[38] Evaluation of Long-Duration Exposure to the Natural Space Environment on Graphite-Polyimide and Graphite-Epoxy Mechanical Properties (A0175)

J. Howard Powell and Douglas W. Welch

Rockwell International Corporation

Tulsa, Oklahoma

Background

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 components. 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.

Objectives

The primary objective of this experiment is to accumulate the needed operational 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 are to evaluate laminar microcracking and crack propagation and to eliminate any concerns associated with " unknowns. " A specific objective of testing the graphite-epoxy material is to validate the mechanical-property "knock-down" factors that were applied to the design and analysis of the Space Shuttle payload bay doors. The assessment of the degree of matrix cracking and crack propagation phenomena resulting from differential expansion of unlike materials coupled with large thermal excursions, and the deletion of unknowns resulting from simultaneous application of multiple environmental factors relative to the payload bay door composite and adhesive system, are secondary objectives in the graphite-epoxy tests.

Approach

The experiment will be mounted in two 3-in.-deep peripheral trays. Graphite-polyimide specimens will occupy 1 1/3 trays and the graphite-epoxy specimens will occupy two-thirds of a tray, as shown in figure 18.

The experiment approach requires two matched sets of specimens with traceable records that are maintained for materials processing and specimen....

[40] ....quality. After fabrication, one set of each test specimen will be sectioned and structurally tested to serve as a data baseline. After the LDEF flight, the other set of specimens will undergo extensive measurements of mechanical properties for comparison with the original data baseline. Figure 19 illustrates the various specimen configurations.

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Structural testing of the graphite-polyimide specimens will provide strength and elastic data in tension, compression, and shear. Transverse tension microcracking and crack propagation will be evaluated by photomicroscopy.

Structural testing of the graphite-epoxy specimens will include 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.

[41] The Effect of Space Environment Exposure on the Properties of Polymer Matrix Composite Materials (A0180)

R. C. Tennyson and J. S. Hansen

Institute for Aerospace Studies, University of Toronto

Downsview, Ontario, Canada

Background

The use of polymer matrix composites in various spacecraft applications is increasing, but the effects of long-term space exposure on the mechanical properties of these materials are not known Although laboratory simulation using a thermal-vacuum chamber can be employed, the correlation between these results and actual in situ behavior has not been established. Consequently, such a correlation should be made in order to provide the design engineer with appropriate correction factors to take into account any degradation of material integrity due to various exposure times in space. Failure to do so will undoubtedly lead to structural failure resulting from material property changes. This can be particularly hazardous when using polymer matrix composites because molecular breakdown (due to radiation), outgassing (due to vacuum), arid internal cracking, accompanied by fiber matrix separation and delamination (from thermal cycling), can result.

Objective

The objective of this experiment is to determine the effect of various lengths of exposure to a space environment on the mechanical properties of selected commercial polymer matrix composite materials. Fiber materials will include graphite, boron, S-glass, and PRD-49. The mechanical properties to be investigated are orthotropic elastic constants, strength parameters (satisfying the tensor polynomial relation), coefficients of thermal expansion, impact resistance, crack propagation, and fracture toughness. In addition, the effect of laminate thickness on property changes will also be investigated.

Approach

Five groups of test articles made up of laminated cylindrical tubes and flat plates will be manufactured from a given batch of prepreg (resin-impregnated) material. This will be done for each material system selected. One group from each material system will be evaluated under ambient laboratory conditions to determine (1) the orthotropic elastic constants E₁₁, E₂₂, v₁₂, and G₁₂, where E is the modulus of elasticity, v is Poisson's ratio, [42] and G is the shear modulus; (2) the tensor polynomial failure parameters F_j, F_ij, and F_ijk; (3) the impact resistance (residual bending strength as a function of impact energy from a projectile; (4) the crack propagation and fracture toughness (measurement of growth of a given crack size in the specimen as a function of temperature and load cycling); and (5) the coefficients of thermal expansion for various laminate configurations. This evaluation will be repeated with a second group of specimens subjected to thermal-vacuum exposure in a laboratory facility. An evaluation of the effects of ultraviolet and electron beam radiation and atomic oxygen impingement will be included. The evaluation will also be repeated with a third group of specimens subjected to actual space environment onboard the LDEF. Finally, two control batches will be evaluated to assess the effects of storage environment, qualifying tests, and aging on LDEF flight articles.

The experiment will occupy one-half of a 3-in.-deep peripheral tray and is divided into three sections. (See fig. 20. ) Each section consists of a layered arrangement of both tubular and flat specimens. The tubular specimens are thin walled (0.02 to 0.06 in.) and approximately 1.75 in. in diameter and 4 in. .....

