each. The reactors will be enclosed in a vacuumtight container and will be surrounded by thermal insulation. The temperature (approximately 35°C) will be regulated and any departures from the desired temperature will be recorded. Power requirements will be provided by LiSO2 batteries.

 


[
9-10]

Figure 5.-Crystal growth experiment.
(a) Reactor configuration.

(a) Reactor configuration.

Figure 5.-Crystal growth experiment.
(b) Tray configuration shown without thermal cover.

(b) Tray configuration shown without thermal cover.

Figure 5.-Crystal growth experiment.

 


[11] Atomic-Oxygen-Stimulated Outgassing (A0034)

Robert L. Scott, Jr.
Southern University
Baton Rouge, Louisiana
 
Roger C. Linton
NASA George C. Marshall Space Flight Center
Huntsville, Alabama

 

Background

Many materials (e.g., thermal control surfaces) are known to produce outgassed products and possible particulate contamination when exposed to a space environment. This 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.

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 surfaces, 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 near-Earth orbit is a major factor in producing optically damaging outgassed products. The expected results will be 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.

 

[12] Approach

Selected thermal control surfaces will be exposed to the atomic oxygen in the near-Earth orbit. Passive collecting samples will collect any induced outgassing resulting from the oxygen impingement. The optical condition of the passive samples will be measured using a ground-based integrating sphere reflectometer and a directional reflectometer.

Two packages are required, each occupying one-sixth of a 3-in.-deep tray. (See fig. 6 ) One package will be positioned on the leading edge and one on the trailing edge of the LDEF. The thermal control surfaces used on Skylab, as well as newly developed surfaces, will be contained in the packages. The atomic oxygen will impinge on the thermal control surfaces contained in the leading-edge container. Passive specular collecting samples will be 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, will be exposed to the available ultraviolet radiation so that environment synergistic effects can be observed. Preexposure and postexposure analysis will include total hemispherical and bidirectional reflectance measurements. The range of these reflectance measurements will be 2.5 to 2500 Å. Postexposure analysis of the leading-edge samples exposed to the atomic oxygen will be compared with the control samples positioned on the trailing edge and shielded from the atomic oxygen.

 


[
13]

Figure 6.-Atomic-oxygen-stimulated experiment shown with top cover removed.

Figure 6.-Atomic-oxygen-stimulated experiment shown with top cover removed.

 


[14] Interaction of Atomic Oxygen With Solid Surfaces at Orbital Altitudes (A0114)

John C. Gregory
The University of Alabama
Huntsville. Alabama
 
Palmer N. Peters
NASA George C. Marshall Space Flight Center
Huntsville, Alabama

 

Background

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 poorly understood, and at this time it is not possible for a spacecraft designer either to allow for them or to disregard them with impunity. This experiment is 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 which are usually designed to point or remain fixed in space. The attitude stabilization mode of LDEF results in the same surface always being presented to the ambient atmospheric flux along the velocity factor, while the opposite surface of the vehicle remains in a hard vacuum. Also, since the LDEF is passive. the contamination problems encountered on spacecraft such as Skylab from venting or leaking fluids do not exist.

 

Objectives

The objectives of this experiment are 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.

 

[15] Approach

The basic approach to this experiment, as previously stated, is to expose a wide variety of material surfaces to the atomic flux in orbit. The experiment is passive and depends on preflight and postflight measurements of the test surfaces in the laboratory. The experiment will also include 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 will be mounted in a panel, as shown in figure 7. The face of this panel will be flown on LDEF normal to the incident stream of oxygen atoms. Each disk will have part of its front surface masked so exposure to the atomic-oxygen reaction will be limited to selected areas. the shadowed areas being used as control surfaces in the measurements.

A typical sample is an optically flat quartz disk overcoated with a film of the material of interest. These include 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, will be solid disks rather than thin films.

The experiment consists of two flight units. Each unit occupies 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 will not be subjected to atomic-oxygen impingement and will serve as control samples. A third set of control samples will be 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 will also be contained in an experiment exposure control canister (EECC) located with LDEF experiment S0010, Exposure of Spacecraft Coatings.

