SP-473 LDEF

 

Power and Propulsion

 


[82] Space Plasma High-Voltage Drainage Experiment (A0054)

William W. L. Taylor and Gene K. Komatsu
TRW Space and Technology Group
Redondo Beach, California

 

Background

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 solar-array 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.

 

Objectives

The objectives of this experiment are to place large numbers of dielectric samples under electric stress in space; to determine their in-space current drainage behavior; 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. 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.

 

Approach

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 [83] frame of the LDEF and are then delivered to the ground return of the bias voltage supply. Figure 38 illustrates the experimental arrangement. 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 are not electrically stressed in flight and hence allow a determination of the effects of merely being present on the LDEF. Figure 39(a) illustrates the dielectric sample construction. The actual experiment will occupy two 3-in.-deep peripheral trays. One tray will be located near the LDEF leading edge and the other will be near the trailing edge. This configuration will allow the determination of charged-particle drainage as a function of plasma density. Figure 39(b) shows a top view of one tray, minus the test samples.

The dielectric sample power processor is equipped with two coulometers. The first of these is in series with the bias voltage lead and determines the integral of the drainage current during the flight. The second coulometer, which has a high-value in-series resistor, is placed between the....

 


Figure 38.- High-voltage drainage experiment concept.

Figure 38.- High-voltage drainage experiment concept.


[
84]

Figure 39.- Space plasma high-voltage drainage experiment.

Figure 39.- Space plasma high-voltage drainage experiment.

 

[85] ...bias potential and LDEF ground to determine the time-integrated applied-bias voltage. An average front-to-back resistance of the dielectric sample is 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 is 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 rule out 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.

 


[86] Solar-Array-Materials Passive LDEF Experiment (A0171)

Ann F. Whitaker, Charles F. Smith, Ir., and Leighton E. Young
NASA George C. Marshall Space Flight Center
Huntsville, Alabama
 
Henry W. Brandhorst, Jr., and A. F. Forestieri
NASA Lewis Research Center
Cleveland, Ohio
 
Edward M. Gaddy and James A. Bass
NASA Goddard Space Flight Center
Greenbelt, Maryland
 
Paul M. Stella
Jet Propulsion Laboratory
Pasadena, California

 

Background

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 will lead to reduced solar-cell electrical output, a situation that is especially critical at high astronomical units (AU's). 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 AU's. Protons, electrons, atomic oxygen, and UV irradiation contribute to surface damage in these array materials. Thin-film materials can become 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.

 

[87] Objective

The objective of this experiment is to evaluate the synergistic effects of the space environment on various solar-array materials, including solar cells, cover slips with various antireflectance (AR) coatings, adhesives, encapsulants, reflector materials, substrate strength materials, mast and harness materials, structural composites, and thermal control treatments.

 

Approach

The experiment is passive and consists of an arrangement of material specimens mounted in a 3-in.-deep peripheral tray. A photograph of the tray, which has been subdivided among the various experiment organizations, is shown in figure 40. The effects of the space environment on the specimens will be determined by comparison of preflight and postflight measurements of mechanical, electrical, and optical properties.

 


Figure 40.-Solar-array materials experiment.

Figure 40.-Solar-array materials experiment.

 


[88] Advanced Photovoltaic Experiment (S0014)

Henry W. Brandhorst, Jr., and A. F. Forestieri
NASA Lewis Research Center
Cleveland, Ohio

 

Background

The advanced photovoltaic experiment consists of a group of three photovoltaics-related experiments for investigating a portion of the solar spectrum and the effect of the space environment on photovoltaics. The information will be used to provide correlation between space and ground testing and also to provide for more accurate performance measurement in the laboratory.

 

Objectives

Specific objectives of these experiments are to provide information on the performance and endurance of advanced and conventional solar cells, to improve reference standards for photovoltaic measurements, and to measure the energy distribution in the extraterrestrial solar spectrum.

 

Approach

The experiment will occupy a 12-in. -deep peripheral tray and will use an experiment power and data system (EPDS) for data recording and LiSO2 batteries to satisfy power requirements. Figure 41 shows a photograph of the experiment.

The experimental approach for the three experiments is detailed below.