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[43] ....long. The flat specimens are of similar thickness and measure 2 in. wide and 5 in. Iong Aluminum end fittings (which function as test fixtures following space exposure) are bonded to each of the test specimens.

Additionally, thermal and strain gauge outputs on various composite specimens and a stainless steel calibration specimen will be sampled simultaneously approximately every 16 hours during the flight. This information will be transferred to a cassette deck for serial recording within a sealed electronics box beneath one of the experiment sections. (See fig. 21.) Two LiSO₂ batteries are required for data recording power and are located beneath the remaining two experiment sections.

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[44] Space Environment Effects on Spacecraft Materials (M0003)

Paul Schall

The Aerospace Corporation

El Segundo, California

Background

Data on materials for unmanned NASA and DOD spacecraft have been valuable, but are limited to those items that can be monitored remotely. Causes of failures or performance degradation can often only be inferred from the telemetry data. The Gemini and Apollo missions included some materials experiments. The return of components from the Surveyor III lunar lander was particularly interesting. The NASA Skylab missions contained a thermal control materials experiment; however, the early Skylab problems resulted in contamination that affected results. Although data from unmanned spacecraft will continue to be used to eval uate the performance of materials in space, the LDEF adds a new dimension to space experiments.

The LDEF provides experimenters with am opportunity to recover specimens that have been exposed for long periods in space. The typical approach for the selection of materials for use in spacecraft has involved laboratory testing and limited measurements in space. Although many materials appear to be satisfactory for a variety of applications, there is insufficient knowledge of the physical and optical properties of these materials after long periods in space. Laboratory tests do not simulate the actual space environment; therefore, it is difficult to predict property changes as a function of environment exposure.

In addition to measuring changes in the macroscopic properties of the returned specimens, microstructural properties will also be examined. Thus, it may be possible to increase our understanding of the changes induced by the environment. This increased understanding can then be used to predict the performance of materials based on knowledge of the space environment and the results of laboratory tests.

This experiment will be a cooperative effort and will provide an opportunity for DOD space programs and laboratories to evaluate materials and components after long exposures to the space environment.

Objectives

The immediate objectives of this experiment are to understand changes in the properties and structure of materials after exposure to the space environment and to compare these changes with predictions based on [45] laboratory experiments. The longer term objectives are to improve the performance and usage of existing materials and to decrease the lead times for application of new materials on DOD space systems.

Approach

The experiment consists of 19 subexperiments involving a number of DOD laboratories and contractor organizations. (See table 6.) In general, the experimental approach with each of the subexperiments will involve comparison of preflight and postflight analyses. Typical analyses will include the measurement of optical properties (reflectance, transmittance, and refractive index), macrophysical properties, and microstructural properties.

The experiment consists of four peripheral trays, two experiment power and data systems, two experiment exposure control canisters, and LiSO₂ batteries to satisfy power requirements. The trays and EECC's will be used to [47] retain a variety of thermal control coatings, composites, laser optics electronic piece parts, fiber optics, solar cells, and LDEF experiment M0002-1.

A 6-in.-deep tray, a 3-in.-deep tray, a data system, and an experiment exposure control canister will be located near the LDEF leading edge with the trays connected by a wiring harness. A similar configuration will be located near the LDEF trailing edge. Environmental exposure to the two locations will be similar except that the leading-edge location will also be exposed to relatively high fluxes of atmospheric constituents (primarily atomic oxygen). Figure 22 shows photographs of two of the experiment trays.

The experiment is equipped to record temperature, strain, and solar-cell output voltage. These data will be recorded approximately every 107 hours (approximately 78 orbits) for the duration of one orbit. The EPDS will be programmed to record data periodically over a span of up to 15 months. Both EECC units will be programmed to open in three stepped increments to vary the UV exposure times. The first opening will occur approximately 10 days after deployment, to minimize contamination. The second stepped opening will occur at approximately one-third of the expected minimum flight duration. The third stepped opening will occur at approximately two-thirds of the expected minimum flight duration. Prior to LDEF retrieval, the EECC units will close to provide protection from contamination during retrieval operations.