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.

 

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 accommodation aspect, three cylindrical ''reflectometers'' will also be included with this experiment. (See fig. 8.) Slits in the panels and cylinders will....

 


[
16]

Figure 7.-Photographs of interaction of atomic-oxygen experiment showing the removable small panel which is painted black to provide a hotter surface than the main surface which is painted white.

Figure 7.-Photographs of interaction of atomic-oxygen experiment showing the removable small panel which is painted black to provide a hotter surface than the main surface which is painted white.

 

[17] ....permit a beam of atomic oxygen to impinge on a given sample at a selected angle of incident. For the first flight of the reflectometers, the objectives are limited to testing the concept of the silver film detectors and distinguishing between specular and cosine law reflectance at the sample surface. Materials chosen include LiF, stainless steel, and aluminum.

 

Spacecraft Attitude Sensor

A unique passive spacecraft attitude sensor has been 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 is designed to measure the angular offset of LDEF from its nominal flight attitude.

 


Figure 8.- Cylindrical reflectometer configuration.

Figure 8.- Cylindrical reflectometer configuration.

 

[18] As a subexperiment provided by Dr. Gerald J. Fishman of NASA Marshall Space Flight Center, a number of activation metal samples will he included with the other samples previously mentioned. After exposure to the space environment. these samples will become slightly more radioactive due to ambient proton and neutron irradiation. upon recovery, the radioactivity will be carefully analyzed by the NASA MSFC Space Sciences Laboratory to provide measurements of the average proton and neutron fluence during the LDEF mission.

 


[19] Influence of Extended Exposure in Space on Mechanical Properties of High-Toughness Graphite-Epoxy Composite Material (A0019)

David K. Felbeck
University of Michigan
Ann Arbor, Michigan

 

Background

Graphite-epoxy composites are promising candidates for structural use in space vehicles because of their high strength and elastic modulus properties. The problem of low fracture toughness has also been solved by use of recently developed techniques of intermittent interlaminar bonding. Before this material can be adapted for space use. however, confidence must he gained that its mechanical properties are not degraded by exposure to the space environment.

 

Objective

The objective of this experiment is 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 recently developed techniques of intermittent interlaminar bonding will be exposed and afterward tested for fracture toughness, tensile strength, and elastic modulus.

 

Approach

The approach of this experiment is to provide a frame on which the specimens can be mounted with their flat sides normal to the LDEF radius, each specimen with an unobstructed exposure of about 2Greek letter pi sr. The specimens will be 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 is passive, nothing is required except the mechanical and thermal anchoring of the test specimens. There will be six fracture toughness specimens and nine tensile modulus specimens utilizing one-sixth of a 3-in.-deep peripheral tray. (See fig. 9.) An identical set of specimens produced at the same time will be stored in the ground laboratory for final testing at the same time as the orbited specimens. A third set of specimems produced at the same time will be tested a short time after fabrication and.....

 


[
20]

Figure 9.-High-toughness graphite-epoxy composite material experiment.

Figure 9.-High-toughness graphite-epoxy composite material experiment.

 

....curing. These three sets of specimens will 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 are tested to fracture, those of special interest will be 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.

 


 

[21] Effect of Space Environment on Space-Based Radar Phased-Array Antenna (A0133)

Richard J. Delasi, Martin L. Rossi, James B. Whiteside, Martin Kesselman, Ronald L. Heuer, and Frederick J. Kuehne
Grumman Aerospace Corporation
Bethpage, New York
 

Background

Large space structures of low areal density are currently being developed for near-term applications such as space-based radar (SBR). 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, in all likelihood, must be polymeric. However, 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.

Based on the performance of numerous polymeric materials following accelerated laboratory testing, Kapton polyimide film has been selected as the baseline material for the Grumman SBR concept. To gain the requisite confidence for long-term service durability, it is desirable to subject material specimens as well as a portion of the SBR antenna directly to the combined space environment and compare property degradation to that caused by laboratory simulation.

