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Description  |
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This invention relates to a novel solar panel which is especially useful in
space.
BACKGROUND OF THE INVENTION
Solar panels are conventionally used as a source of electrical power for
spacecraft, such as satellites. The solar panels typically used for
spacecraft include a substrate and a plurality of individual photovoltaic
solar cells which are secured to a face surface of the substrate. The
individual solar cells are electrically connected together to form a
series-parallel solar cell array which, when oriented properly toward the
sun, converts solar energy into electrical energy.
The most important consideration for solar panels used on spacecraft is
reliability. If a solar panel fails in space, it is difficult, if not
impossible, to correct or compensate for the resulting loss of electrical
power with the result that the useful life of the entire spacecraft is
often prematurely ended.
Reliability of solar panels in space applications is difficult to obtain
because the solar panels are subjected to a wide variety of conditions in
use, many of which are extremely harsh and tend to damage or destroy the
solar panels. For example, during the launch of a spacecraft such as a
satellite into outer space, the solar panels, which are typically stored
for launch in a compact, folded configuration, are subjected to the
extreme vibrations and high gravitational forces encountered during blast
off. After the satellite has been separated from the launch vehicle and
placed into orbit, the solar panels are deployed from their compact,
folded configuration and extended to an open configuration with the array
of solar cells oriented toward the sun. In the deployed configuration, the
solar panels are subjected to substantial thermal stresses; the solar
cells and the face surfaces of the substrates are subjected to the intense
heat of the sun while the back surfaces of the substrates are subjected to
the extreme cold of outer space. Furthermore, in order to have the solar
array operate at maximum efficiency it is necessary that the substrate be
relatively rigid so that it can maintain all the solar cells in the
correct alignment with respect to the sun and be sufficiently strong so as
not to break under the forces inherently applied to the deployed solar
panels in space. Because of the above-noted requirements it has been
extremely difficult to form the solar panels with the required
reliability.
The most important consideration after reliability in the construction of
solar panels for spacecraft, such as satellites, is weight. The solar
panels should be as light in weight as possible to allow increased
payloads and increased amounts of fuel to be stored on board the
satellite. In this regard, it should be noted that the relative amount of
fuel stored on a satellite is particularly important as this fuel is used
for the rocket thrusters which are periodically activated to reorient the
satellite to the proper attitude. Any additional fuel that can be stored
on board by using lighter weight solar panels can be used to extend the
useful life of the satellite.
It has not heretofore been possible to construct a solar panel having the
desired combination of high reliability and light weight. Suggestions made
heretofore to improve the reliability of the solar panels typically
increased the weight of the solar panels, and suggestions made to reduce
the weight of the solar panels caused a reduction in the overall
properties and particularly the flexural strength of the solar panels and
thus reduced reliability. For example, it was suggested to use thicker,
stiffer substrates and/or to add reinforcement ribs to stiffen the
substrate of the solar panels. These proposals and other similar proposals
significantly increased the weight of the solar panels. It was also
suggested to use thin plastic films as the supporting substrate to lighten
the solar panels, but this caused the substrate to be excessively flexible
and unstable which decreased its reliability.
State of the art substrates are comprised of a honeycomb core having an
outer skin. These substrates are made with various combinations of
materials, including aluminum honeycomb cores with aluminum skins, fiber
reinforced plastic honeycomb cores with fiber reinforced plastic skins,
and aluminum honeycomb cores with fiber reinforced plastic skins. These
substrates require, in addition to the honeycomb core and associated skin,
reinforcement ribs to provide adequate strength. These ribs, however,
increased the total weight. Furthermore, the relatively complex honeycomb
substrates are expensive to manufacture and prone to mechanical failure,
especially at the junctions of the skin with the honeycomb core.
What would be highly desirable would be a solar panel having a substrate
which is relatively strong, simple in construction, has a high degree of
reliability, and which is relatively light weight as compared to the solar
panel heretofore employed.
SUMMARY OF THE INVENTION
A novel solar panel is disclosed which is especially useful for spacecraft
applications. The solar panel has a supporting substrate which is
comprised of one or more plies of a resin reinforced novoloid fabric. The
novel solar panel is lighter in weight and has at least equivalent overall
physical properties as compared to conventional solar panels.
BRIEF DESCRIPTION OF THE DRAWING
FIG. 1 is a pictorial illustration of a satellite showing the solar panels
thereof in the fully extended position.
FIG. 2 is an isometric projection in partial cross section of a portion of
a solar panel of this invention.
FIG. 3 is a cross-sectional illustration of the preferred embodiment of the
solar panel of the present invention.
DETAILED DESCRIPTION OF THE INVENTION
In FIG. 1 there is illustrated a communications satellite 10 which includes
main reflectors 12, 14, and subreflectors 16, 18 attached to the body 20
of the satellite 10. The body 20 of the satellite 10 also contains the
fuel supply and the rocket thrusters (not shown) which are used to
reorient the satellite 10. Extending outward from each side of the body 20
of the satellite 10 are booms 22, 24 on which are mounted a number of
individual solar panel units 26, 28, 30, 32, 34, 36. The solar panel units
in use are oriented towards the sun and convert solar energy into
electrical energy.