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The fiber optics experiment will investigate the resistance of a fiber optic cable to the effects of the space environment and will involve the illumination of a l-km fiber by a light-emitting diode (LED) source. A detector will be used to monitor the radiance of the LED and a second detector will monitor the output of the fiber. The difference between the two detectors will represent the loss in the fiber. The fiber will be unsheathed so that radiation effects can be introduced. The LED source, detectors, and fiber will be electrically and optically characterized prior to the test. All composites will be subjected to postflight analysis for comparison to the initial data. Failure analysis will be used as required to identify failure modes and/or mechanisms and indicate potential solutions to identified problems. The detectors will be matched to minimize the difference between the individual detector characteristics.

[49] Balloon Materials Degradation (S1006)

David H. Allen

Texas A. & M. University

College Station, Texas

Background

There exists a need to expose a variety of very thin films to the space radiation environment in order to gain sufficient data to properly support other 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 for the past 3 years. 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.

Objective

The objective of this experiment is 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 will be observed by a series of tests on exposed materials.

Approach

The experiment is passive and will test candidate balloon films, tapes, and lines. The experiment will occupy one-third of a 3-in.-deep peripheral tray, as shown in figure 23. The materials to be tested are listed in table 7. Two additional identical sets of material will be prepared. The first set will be tested immediately and the second will be held in a controlled environment until the recovery of the samples placed in orbit. Tests will then be performed on this second set to determine any effects of aging. The specimens that are recovered from the LDEF will also be tested and the effects of long-duration exposure noted. In addition to these specimens, another set of specimens will be exposed to the Texas A. & M. University accelerated exposure facility and the results will be compared with those of specimens exposed in situ.

Subsequent to exposure, two types of tests wilt be performed on each specimen The films will be subjected to a uniaxial state of stress at room...

Table 7.- Balloon Materials Specimens

Films:

0.5-mil nylon 12 TD^a

0.5-mil nylon 12 MD^b

1.0-mil Stratofilm TD

1.0-mil Stratofilm MD

0.5-mil Stratofilm TD

0.5-mil Stratofilm MD

1.0-mil SFX TD

1.0-mil SFX MD

0.5-mil SFX TD

0.5-mil SFX MD

0.35-mil SFX TD

0.35-mil SFX MD

0.35-mil aluminized polyester

0.92-mil Hostaphan 2000

0.48-mil to .48-mil polyester with 400-denier Kevlar 29

0.92-mil to .92-mil laminated Hostaphan 2000

Tapes:

Nylon-reinforced polyester (500 lb)

Kevlar-reinforced polyester (1000 lb)

Pressure-sensitive adhesive tape (polyethylene substrate with reinforcing polyester backing and silicon adhesive)

Lines:

Nylon (500 lb)

Kevlar (500 lb)

Kevlar (1000 lb)

^a TD = transverse direction

^b MD = meridional directional

[51] ....temperature and at - 80°C with a constant strain rate of 0.2 percent per minute. The load and deformation will be recorded as a function of time and stress-strain diagrams will be prepared. Five specimens of each film type will be used to insure repeatability. After detailed elastic data have been taken, the film will be loaded to failure.

It is anticipated that significant chemical changes will occur which will affect the effective absorptivity and emissivity properties of the film. Therefore, care will be taken not to clean or otherwise alter the surface of the exposed films.

[52] Thermal Control Coatings Experiment (A0138-6)

A. Paillous

CERT/ONERA-DERTS

Toulouse, France

 

J-C. Guillaumon

CNES/CST

Toulouse, France

Background

In order to assess the degradation of thermo-optical properties of coatings used on satellites, a space environment simulation is needed. 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 degradations obtained in the laboratory with degradations obtained in space would be very valuable.

Objectives

The objectives of this experiment are 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, and to compare the behavior in space of some materials for which the available ultraviolet solar simulation is inadequate (especially in the far ultraviolet).

Approach

The experimental approach is to passively expose samples of the thermal coatings of interest. These coatings include black paint, aluminum paint, white paint, a solar absorber, an optical surface reflector, second-surface mirrors, metal coatings, and silica fabrics. Preflight and postflight measurements of thermo-optical properties will be compared to determine the effects of space environment exposure.

The experiment will be located with nine other experiments from France in a 12-in. -deep peripheral tray. The thermal coating samples will be housed in one of the three FRECOPA boxes located in the tray. (See fig. 12.) The [53] FRECOPA box (fig. 13) will protect the samples from contamination during the launch and reentry phases of the LDEF mission.

Samples will be independently maintained in sample holders that allow their front face to receive maximum solar illumination when the FRECOPA box is open. (See fig. 24.) Thirty samples will be tested. Twenty-nine samples are 3/4 in. by 3/4 in. and one sample is 1 1/2 in. by 1 1/2 in. Sample thickness is less than 1/8 in. The maximum temperature during space exposure will be recorded by passive temperature indicators fixed to the sample mounting plate. Additionally, the ionizing radiation dose will be measured by a passive LiF dosimeter.

