A solar panel satellite subsystem includes a first and second set of solar panels and booms joined together end to end in series and to opposite sides of the spacecraft by spring bias hinges. These hinges include viscous fluid dampers. Each damper has its own heater and thermostat coupled thereto to maintain essentially the same damper temperature as the ground test temperature at each hinge and maintain essentially the same desired rate of deployment at each hinge.

This invention relates to a deployable folded satellite subsystem such as 
the solar panel subsystem of a spacecraft, which is deployed from a folded 
to an extended position and more particularly to a freely deployed 
subsystem which does not include tether lines to control deployment. 
Satellites often have large deployable panels and/or booms which are hinged 
together or hinged to the main spacecraft body. The deployment is 
triggered by a control signal to the spacecraft which causes the release 
of some retainer mechanism. This retainer mechanism may be released by, 
for example, an electronically triggered pyrotechnic device. Then by a 
spring mechanism in the hinge, for example, the panels are deployed. 
It sometimes occurs that the panels are sprung open too fast or some panels 
are sprung at a different rate of speed than other panels. This can cause 
damage to the lightweight panels or booms supporting them and can cause 
high disturbance forces to be imparted (nutation for example) to the 
spacecraft. A damping mechanism, such as a viscous damper may be employed 
to slow down or retard the deployment of the booms and/or panels. Viscous 
damping is highly dependent on the ambient temperature which affects the 
viscosity of the fluid in the damper. Satellites are often subjected to 
large differences in temperature, from relatively high values when under 
direct sunlight to low values when shadowed from the sun, so that a 
viscous damper when used in a satellite often experiences large 
differences in its damping rate. For a large deployable multielement 
structure these temperatures may vary dramatically from element to element 
over the length of the structure. 
The proper deployment of solar panels for example, is an important matter. 
Failure of such panels, and/or booms, to operate and deploy correctly can 
cause failure of the complete mission which can cost many millions of 
dollars. It has been reported that at least one satellite designed for a 
seven year life ran out of station keeping fuel in about six months time 
after much of it had been exhausted shaking loose a sticky hinge antenna 
and a solar sail. 
This deployment problem is further increased when one is deploying a 
plurality of solar panels which extend from opposite sides of the 
spacecraft with the joints between the panels being exposed to differing 
amounts of the sun depending on the position of the sun with respect to 
the spacecraft. The temperature variations can differ from about 
-50.degree. C. to +50.degree. C. 
In accordance with an embodiment of the present invention an improved 
deployable folded satellite subsystem, deployable from a main spacecraft 
body includes a set of structural elements with the structural elements 
joined end to end using a spring biased hinge between the structural 
elements and one end of the set joined by a spring bias hinge to one 
surface of the main body, to form a deployable folded structure so that 
the set of structure elements are deployed from a folded position against 
and parallel to the one surface of the main body to an extended position 
generally perpendicular to the one surface due to forces provided by the 
spring bias. Each spring bias hinge further includes a viscous damper for 
damping the rate of deployment of the elements. Each of the hinges 
includes its own heater and thermostat at the damper for maintaining the 
temperature of the fluid in the damper of each hinge at substantially the 
same predetermined level and thereby maintain the deployment rate of all 
of said structure elements equal.

Referring to FIG. 1 there is illustrated a satellite 10 comprising a main 
satellite body 11 with antenna 10a and a first set 12 of series connected 
structural elements 13 through 17. These structural elements are, for 
example, elements of a solar array comprising series connected lightweight 
booms 13 and 14 connected in series with series connected lightweight 
solar panels 15, 16 and 17. 
These series connected elements 13 through 17 are connected together using 
spring bias hinges 18. The boom 13 is connected to the one surface 11a of 
the main spacecraft body via a 90.degree. spring bias hinge 18a. Boom 14 
is connected to boom 13 via a 180.degree. spring bias hinge 18b. The boom 
14 is connected in series to solar panel 15 via 180.degree. spring bias 
hinge 18c. The solar panel 16 is connected to solar panel 15 via 
180.degree. spring bias hinges 18d and 19 and the solar panel 17 is 
connected to solar panel 16 via 180.degree. spring bias hinges 18e and 19. 