 

Objective

The overall objective of this program is to evaluate the effect of the space environment on polymeric materials currently being considered for the Grumman SBR Phased-Array Antenna. Degradation mechanisms caused by thermal cycling, ultraviolet and charged-particle irradiation, applied load and high-voltage plasma interaction will be evaluated.

 

Approach

The experiment occupies a 6-in.-deep end comer tray located on the space end of LDEF and consists of both passive and active parts. The passive [22] part addresses 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 10-3 to 10-4 in./in. will be measured on 10-in.-long specimens. The specimen array shown in figure 10 contains 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 will undergo limited thermal excursions and will be exposed to minimal solar and charged-particle radiation. Each specimen contains a bonded splice located so that both spliced and continuous Kapton regions can be tested after exposure. Four stress levels (30, 150, 300, and 450 psi) 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.

The active part of the experiment addresses the issue of the interaction between high voltage and low-Earth-orbit plasma. A 14- by 28-in. section of the Grumman SBR antenna (two Kapton antenna planes and a perforated aluminum ground plane) has been selected as the test specimen. The electrodes provided by copper dipole elements deposited on the Kapton plane will be held at I and 2 kV. A counting circuit and recently available Grumman-designed microprocessor memory using EEPROM's (a nonvolatile, electrically erasable, programmable read-only memory) will be used to record the number of electrical discharges. The experiment timer-sequencer delays the application of high voltage for 16 days and powers up the memory subsystem every 20 minutes for 0.6 second. Special testing circuitry has been included to assure proper circuit operation, and mission time scale speedup capability has been included to permit simulation of the entire flight.

 


[
23]

Figure 10.-Space-based radar phased-array antenna experiment.

Figure 10.-Space-based radar phased-array antenna experiment.

 


[24] Space Exposure of Composite Materials for Large Space Structures (A0134)

Wayne S. Slemp
NASA Langley Research Center
Hampton, Virginia

 

Background

As space systems become large, more complex, and expensive, they will 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 is 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.

 

Approach

The experiment is passive and occupies about one-half of a 6-in.-deep peripheral tray, as shown in figure 11. Specimens of composite materials and polymeric and resin films are arranged above and below the experiment....

 


[
25]

Figure 11.-Space exposure of composite materials experiment shown integrated with Exposure of Spacecraft Coatings experiment (50010).

Figure 11.-Space exposure of composite materials experiment shown integrated with Exposure of Spacecraft Coatings experiment (50010).

 

....mounting plate to enable both exposure and nonexposure to sunlight. This provides 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 is thermally isolated from the LDEF structure to allow the material specimens to experience a wide range of thermal cycles.

Tensile and compression specimens will be used to evaluate the laminated composite materials. A number of the specimens are precut and ready for testing after space exposure, whereas other specimens will be prepared from larger samples. Both O.005-in.- and 0.003-in.-ply thicknesses of prepreg (resin-impregnated material) will be used. The tensile specimens will be fabricated with ±45° layup. The effects of flight exposure will be evaluated by determining the stress-strain and ultimate tensile strength before and after flight exposure. Table I summarizes the specific composite materials to be evaluated. Metal matrix composites are 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) will be used to provide additional data on the behavior of polymers in space. Data will be 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 will also be performed before and after flight exposure.

 

[26] Table 1.-A0134 Composite Specimens.

Resin and/or metal

Reinforcement

.

Epoxy, 5208

Graphite, T-300

Epoxy, 934

Graphite, T-300 (0.003 in./ply)

.

Graphite, T-300 (0.005 in./ply)

Polysulfone, P-1700

Graphite, Celion 3000 (0.003 in./ply)

.

Graphite, Celion 6000 (0.005 in./ply)

Epoxy, 930

Graphite, GY-70

Mg

Graphite, P-100

Al

Graphite, P-100

A series of laboratory tests will be conducted on all materials to simulate LDEF space flight conditions. The laboratory test program will include one set of specimens exposed to UV radiation at one solar constant in 10-7 torr vacuum for 1000 hours and another set of specimens exposed to vacuum for the duration of LDEF flight exposure. These specimens will provide data to help isolate the effects of ultraviolet light, vacuum, and time on the flight specimens.