The FRECOPA box will be closed in space after exposure and will be kept under vacuum until optical measurements are completed in the laboratory. The entire closed box containing samples under vacuum will be placed in a vacuum chamber, where the optical reflectance spectrum of each sample will be recorded using an integrating sphere. An additional set of samples will be maintained in the laboratory for comparison with samples subjected to space exposure.

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[54] Exposure of Spacecraft Coatings (S0010)

Wayne S. Slemp

NASA Langley Research Center

Hampton, Virginia

Background

The degradation of thermal control coatings due to space radiation exposure has caused spacecraft to overheat, leading to problems with subsystems and mission lifetimes. To prevent such problems, designers need to be able to accurately predict the performance of thermal control coatings. Several flight experiments have been conducted to obtain the necessary coating performance data. Unfortunately, these data were limited to telemetry information and the experiments were not returned for postflight evaluation. Coating performance was determined from temperature measurements made on sample coatings and increases in sample temperature were interpreted as being caused by space radiation. With these experiments it was not possible to distinguish between damage caused by space radiation and that caused by some other means, such as mechanical stress or contamination. To properly isolate the cause of coating degradation, an experiment is necessary which provides for the return of coating samples after space exposure for ground laboratory evaluation.

Objectives

The objectives of this experiment are to determine the effects of both the Shuttle-induced environment and the space radiation environment on selected sets of spacecraft thermal control coatings.

Approach

The experimental approach is to passively expose samples of thermal control coatings to Shuttle-induced and space radiation environments and to return the samples for postflight evaluation and comparison with preflight measurements to determine the effects of the envirommental exposure. Two additional sets of samples will remain in the laboratory and will be analyzed for comparison with the flight data. Optical measurements of the samples will include total normal emittance and spectral reflectance.

The experiment will utilize a 6-in.-deep peripheral tray and an experiment exposure control canister (EECC). (See fig. 25.) The EECC will provide protection for some of the samples against exposure to the launch and...

...reentry environments. The EECC will be programmed to open about 2 weeks after LDEF deployment and close prior to LDEF retrieval by the Shuttle and reentry.

Some samples will not be housed in the EECC and will be exposed to the Shuttle-induced environment during launch and reentry. Comparison of the data from these samples with data from samples in the EECC exposed to only the space radiation environment will yield information about possible contamination-induced degradation effects.

Table 8 provides a list of the thermal control coatings that will be included in this experiment. Additionally, several materials specimens will be located in the EECC as control specimens for other LDEF experiments (e.g., A0114 and A0187).

[57] Thermal Control Surfaces Experiment (S0069)

Donald R. Wilkes and Harry M. King

NASA George C. Marshall Space Flight Center

Huntsville, Alabama

Background

The optical properties of thermal control surfaces in the solar region of the spectrum are of primary interest to spacecraft thermal designers since these properties govern the solar-heat input to exposed surfaces (such as the thermal radiators) and therefore influence the temperature of the spacecraft. These properties, however, have been shown to be altered considerably under the space environment, which includes solar irradiation, thermal vacuum, micrometeoroid bombardment, and contamination. One such mechanism of solar ultraviolet degradation is caused by photodesorption of oxygen, which is immediately and completely reversible upon exposure to a very small amount of oxygen ( 10^-4 to 10^-6 torr partial pressure). This type of bleaching mechanism shows the necessity of in situ measurements of the optical properties of environmentally damaged surfaces (i.e., in vacuum before repressurization).

Until now, no optical measurements of thermal control surfaces have been made in space. Temperature measurements of thermally isolated samples have been used to back-calculate solar absorptance and thermal emittance. This type of measurement is not as definitive as required and does not describe the spectral character of the sample surface. Spectral reflectance measurements of the samples are required to differentiate between different damage mechanisms of environmental effects and to separate contamination effects.

Additionally, because of the inability to simulate exactly the conditions of the coating surface temperature and the solar spectrum, there is a major difference between laboratory test data and in-flight experiment data. Although the current generation of laboratory test apparatus is extremely complex and well thought out, it provides only relative data on the degradation of coatings in actual space conditions. The only accepted test for flight qualification of new coatings is to have them evaluated in actual conditions of space flight in the space environment where they will be used.