Space Exposure of Solar Cells

Space exposure of advanced and conventional cells will provide information on the performance and endurance of such cells in the space environment. Correlation between space environment and ground simulation test results will also be verified by this experiment.

Data to be obtained will include temperatures and short-circuit current of the samples. Six-point current-voltage (I-V) characteristics will be obtained for selected samples. These data will be recorded once a day during the flight. Orbit data will be correlated with preflight and postflight measurements of the samples.

 


[
89]

Figure 41.-Advanced photovoltaic experiment.

Figure 41.-Advanced photovoltaic experiment.

 

Reference Solar-Cell Calibration

Various reference cells, including some previously measured on balloon, aircraft, or rocket flights, will be measured before flight and throughout the flight to determine their outputs (short-circuit current). Upon return, those cells with known output in space will serve as laboratory standards for accurate determination of space output from other cells and arrays. The flight of previously calibrated cells will permit verification of the accuracy of the various calibration techniques.

 

Solar Spectrum Energy Distribution

A series of optical bandpass filters coupled to solar-cell detectors will be used to determine the energy in 16 spectral regions between 0.3 and 1.1 µm In addition, the total energy in the spectrum above and below 0.5 µm will be measured using a dichroic 45° mirror. The characteristics of the filters will be measured both preflight and postflight. The energy within the appropriate bandpass will be determined from the short-circuit current of the detector solar cell measured in space. These measurements will be used to assess the accuracy of laboratory instruments such as the filter wheel solar simulator.

[90] Finally, the total energy in the solar spectrum will be determined with an absolute radiometer detector.

The following participants have supplied samples for these experiments: Air Force Wright Aeronautical Laboratory; Applied Solar Energy Corporation; Comsat Laboratory; European Space Agency; Hughes Research Laboratory; Jet Propulsion Laboratory; Lockheed Missiles and Space Company, Inc.; NASA Langley Research Center; NASA Lewis Research Center; NASA Marshall Space Flight Center; Rockwell international Corporation; Solarex Corporation; Spectrolab inc.; Spire Corporation; and Varian Associates.

Experiment operation will be automatically timed by the EPDS clock, which begins with an initiate command at LDEF deployment. Data will be recorded once each day when a maximum Sun angle less than 20° is reached. (This will be determined by a two-axis Sun angle sensor, which detects the maximum cosine angle for the data period.) A scan of data will consist of timing, Sun sensor output, temperatures, six-point current-voltage data of 16 solar cells, and short-circuit currents of 120 cells. Approximately I month prior to planned retrieval of the LDEF, the experiment will be terminated.

 


[91] Investigation of Critical Surface Degradation Effects on Coatings and Solar Cells Developed in Germany (S1002)

 

Ludwig Preuss
Space Division, Messerschmitt-Bolkow-Blohm
Munich, Federal Republic of Germany

 

Background

Various coatings developed in the FRG (i.e., second-surface mirrors with interference filters with and without conductive layers, conductive layers on solar-cell covers, and selective absorber coatings) have been qualified by accelerated tests under simulated space environment conditions. Experiments with coatings and solar cells have shown, however, that the thermo-optical behavior can differ considerably when performed on the ground and in space because of the great difficulty in simulating the space environment realistically.

 

Objectives

The objective of this experiment is to qualify these coatings under realistic space environment conditions. In addition, the experiment will provide design criteria, techniques, and test methods to insure control of the combined space and spacecraft environment effects, such as contamination, electrical conductance, and optical degradation, on the coatings.

 

Approach

Figures 42 and 43 show the experiment arrangement and electronics block diagram, and table 10 lists the samples to be investigated. Test samples will be installed in an experiment exposure control canister (EECC) and on a cover sheet near the upper surface of the 6-in.-deep peripheral tray. The samples in the canister will be exposed only to the space and spacecraft environment because the canister will be opened after LDEF deployment and closed prior to LDEF retrieval. The other samples will be exposed to the complete mission environment. Data to be measured include the temperature of the samples, the electrical resistance of the conductive layers of the samples, the short circuit current of the solar-cell modules, and the deposition of contaminants on the samples (using quartz crystal microbalances (QCM's)).