The structure of the novel solar panel 38 of this invention is best
illustrated in FIGS. 2 and 3. The solar panel 38 is comprised of a
substrate 40 having a face surface 42 on which an array of solar cells 44
are attached. In addition to the substrate 40 and the array of solar cells
44, the solar panel 38 can also include other conventional elements which
are not illustrated, such as protective covers for the solar cell array
44, electrical networks for interconnecting the individual solar cells,
brackets and support structures for securing the solar panels to the booms
22, 24 and the like.
The substrate 40 of the solar panel 38 of this invention is comprised of at
least one ply of a resin reinforced novoloid fabric. The novoloid fabrics
are manufactured from novoloid fibers. Novoloid fibers chemically are
cross-linked phenolic-aldehyde fibers which are prepared by acid catalyzed
cross linking of a melt spun novolac resin with formaldehyde. The generic
term novoloid is recognized by the Federal Trade Commission as designating
a manufactured fiber containing at least 85% by weight of a cross-linked
novolac. Novoloid fibers are commercially available under the trademark
Kynol from Nippon Kynol, Inc. (Japan) and American Kynol, Inc. They are
available in a range of diameters from about 14 to 33 .mu.m. For purposes
of this invention, it has been found that the smaller diameter novoloid
fibers are preferable because stronger bonding is possible with the resin
used to form the matrix of the substrate 40.
Novoloid fibers have a unique combination of physical properties which make
the novoloid fibers especially useful for solar panels which will be used
in outer space. The novoloid fibers are stable when exposed both to
relatively high temperatures for prolonged periods of time especially in
the absence of oxygen as in outer space, and at very low temperatures.
They are also stable to ultraviolet radiation.
In the manufacture of the solar panel 38 of this invention, the novoloid
fibers are initially formed into either woven or knitted fabrics using
conventional textile manufacturing methods. The use of fabric as opposed
to fiber facilitates the manufacture of the substrate 40 and also
substantially increases the overall strength of the substrate 40. The
fabric can be manufactured in various weaves and different knits. It is
preferable, however, to form the fabric with a relatively lightweight,
open structures, so as to insure proper wetting of the novoloid fibers
with the resin employed in the preparation of the substrate 40. It has
been found that the optimum results are obtained with either plain weaves
or twills having a weight of about 9 to about 10 ounces per square yard
(278-286 grams/sq. meter).
Various resins can be used in the manufacture of the substrate 40,
providing they have good strength and good stability to environmental
conditions encountered in space. Epoxy-based resins are known to be
suitable for this application and can be used in the present invention.
Certain polyester resins, and in particular the isophthalic polyester
resins, are particularly suitable for the manufacture of substrates 40 of
this invention. Suitable polyester resins are commercially available from
various sources, such as MR 14042 sold by U.S.S. Chemicals.
The substrate 40 is manufactured using conventional molding techniques. The
number of plies of novoloid fabric required to form a specified substrate
are initially cut to size. The number of plies can be as few as one
relatively heavy ply, but preferably a number of lighter weight plies 48,
50, 52, 54, 56, 58 as shown in FIG. 3 are employed rather than one heavier
ply. In practice, it has been found that it is preferable to use about six
plies of fabric since this does not unduly complicate the manufacturing
process and results in the final substrate 40 having an excellent
combination of physical properties.
The plies of the novoloid fabric are impregnated with the resin making sure
that the individual novoloid fibers are well wet during the process. The
relative amount of the novoloid fabric and the resin employed in the
manufacture of the substrate 40 can be varied somewhat, but the substrate
40 should contain about 40 to 60% by weight of the novoloid fabric, with
the remainder being the resin. The impregnated plies of the novoloid
fabric are then laid up on each other and the assembly is compression
molded to form a laminate which, after trimming, is used as the substrate
40 of the solar panel 38 of this invention. The resin employed in the
manufacture of the substrate 40 should form an interlocking matrix between
and through the plies 48, 50, 52, 54, 56, 58 of the novoloid fabric of the
substrate 40.
Once the substrate 40 is obtained, the solar cell array 44 and any other
required elements are attached to the substrate 40 in conventional manner.
The substrate 40 which is obtained in accordance with this invention is at
least about 10 to 11% lighter in weight than the equivalent strength
substrate made from conventional materials such as fiberglass or aramid
fibers. This substantial reduction in weight allows the satellite 10 to
carry additional fuel which as noted above can substantially extend the
life of the satellite 10 by permitting additional reorientation of the
satellite 10 as required in use.
In order to obtain a more direct comparison of the advantages of the
substrate 40 of the present invention as compared to the prior art
substrate, a series of substrates were prepared using various types of
fabrics and resin systems. The amounts and weights of fabrics used were
selected so as to be most suitable for the particular system evaluated.
The below-tabulated results were obtained.
TABLE
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FABRIC REIN-
FORCEMENT ARAMID FIBERGLASS NOVOLOID
RESIN MATRIX
EPOXY EPOXY POLYESTER
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Fabric Content,
55 47 40
Weight %
Resin Formula,
45 53 60
Weight %
Specific Gravity
1.32 1.86 1.23
Tensile Strength,
28,500 14,400 32,800
psi
Tensile Modulus,
2.96 2.23 2.98
10.sup.6 psi
Flexural Strength,
35,500 29,100 39,500
psi
Flexural Modulus,
2.67 2.33 2.97
10.sup.6 psi
IZOD Impact
17.4 19 18.4
Strength, ft/lb in.
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