Objectives

The objectives of this experiment are to determine the effects of the near-Earth orbital environment and the Shuttle-induced environment on [58] spacecraft thermal control surfaces. Spectral reflectance measurements will be obtained and used to differentiate between different solid-state damage mechanisms of environmental damage, to separate the effects of contamination from those of natural-environmental damage, and for comparison and correlation with laboratory test data.

Approach

The experiment is designed to measure certain physical properties of 25 "active" spacecraft thermal surface samples in an environment that approximates their normal use. The parameters to be measured include the hemispherical reflectance as a function of wavelength ( 100 wavelength steps from 0.25 to 2.5 µm) and the temperature of these samples as a function of time in a calorimeter configuration. The latter measurements will be made in two different physical configurations that allow calculation of the emittance and the ratio of solar absorption to emittance for each sample. In addition, 24 passive samples will be exposed to approximately the same environment as the active samples.

Figure 26 shows a simplified block diagram of the experiment, figure 27 shows the experiment layout in a 12-in.-deep tray, and figure 28 is a photograph of the flight hardware. The active samples are contained in calorimeter assemblies and are mounted along with the passive samples on the carousel. In addition, three radiometers (solar and Earth albedo, Earth albedo, and earthshine) are also mounted on the carousel. The radiometers are used to measure the radiant energy incident upon the samples, which is required for calculating the ratio of absorption to emittance, and to provide a record of the total exposure of the samples to the solar ultraviolet.

The carousel has two fixed positions, referred to as in IN, or protected, and OUT, or exposed. The OUT position exposes the samples to the environment. The samples are in this position approximately 23 1/2 hours for every Earth day, including the 1 1/2-hour period each day when temperature and radiometer measurements are being recorded to determine the ratio of absorption to emittance. The carousel is rotated 180° from the OUT position to the IN position for the emittance measurements for approximately 1/2 hour each day. For these measurements, the samples view a massive heat sink (aluminum "emittance" plate) which maintains a relatively constant temperature, and temperature change as a function of time is recorded for each sample.

The IN, or stowed, position also places the samples and radiometers in a protected enclosure for launch and reentry. This position is also maintained for 10 days after launch to allow volatiles to outgas prior to starting experimental operations.

[59] The reflectometer assembly includes an integrating sphere, which is located at the bottom of the carousel assembly. The carousel is rotated by a stepping motor through a geneva drive mechanism to position each of the 25 active samples in the integrating sphere aperture, where sample reflectance can be measured. Each sample will be measured 20 times during the LDEF mission, nominally once per month with measurements slightly more often near the beginning of the mission.

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[62] Ion-Beam-Textured and Coated Surfaces Experiment (S1003)

Michael J. Mirtich, Jr.

NASA Lewis Research Center

Cleveland, Ohio

Background

Future spacecraft relying on thermal control surfaces or solar thermal power generation will be subjected to the near-Earth Shuttle environment prior to insertion into a geosynchronous orbit. The combined effects of the near-Earth Shuttle environment may be synergistic and may cause appreciable degradation prior to geosynchronous-orbit operations. In situ exposure of various candidate surfaces is required to evaluate material, optical, and/or electrical property durability so that a choice of surface materials can be made with respect to optical and or electrical performance, durability, and contamination protection requirements.

Objective

The objective of this experiment is to measure the effects of exposure to the Shuttle launch and near-Earth space environments on the optical properties of ion-beam-textured high-absorptance solar thermal control surfaces, the optical and electrical properties of ion-beam-sputtered conductive solar thermal control surfaces, and the weight loss of ion-beam-deposited oxide-polymer films.

The various types of surfaces to be tested include six major categories: (1) ion-beam textured surfaces suitable for space solar-thermal (solar concentrator) application (e.g., materials such as copper, aluminum, Inconel, stainless steel, and silver); (2) painted and or state-of-the-art solar thermal surfaces (e.g., black chrome); (3) ion-beam-sputtered conductive coatings for thermal and space charge control (e.g., indium-oxide-coated metalized FEP Teflon); (4) ion-beamsputtered conductive coated solar-sail materials for space charge control and cooling through emittance (e.g., sputtered coatings on Kapton such as indium oxide, aluminum, and chromium); (5) micrometeoroid-sensitive samples whose optical properties change only as a result of micrometeoroid impact; and (6) Kapton coated with oxide-polymer films to minimize oxygen degradation at near-Earth-orbit altitudes.