The 90.degree. hinge 18a has an end stop that allow the boom 13 to rotate 
90.degree. till it is perpendicular to surface 11a. The 180.degree. hinges 
have end stops at 180.degree. so the panels and the booms unfold to be 
coplanar with each other. 
Similarly, a second set 20 of lightweight satellite elements 23 through 27 
is connected to the opposite surface 11b of the satellite main body 11 to 
extend colinearly with said first set 12. The second set 20 of elements is 
identical to the first set and includes lightweight booms 23 and 24 
connected in series by a 180.degree. spring bias hinge 18f where boom 23 
is connected to the main body via a 90.degree. spring bias hinge 18g. The 
solar panels 25, 26 and 27 are connected together and to boom 24 via 
180.degree. spring bias hinges 18 and 19 as shown. 
The solar panels and booms are connected together and hinged such that when 
the panels and boom are not deployed the booms and the panels are folded 
over each other as illustrated in FIG. 2. When the spacecraft is in the 
required orbit position a retainer mechanism, not shown, is released by 
ground command and the spring bias on the hinges causes the panels and 
booms to be deployed as illustrated in FIG. 3. These panels continue to 
unfold to the position illustrated in FIG. 1. 
Since these panels and booms are generally made of very lightweight 
structures, the deployment rates of panels must be controlled to prevent 
damage to the panels and booms. A sufficient spring bias however, must be 
maintained in order to assure that the panels and booms will unfold. Each 
hinge contains a spring (for example a leaf spring) to provide the force 
required to deploy the system. Each hinge also contains a viscous fluid 
damper to control the rate of deployment. 
When the panels and booms are deployed a force is applied on the spacecraft 
that depends on the rate and profile of deployment. The profile of 
deployment is the side view shape of the deployment as illustrated in FIG. 
3. If the rate and profile of deployment of the panels and booms is not 
controlled, then not only can the panels and boom be damaged but the 
forces applied to the spacecraft could make it difficult to control the 
spacecraft. One of the primary causes of deployment control problems is 
the variety of temperatures seen by the dampers. 
In accordance with the present invention the deployment system is made very 
reliable and predictable in rate and profile of deployment by maintaining 
the temperature of all of the dampers at substantially the same level (and 
at the test temperature) so the damping rate is the same as during ground 
test and therefore predictable. Further, by having the two similar 
subsystems with similar solar panels and booms, as shown in FIG. 1, 
deployed from opposite sides of the spacecraft and by having these hinges 
temperature controlled to substantially the same temperature, the 
resultant forces on the spacecraft tend to cancel. The two systems or sets 
of panels and booms (set 12 and set 20), deploy at substantially the same 
rate and profile with minimal disturbance on the spacecraft. After 
deployment, the power to the heater is turned off by ground command. 
Referring to FIG. 4, there is illustrated a sketch of a spring bias hinge 
18 according to the present invention. The hinge 18 includes a shaft 30 to 
which a panel, for example panel 15, is fixed via a U shaped clevis 
(flange) 31. The adjacent element, such as boom 14 for example, is 
rotatable about the shaft 30 via a flange 32 coupled to the boom 14. 
Flange 32 has apertures 34 therein through which the shaft 30 freely 
passes. The apertures 34 may include ball bearings to aid rotation of the 
shaft 30. 
The hinge 18 may be like that shown and described in U.S. Pat. No. 
4,290,168 of Derick S. Binge, which patent is incorporated herein by 
reference. A leaf spring 40 which may be made of spring sheet metal is 
coupled between the clevis 31 and flange 32 to provide biasing means for 
unfolding boom 14 about shaft 30. 