The data will be measured according to a defined time program and will be amplified, digitized, and stored on a data recorder. After the return of the...

 


[
92]

Figure 42.- Experiment arrangement.

Figure 42.- Experiment arrangement.

 


Figure 43.- Electronics block diagram.

Figure 43.- Electronics block diagram.

 

...experiment, the stored data will be evaluated along with the data for attitude and related solar aspect angles to determine relations between space conditions and surface effects on the test samples. In addition, contamination will be investigated by means of infrared spectroscopy, and the electrical characteristics of the solar cells will be determined.

 

[93] Table 10.-Test Samples To Be investigated.

Components to be investigated
 
Second-surface mirror with
Ag reflector and Inconel protection layer on rear
Interference filter on front face
Second-surface mirror with
Ag reflector and Inconel protection layer on rear
Interference filter on front face
Doped In2O3 layer on interference filter
Chromium black selective absorbers
Solar-cell modules with doped In2O3 layer on cover glass
 
Reference components
Second-surface mirror with Ag reflector and Inconel protection layer on rear
Optical solar reflector with Ag reflector and Inconel protection layer on rear
Solar-cell module

 


[94] Space Aging of Solid Rocket Materials (P0005)

Leon L. Jones and R. B. Smalley, Jr.
Morton-Thiokol, inc.
Brigham City, Utah

 

Background

Solid-rocket motors continue to be used extensively in space applications, and future missions have been identified in which the solid motor may be stored in space for an extended time before firing. This indicates a need to gather direct information on the effects of extended storage of rocket materials in the combined vacuum and thermal conditions of space. Tests have been performed in high vacuum to determine outgassing characteristics, but only limited testing has been done on the effects of vacuum on the mechanical and ballistic properties of the materials themselves. Most vacuum aging has been plagued by mechanical problems and subsequent back contamination of the material samples.

 

Objective

The objective of this experiment is to determine the effects of long-term orbital exposure on the materials used in solid-rocket space motors. Specifically, structural materials and propellants from the STAR PAM-D series motors and the PAM DII/IPSM-II motors will be tested, as well as advanced composite case and nozzle materials planned for future use.

 

Approach

The experiment approach is to expose samples of solid-rocket propellant, liner, insulation, case, and nozzle specimens to the space environment and to compare preflight and postflight measurements of various mechanical, chemical, and ballistic properties to determine the effects of long-term orbital exposure. A parallel program will be conducted on ground storage samples; thus the data will be applicable on a generic basis as well as to the specific materials being tested.

Figure 44 shows a view of the IPSM-II space motor with some of the sample materials identified. Table 11 lists the STAR materials and the tests to be performed. These materials will be packaged within a 5- by 6.5- by 11.5-in. aluminum container and attached to an interior plate on the LDEF center ring. The container will be flushed with dry nitrogen and sealed before installation on the LDEF. An air-pressure-activated valve has been designed....

 


[
95]

Figure 44.-Sketch of IPSM-II motor showing examples of aging samples.

Figure 44.-Sketch of IPSM-II motor showing examples of aging samples.

 

.....to vent the box when it is subjected to pressure below 0.2 atm, leaving the box open to the vacuum. On reentering, the valve will close at 0.5 atm external pressure.

 

[96] Table 11.- Solid Rocket Materials a

Mechanical properties: tensile, relaxation, dynamic

Propellant materials:
IPSM-II/PAM-DII
STAR/PAM-D
 
Insulation:
EPDM rubber
NBR rubber
 
Nozzle materials:
Carbon-carbon
Silica-phenolic
Carbon-phenolic
 
Case materials:
Kevlar-epoxy
Graphite-epoxy
Glass-epoxy

Interface strength: tensile, peel, constant load

Titanium-insulation
Kevlar-insulation
Graphite-insulation
Kevlar-NBR-glass
Glass-insulation
Three propellants-insulation
Igniter housing-propellant

Ballistic properties: burning rate

IPSM-II/PAM-DII propellant and igniter assembly
STAR PAM-D propellant and igniter
Boron-potassium nitrate pellets

 

a AII of the samples will be subjected to determination of weight loss with attendant chemical analysis.

 
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