The objective for the first two categories of samples is to verify that the optical properties of the microscopic cone or ridge-type ion-beam-textured surfaces are more resistant to degradation than conventional solar thermal surfaces. The objective for the third and fourth categories of samples is to evaluate the electrical and optical durability of conductive coatings for thermal control and solar-sail radiative cooling applications. The objective for the [63] fifth-category sample is to identify changes in the optical properties which can be attributed to micrometeoroid impact. The objective for the sixth category is to measure any changes in optical or material properties of oxide-polymer-coated Kapton after exposure to the oxygen atom environment in near-Earth (Shuttle) orbit.

Approach

The experimental approach is to passively expose the samples to all environments of the entire mission. The optical properties (absorptance and emittance) of each surface will be measured in ground tests both before and after exposure to the environment. This will be done by experimentally measuring the spectral reflectivity between 0.33 and 2.16 µm using a Gier-Dunkle integrating sphere to obtain the solar absorptance. The emittance will be obtained by measuring the spectral reflectance in the infrared between 1.5 to 15.5 µm using a Holraum reflectometer.

Electrical conductive coatings will be resistance documented before and after the LDEF flight. Comparisons will be made between the durability of the painted surfaces and the ion-beam-textured or sputtered surfaces. Additional tests, including weight loss. Auger and SEM measurements and or chemical analyses may also be performed as the data warrants.

The experiment requires one-sixth of a 3-in.-deep peripheral tray. Figure 29 illustrates the experiment configuration and table 9 lists the samples that will be tested.

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[66] Cascade Variable-Conductance Heat Pipe (AOO 76)

Michael G. Grote and Leslie D. Calhoun II

McDonnell Douglas Astronautics Company

St. Louis, Missouri

Background

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 will provide this capability, but its performance capabilities have not been adequately demonstrated in flight.

Objective

The objective of this experiment is 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.

Approach

The approach to conducting the experiment is consistent with the LDEF capabilities (i.e., relatively long duration, zero gravity environment, and minimal electrical power and data system capability). Solar energy is the heat source and space the heat sink for thermally loading two series-connected variable conductance heat pipes. Electronics and power supply equipment requirements are minimal. The experiment power data system (EPDS) in LDEF experiment S1001 (Low-Temperature Heat Pipe Experiment Package (HEPP) for LDEF) will be used for data recording. A 7.5-V lithium battery supplies the power for thermistor-type temperature sensors for monitoring system performance, and a 28-V lithium battery supplies power for valve actuation.

The experiment will occupy a 6-in.-deep peripheral tray located on the leading edge of the LDEF. Two external-surface subpanels will be employed which are thermally coupled to opposite ends of two series-connected variable conductance heat pipes. One panel will be designed as a heat absorber through application of a high /(absorptivity/emissivity) surface coating, and the second will be a radiator with a low / surface coating. Multilayer insulation and fiberglass structural attachments are used to thermally isolate the experiment from the LDEF tray structure and interior. Each of the heat pipe evaporators will be maintained within preselected temperature ranges by sizing the collector, [67] radiator, and insulation and by servicing the noncondensible dry gas reservoirs, thus demonstrating passive variable-conductance heat pipe operation.

Figure 30 shows the CVCHP experiment configuration, which uses two gas-loaded, dry-reservoir VCHP's in series. The coarse-control heat pipe temperature is controlled to plus or minus 3°C and is used as a sink for the fine-control heat pipe. The dry gas reservoir temperatures are controlled by locating the reservoirs next to the heat pipe evaporators. The principal concern is drift in the set point temperature due to many heat load and or environment temperature cycles. The cyclic operation will move vapor into or out of the...

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[68] .... noncondensible gas reservoir, which changes the set point temperature by introducing a varying working-fluid partial pressure into the gas reservoir. The capillary tube is located between the heat pipe and the reservoir to prevent working fluid from entering the reservoir. The diameter and length of the capillary are 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, shown in figure 31, has a single-pedestal artery with seven tubelets enclosed in a sheath of two layers of 400-mesh screen. The sheath provides the high capillary pumping pressure and the tubelets provide high permeability to reduce the axial pressure drop. The wall wick has a 200-mesh outer layer for low pressure drop and a middle layer of 400 mesh for capillary pumping.

Ammonia is used as the working fluid for both heat pipes. The coarse control VCHP has a reservoir-to-condensor volume ratio of 20, which will yield...

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[69] ....a control band of plus or minus 3°C. A larger 90-to- I volume ratio is used on the fine-control heat pipe to attain a control of plus or minus 0.3°C. The fine-control VCHP has 8.5 m of capillary, and the course-control VCHP has a larger 13.4 m section because of its larger condenser volume.