 


[27] Effect of Space Exposure of Some Epoxy Matrix Composites on Their Thermal Expansion and Mechanical Properties (A0138-8)

Robert Elberg
Space Division' Matra S.A.
Le Chesnay, France

 

Background

Carbon and Kevlar fiber-reinforced plastic composites are being used increasingly in space structures (e.g., launch vehicles, spacecraft, payload elements such as antennas and optical benches). This extensive interest in composites is due on one hand 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 wishes to examine is 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.

 

Objectives

This experiment has three objectives. The first and main objective is to detect a possible variation in the coefficient of thermal expansion of composite samples during a l-year exposure to the near-Earth orbital environment. A second objective is to detect a possible change in the mechanical integrity of composite products, both simple elements and honeycomb sandwich assemblies. A third objective is to compare the behavior of two epoxy resins commonly used in space structural production.

 

Approach

The experimental approach is to passively expose samples of epoxy matrix composite materials to the space environment and to compare preflight [28] and postflight measurements of mechanical properties. The experiment will be located in one of the three FRECOPA (French cooperative payload) boxes in a 12-in.-deep peripheral tray that contains nine other experiments from France. (See figs. 12 and 13.) The FRECOPA box will protect the samples from contamination during the launch and reentry phases of the LDEF mission.

A list of the samples to be tested and their composition is given in table 2. Two identical samples of each type are foreseen and four different configurations of samples are used, as described in figure 14. Two common characteristics of configurations A, B, and C are the length (4 in.) and the existence at each end of the samples of three protrusions. These are the reference points for the measurement of the coefficient of thermal expansion (CTE).

The coefficients of thermal expansion are measured on Earth before and after space exposure. This measurement is based on a laser interferometry method working in a vacuum. The method used consists 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 monomode He-Ne laser beam. Length variation due to temperature variation results in a fringe motion, which is measured. CTE measurements are made at ambient temperature with a temperature variation of about 20°C. The accuracy for a sample length of 4 in. is about ±0.1 x 10-6 per °C.

 


Figure 12.-Layout of French cooperative payload (FRECOPA).

Figure 12.-Layout of French cooperative payload (FRECOPA).

 


[
29]

Figure 13.-Photograph of FRECOPA box (open).

Figure 13.-Photograph of FRECOPA box (open).

 

The test program consists of the following measurements.

Weight measurements of all the samples will be made before launch (samples previously dried and outgassed) and after launch. Any weight variation will be useful information to establish the efficiency of the drying and outgassing process for composite materials.

The coefficient of thermal expansion will be 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 l-year exposure at ultrahigh vacuum will be detected.

Micrographic inspections will be 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 will be compared to similar views of identical materials kept on Earth for reference.

Mechanical tests for interlaminar shear strength, flexural strength, and flatwise tensile strength will be made on elements cut from the samples after return to Earth. These test results will be compared to results from similar samples that were not submitted to the space environment. Although vacuum chamber tests of epoxy matrix composites show good mechanical and thermal [30] stability, this experiment is expected to increase confidence in the performance of the tested composite specimens, particularly relative to the thermal stability.

 

Table 2.- List and Composition of Epoxy Matrix Composite Samples

Reference Number

Sample Type (Shape referencea)

Fiber-Resin Material

Additional Material

Fiber Arrangement

.

1-1'

Circular tube (A)

GY70/934 Unidirectional, eb = 0.005 in.

None

4 ply (4.0°/ ±35°/4.0°)

2-2'

Sandwich: aluminum honeycomb, CFRP face sheets (B)

GY70/934 Unidirectional, e = 0.005 in.

Aluminum honeycomb, bond film, 312

(0°/ ±60°/0°/90°/±42°)s

3-3'

Rectangular tube (C)

GY70/934 Unidirectional, e = 0.005 in.