FIG. 5 represents a sketch of the spring 40 as viewed in the direction of 
arrow 41 in FIG. 4. The spring 40 is wrapped around and fixed at one end 
to cylinder 40a. Cylinder 40a is engaged with shaft 40b which is attached 
to clevis (flange) 31. The other end of spring 40 is wrapped around and 
fixed to cylinder 40c on flange 32. As the spring exerts pressure, the 
shaft 30 rotates and this rotation is damped by a damper 43 (see FIG. 4) 
coupled to the shaft 30. 
Referring to FIG. 6 there is illustrated, in a cross sectional view, the 
damper 43 as viewed along arrow 53. Damper 43 comprises a hollow cylinder 
defined by wall 45 with a central shaft 47 coupled to the shaft 30 and 
rotatable therewith. The damper 43 has a solid barrier portion 46 
extending inward from wall 46. A fin 49 is coupled to the shaft 30 via 
shaft 47 and is rotatable therewith. Chambers 50 and 51 are formed by fin 
49 and barrier position 46. Rotation of fin 49 increases the volume of one 
of chambers 50 and 51 while decreasing the other. Chambers 50 and 51 are 
filled with a viscous fluid. This fluid is for example a silicone fluid 
Type 200 of Dow Corning of Midland, Mich. The viscosity of the fluid is 
50,000 centistokes at 25.degree. C. The fin 49 includes an aperture 49a 
between the two chambers for allowing the passing of restricted amounts of 
fluid. This aperture size is adjusted to control the damping rate. It is 
0.178 centimeter in diameter for present example. As the shaft 30 rotates 
so does shaft 47 and fin 49 which compresses the fluid in one chamber and 
the fluid flows through the aperture 49a into the other chamber to damp 
the spring bias provided by the spring 40. 
In accordance with the present invention each of the dampers and each of 
the hinges includes a pair of heating elements 61 and 62 (See FIGS. 4 and 
6), wrapped around the periphery of the housing 45 on the damper for 
heating the fluid in the damper. Each damper further includes a thermostat 
63 coupled to the outer surface of the housing with the thermostat 63 
coupled via electrical wiring 65 between a DC power source and the heating 
elements 61 and 62. These heating elements may be for example, thermofoil 
type heaters supplied by MINCO products of Minneapolis, Minn. The elements 
may be for example, 1.07 centimeters wide and about 36.6 centimeters long. 
Two heating elements are used for redundancy. The heating elements 61 and 
62 are each 709 ohm resistive elements and they are coupled in parallel 
across the supply as shown in FIG. 7. As illustrated in FIG. 4, the damper 
43 is thermally isolated from the rest of hinge 18 by insulator 33 so that 
the heaters only have to heat the damper. 
Each of the hinges 18a, 18b, 18c, 18d, 18e etc. are similar to the one 
shown in FIG. 4 wherein they each contain their own thermostats and their 
own heating elements and are coupled in parallel across the power source 
as shown in FIG. 7. The thermostat for each pair of heating elements is 
set at a particular predetermined temperature such as about 22.degree. C. 
This is approximately the temperature of the room where the hinges were 
originally tested. Also this temperature is at about the maximum 
temperature any part of this example satellite will experience for the 
particular orbit and orientation of this example satellite. If the 
satellite was in an orientation and orbit where the hinges would see 
+50.degree. C., the predetermined temperature would be 50.degree. C. and 
the room testing temperature would be +50.degree. C. In this manner each 
of the hinges is heated to substantially the same temperature and 
therefore the viscosity of the fluid in each of the dampers should be 
substantially the same. The deployment rate for each of the panels and 
booms is therefore substantially the same. The fluid selected and the 
temperature selected is such that the rate of deployment is, for example, 
approximately 30 seconds. This rate of deployment has been found to be 
ideal for the type of lightweight structures involved. Shorter times on 
the order of 10 seconds have been found to cause damage to the lightweight 
structures and therefore highly undesirable.