Experiment operation begins when the battery circuit is initiated at LDEF deployment from the Shuttle. When the temperature of the fine-coarse heat exchanger falls below - 7°C, a valve opens to permit initiation of the coarse control VCHP and allow the heat pipes to prime prior to beginning VCHP operation. This temperature-controlled initiation ensures that heat pipe temperatures are below their operating set point so that vapor is not forced into the reservoirs. Twenty-five hours later, another valve opens to initiate the fine control VCHP operation. This delayed opening permits data to be taken on a stabilized coarse-control VCHP before fine-control VCHP operation. Data will be collected at least twice daily during the LDEF mission and will be stored on magnetic tape for subsequent retrieval and playback.

[70] Low-Temperature Heat Pipe Experiment Package (HEPP) for LDEF (S1001)

Roy Mclntosh, Jr., and Stanford Ollendorf

NASA Goddard Space Flight Center

Greenbelt, Maryland

Craig R. McCreight

NASA Ames Research Center

Moffett Field, California

Background

Experience gained in the development of heat pipes has demonstrated the necessity of obtaining performance data in the space environment. This is due to the fact that the pumping in a heat pipe is derived from relatively weak capillary forces. As a result, particularly in the case of the axially grooved geometry, reliable l-g performance measurements are often very difficult to obtain. Also, many of the candidate low-temperature and cryogenic fluids have relatively low surface tensions and wicking heights, which compound the problem of l-g tests.

Objectives

The principal objectives of the experiment are to determine zero-g start-up performance for conventional and diode low-temperature heat pipes, to evaluate heat pipe performance in zero-g for an extended period of time, to determine zero-g transport capability of each heat pipe, and to determine diode operation, including forward conductance, turndown ratio, and transient behavior.

Approach

The heat pipe experiment package (HEPP) is designed to test the performance of low-temperature (< 190 K) heat pipes on the Long Duration Exposure Facility (LDEF). Two heat pipes, a fixed-conductance transporter heat pipe and a thermal-diode heat pipe, are coupled with a radiant cooler system.

Both pipes are charged with ethane. Also integrated with the radiator is a phase change material (PCM) canister which provides temperature stability during transport tests. N-heptane, which has a melting/freezing point of 182 K, is used as the PCM. The high heat capacity (28 W-hr of latent heat) provided by the canister permits high-power heat pipe testing (e.g., 40 W for 40 minutes) at constant temperature. Multilayer insulation blankets are employed, and a shielding [71] configuration was developed to minimize parasitic inputs from the Earth and maximize radiation to deep space.

HEPP is a completely self-contained and thermally isolated package designed to fit in a 12-in.-deep peripheral tray, as shown in figure 32. The necessary electrical equipment, including electronics for signal conditioning sequencing and command functions, will also be contained within the experiment tray. A standard LDEF experiment power and data system (EPDS) will be used for data collection and recording. Power for tine experiment will be provided by a dedicated solar-panel Ni-Cd battery system in another 12-in. -deep tray (fig. 33). This tray will be located on the space-facing end of the LDEF to take maximum advantage of Sun input for experiment operation.

After completion of the LDEF mission, the recorded data will be unpacked and converted to engineering units. The existing thermal model will be used to analyze and correlate the flight data. The flight data will be used to establish boundary conditions; the thermal program will then determine individual heat pipe heat flows and conductances, diode shutdown energy, and PCM performance. Data tabulations and plots will be generated and compared with preflight predictions and thermal-vacuum test results. Performance results derived from the analysis will be reported following preflight and postflight tests.

Additionally, as a result of Kapton erosion seen on Shuttle flights, control samples have been added to some of the trays as an atomic-oxygen coatings investigation to determine ways of protecting Kapton (polyimide) film from atomic-oxygen degradation. Several specimens (a Kapton control, Kapton with coatings of In₂0₃, urethane-acrylic, and silicones) will be taped to Kapton film (14 in. square) using Kapton-backed pressure-sensitive tape. This sheet will then be taped to the HEPP tray Kapton blanket. A duplicate set of specimens will be similarly taped to the CVCHP blanket. In the case of the power tray, a lesser number of specimens will be taped directly to the metal lip of the tray.

[74] Transverse Flat-Plate Heat Pipe Experiment (S1005)

James W. Owen

NASA George C. Marshall Space Flight Center

Huntsville, Alabama

 

Fred Edelstein

Grumman Aerospace Corporation

Bethpage, New York

Background

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 conditions 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 payloads 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 is 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 versus flight data correlation. It is anticipated that the self-regulated transverse flat-plate heat pipe will maintain the temperature control areas of the experiment within the tolerance specified with varying heat inputs independent of LDEF orientation.