None

(0°/ ±60°/0°/90°/±42°)s

4-4'

Rectangular tube (C)

GY70/934 Unidirectional, e = 0.005 in.

None

5-5'

Rectangular tube (C)

GY70/V108 Unidirectional, e = 0.005 in.

None

6-6'

Sandwich: Kevlar honeycomb, Kevlar face sheets (D)

Kevlar/V108, e = 0.0075 in.

Kevlar honeycomb

2 ply (0°/90°)

a Refer to figure 14.

b e = fiber diameter.

 


[
31]

Figure 14.- Experiment A0138-8 sample configurations.

Figure 14.- Experiment A0138-8 sample configurations.

 


[32] The Effect of the Space Environment on Composite Materials (A0138-9)

Michel Parcelier
Aerospatiale
Les Mureaux, France

 

Background

It is the duty of a manufacturer to verify the characteristics of the products he designs and manufactures. Until now, the space industry has complied with this rule with laboratory simulations and evaluations. The launching of the LDEF by the Space Shuttle will provide an opportunity to observe the actual behavior of materials under exposure to the space environment and will make it possible to correlate with artificial aging tests.

 

Objective

The objective is to test different types of materials (laminates, thermal coatings, and adhesives) to determine their actual useful lifetime. These experiments will also make it possible to integrate the histories of the thermal and mechanical characteristics into models of the composite structures.

 

Approach

The experiment is passive and is located in one of the FRECOPA boxes in a 12-in.-deep peripheral tray with nine other experiments from France. (See figs. 12 and 13.) The FRECOPA box will provide protection for the samples from contamination during the launch and reentry phases of the LDEF mission. The experiment revolves around four themes of study: thermal coatings, adhesives, dimensional stability, and mechanical characteristics.

The various materials will be arranged in six levels within the FRECOPA box, so only the first level will be subjected to direct solar radiation.. (See fig. 15 and table 3.) Each level will consist of plates from which test specimens will be cut after the mission.

 


[
33]

Figure 15.-Section of FRECOPA box showing arrangement of experiment A0138-9.

Figure 15.-Section of FRECOPA box showing arrangement of experiment A0138-9.

 

Table 3.- Composite Materials Test Specimens

Level

Test specimens

.

1A

OSR and SSM on graphite-epoxy composite

1B

OSR and SSM on aluminum support

1A

Sandwich (GY70/code 87 + BSL 312L)

1B

Sandwich (GY70/BSL 914 + BSL 319L)

2

BSL 312L and BSL 319 L specimens

2A

Redux 408 on graphite-graphite support

2B

Redux 312 L on graphite-aluminum support

3A, 4A

GY70/code 87

3B, 4B

GY70/BSL 914

5A, 6A

T300/V108

5B

GY70/V 108

6B

G837/V 108

Thermal Coatings

Optical solar reflectors and second-surface mirrors will be laid on aluminum and carbon supports. The tests will evaluate the level of degradation, the mass of contaminants received, and the alteration of the thermo-optical properties.

The tests will measure the shear in a joint bonded with 408 (room-temperature curing) and 312 L (high-temperature curing) adhesives. The purpose is to observe the effects of the thermal stresses due to the assembly of [34] materials having different expansion coefficients (carbon and aluminum) and the thermal cycling in the low orbit of the LDEF.

 

Dimensional Stability

Tests will be carried out to verify the predicted thermoelastic deformations in sandwich structures that have withstood the space environment. Expansion tests will be 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 are GY70/87 (0° and 90° orientation) and 312 L. The sandwich will be structured as shown in figure 16. The same experiment will be carried out on a sandwich test specimen cocured with 914 and 319 L.

 


Figure 16.-Structure of sandwich specimen.

Figure 16.-Structure of sandwich specimen.

 

The degradation of the mechanical properties (tensile, flexural, and interlaminar tests) will be evaluated on the following materials: GY 70/87 (tape), GY70/914 (tape), T300/V108 (tape), and G837/V 108 (fabric).