[75] Objective

The objective of this experiment is to evaluate the zero-g performance of a number of transverse flat-plate heat pipe modules. Performance will include the transport capability of the pipes, the temperature drop, and the ability to maintain temperature over varying duty cycles and environments. Additionally, performance degradation, if any, will be monitored 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.

Approach

As shown in figure 34, three transverse flat-plate heat pipe modules will be installed in a 12-in.-deep peripheral tray. Heat will be supplied to the evaporator side of the module by a battery power supply that will simulate various watt density equipment heat dissipators. This heat will be radiated to space from the outboard-facing radiator surface of the modules. Pretimed heater duty cycles will provide load inputs at discrete mission times. Thermocouple data recording the performance of the heat pipes will be stored on magnetic tape for analysis after retrieval of the experiment. The entire experiment will be self contained with respect to power supply, data storage, and on-orbit cycling. An experiment power and data system (EPDS) will be used for the data recording and LiSO₂ batteries will provide EPDS power.

Heat will be supplied to the modules through foil heaters bonded directly to the interior surface of the modules. The batteries and other components are not utilized to supply heat to the modules because it was desired to be able to accurately control the module environments and vary the heat loads to allow a more detailed verification of the experiment capability. Heater power will be provided by 28-V lithium monofluorographite batteries.

The experiment timeline is shown in figure 35. Three identical experiment "on" times are planned during the mission. Each "on" time will last approximately 13 hours (8.6 orbits) and will have heat input to the modules as shown in figure 36. Each "on" period will be subdivided into two 4.3-orbit heater input periods to verify proper operation of each module. The initial ' on" will occur approximately I month after launch, the second "on" 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.

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[78] LDEF Thermal Measurements System (P0003)

Robert F. Greene, Jr.

NASA Langley Research Center

Hampton, Virginia

Background

Many of the passive experiments flying on LDEF will be significantly enhanced if data are available postflight to indicate the temperature-time histories of test materials and other specimens exposed in the experiments. The baseline LDEF approach was to provide postflight calculated temperature histories of experiment boundaries and solar flux data for the mission, which can in turn be used by each investigator to calculate the temperature-time histories for critical experiment components. Without in-flight temperature measurements, a substantial uncertainty (±40°F) will exist in the calculated temperatures. The data measured by the thermal measurement system (THERM) will significantly improve postflight knowledge of temperatures experienced by LDEF experiments. The THERM data will also be valuable in validating the LDEF thermal design concept and in providing better design data for experimenters on future LDEF missions.

Objectives

The objectives of this experiment are to determine the history of the interior average temperatures of the LDEF for the total orbital mission and to measure the temperatures of selected components and thermal boundary conditions.

Approach

The THERM system consists of six copper-constantan thermocouples (T/C's), two thermistor reference measurements, an electronic system, one 7.5-V battery, and an interface harness with the HEPP experiment. Data are recorded on dedicated channels of a shared EPDS tape recorder in the low-temperature heat pipe experiment package (HEPP) (S1001).

The THERM hardware locations are shown on LDEF in figure 37. Measurement I provides the temperature of the center structure and a backup measurement of the average temperature. Measurement 3 is on the top of the dome surrounding the viscous magnetic damper. Measurement 4 is on a radiometer suspended in the middle of the center ring and is designed to track the average interior temperature of LDEF. Measurement 5 is on a side longeron structure that is expected to see the maximum structural tempera-....

...-ture. This measurement can also be used for rough attitude determination of the LDEF. Measurement 6 is on the space end of the structure and provides a representative boundary condition for the experiments mounted on the space end. Measurement 7 gives similar data for the Earth-facing end. Measurements 2 and 8 are thermistors that measure the reference junction temperature in the THERM electronics.

Operationally, THERM will be activated when it receives its own initiate or " set" signal from LDEF just prior to LDEF deployment into orbit. Routine scans of data will be taken about 12 times daily; however, on occasions during the mission the HEPP logic will trigger the EPDS to record data in the high-frequency data recording mode for periods up to 15 days. During the high-frequency mode, data scans will be taken every 5 minutes to [80] provide temperature profiles throughout typical orbits. The THERM data will therefore provide both long-term and transient temperatures. Total system accuracy is within plus or minus 10°F for all measurements over a range from -30°F to 170°F.

The THERM data, other experiment temperature data, and LDEF attitude information will be reduced and analyzed postflight to provide each experimenter with an improved time history of the experiment boundary conditions encountered during the LDEF mission.

 

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