 


[35] Microwelding of Various Metallic Materials Under Ultravacuum (A0138-10)

Jean Pierre Assie
Aerospatiale
Cannes, France

 

Background

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. Such microwelds result in a continuing increase in the friction factors and are even liable to jam the mechanisms altogether.

 

Objective

The objective of this experiment is to check 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.

 

Approach

The experimental approach is 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 experiment will be located in one of the FRECOPA boxes in a 12-in.-deep peripheral tray that contains nine other experiments from France. (See figs. 12 and 13.) Table 4 lists the materials to be tested and table 5 and figure 17 show the test sample arrangement and experiment layout.

 

[36] Table 4.-Experiment A0138-10 Test Materials [Shape: round bars]

Type of material

Description

Condition

Applicable standard

.

Aluminum alloy

AZ5GU (7075)

T 7351

ASN A 3086

AU4G1 (2024)

T 351

ASN A 3058

A5 (1050)

H 24

NA

Copper alloy

CuBe1.9

TH2

ASN A 3416

CuNi35i (UN 3S)

TF

ASN A 3405

Titanium alloy

TA6V

Annealed

ASN A 3307

TA6V

Quenched and tempered

ASN A 3306

Stainless steel

EZ6NCT25

Quenched for 960 MPa

ASN A 3412

Z100CD17 (440C)

Treated

ASN A 3376

Z6CNT18/11 (321)

Hyperquenched

ASN A 3140

 


Figure 17.-Layout of the eight washer <<spools>> within frecopa box.

Figure 17.-Layout of the eight washer "spools" within FRECOPA box.

 

Table 5.- Description and Combination of Constituent Materials of Washers.

1

2

3

4

.

EZ6NCT25

CuBe1.9

AZ5GU a

AZ5GU

EZ6NCT25

CuBe1.9

AZ5GU a

AZ5GU

AZ5GU a

AZ5GU

CuBe1.9

CuNi3Si

TA6V annealed b

TA6V annealed b

CuNi3Si

CuNi3Si

AZ5GU a

AZ5GU

CuBe1.9

CuNi3Si

TA6V annealed

TA6V annealed

EZ6NCT25

EZ6NCT25

AZ5GU a

AZ5GU

TA6V annealed b

TA6V annealed

CuBe1.9

CuBe1.9

CuBe1.9

CuBe1.9

AZ5GU a

AZ5GU

TA6V annealed b

TA6V annealed

CuNi3Si

CuNi3Si

CuNi3Si

CuNi3Si

AZ5GU a

AZ5GU

TA6V tempered b

TA6V annealed

EZ6NCT25

EZ6NCT25

EZ6NCT25

EZ6NCT25

5

6

7

(Analysis of influence of pressure)

8

(Analysis of influence of lubrication)

.

TA6V annealed b

Z1OOCD17

P1

CuBe1.9

CuBe1.9/Molykote Z

TA6V annealed b

Z1OOCD17

.

CuBe1.9

CuBe1.9

TA6V annealed

Au on AZ5GU c

P2

CuBe1.9

CuBe1.9/Molykote Z

TA6V annealed

Au on AZ5GU c

.

CuBe1.9

MoS2 d/AZ5GU a

Au on AZ5GU c

Z6CNT18/11

P3

CuBe1.9

CuBe1.9/Molykote Z

Au on AZ5GU c

Z6CNT18/11

.

CuBe1.9

MoS2 d / TA6V annealed b

Ag on AZ5GU c

A5

P4

TA6V annealed

TA6V annealed b/ MoS2 d

Ag on AZ5GU c

A5

.

TA6V annealed

TA6V annealed b

Cr on AZ5GU

AU4G1 a

P5

TA6V annealed

AU4G1 a / MoS2 d

Cr on AZ5GU

AU4G1

.

TA6V annealed

TA6V annealed b

Au on AZ5GU c

TA6V quenched

P6

TA6V annealed

AU4G1 a / MoS2 d

Ag on AZ5GU c

TA6V quenched

.

TA6V annealed

AU4G1 a

a Chromium anodized.
b Sulfur anodized.
c Au and Ag are reference couples.
d Physical vapor deposition.

 
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