Modular communication satellite

A Modular Communication Satellite (10) for a Satellite Communication System is disclosed. The preferred embodiment of the invention includes a foldable, high-gain, electronically steered antenna array (12) that is always pointed toward the Earth (E). The unfolded spacecraft resemble an oblate flower. Polygonal antenna panels (92, 94, 96, 102, 104, & 106) are attached to each other and to a primary bus structure (22) by antenna deployment hinges (90). The upper portion of the satellite (10) incorporates intersatellite antenna arrays (26) of individual intersatellite antennas (28), which are always pointed tangentially to the Earth (E). An Astromast.TM. boom (32) is mounted between the space facing surface of the primary bus structure (22) and an assembly of solar array storage booms (36). The Astromast.TM. boom (32) can expand and rotate amorphous silicon solar arrays (38) which are unfurled from within the solar array storage booms (36). The amorphous silicon solar arrays (38) gather solar radiation to provide power the satellite (10), and furnish thermal control for the satellite (10) by shielding it from solar radiation. The satellite is capable of being nested or stacked in a compact arrangement that fits within a payload bay of a launch vehicle (LV).

CROSS-REFERENCES TO RELATED PATENT APPLICATIONS 
The present patent application is related to the following commonly-owned 
and commonly-assigned patent applications: 
Satellite Communication System by Edward F. Tuck et al., filed on Oct. 28, 
1991 and assigned U.S. Ser. No. 07/783,754; 
Terrestrial Antennas for Satellite Communication System by Asu Ram Jha, a 
Continuation-in-Part application filed on Dec. 2, 1992 and assigned U.S. 
Ser. No. 07/984,609 and claiming the benefit of priority of a parent 
application entitled Terrestrial Antennas for Satellite Communication 
System by Asu Ram Jha, filed Nov. 8, 1991 and assigned U.S. Ser. No. 
07/790,273, now abandoned; 
Switching Methods for Satellite Communication System by David Palmer 
Patterson & Moshe Lerner Liron, filed on Nov. 8, 1991 and assigned U.S. 
Ser. No. 07/790,805, now abandoned; 
Beam Compensation Methods for Satellite Communication System by David 
Palmer Patterson and Mark Alan Sturza, filed on Nov. 8, 1992 and assigned 
U.S. Ser. No. 07/790,318; 
Spacecraft Antennas & Beam Steering Methods for Satellite Communication 
System by Douglas Gene Lockie and Mark Thomson, filed on Oct. 28, 1992 and 
assigned U.S. Ser. No. 07/967,988 and claiming the benefit of priority of 
a parent application entitled Spacecraft Antennas & Beam Steering Methods 
for Satellite Communication System by Douglas Gene Lockie, filed on Nov. 
8, 1991 and assigned U.S. Ser. No. 07/790,271, now abandoned; 
Spacecraft Designs for Satellite Communication System by James R. Smart and 
David P. Patterson, filed on Aug. 18, 1992 and assigned U.S. Ser. No. 
07/931,625, now U.S. Pat. No. 5,386,953 and claiming the benefit of 
priority of a parent application entitled Spacecraft Designs for Satellite 
Communication System by James R. Smart filed on Nov. 8, 1991 and assigned 
U.S. Ser. No. 07/790,748, now abandoned; and 
Traffic Routing for Satellite Communication System by Moshe Lerner Liron, 
filed on Feb. 9, 1993 and assigned U.S. Ser. No. 08/016,204, now 
abandoned. 
The specifications of the patent applications listed above are hereby 
incorporated by reference. 
FIELD OF THE INVENTION 
The present invention relates to the field of satellite designs. More 
particularly, this invention is a part of a constellation of 840 extremely 
high power and ultra-lightweight spacecraft grouped in sets of 40 equally 
spaced satellites which circle the globe in 21 low Earth orbits. The 
satellites operate in 700 km (435 mile) circular, sun-synchronous orbits 
which are inclined approximately 98.2 degrees to the equator. 
BACKGROUND OF THE INVENTION 
Communications satellites operating in Earth orbit were first seriously 
proposed during the middle of this century. A relatively small portion of 
current telephone traffic is relayed between ground stations by spacecraft 
carrying transponders that are located over a fixed portion on the Earth 
in 22,300 mile geosynchronous orbits. Over the past few decades, public 
phone systems have relied primarily on land lines and microwave repeaters 
to handle call traffic. Cellular networks now provide service which 
extends previous network capabilities. Customers using hand-held portable 
phones or carphones are now able to access the conventional, centralized 
land-based system without using a traditional fixed phone, as long as 
their transportable terminals are within the range of land-based antenna 
towers called "cell sites." Even in the United States, these cell sites 
are not universally prevalent, since market forces restrict cellular 
service to only the most densely populated urban portions of our country. 
Since cellular service is available to only a small minority of privileged 
users in wealthy countries, and is virtually non-existent in lesser 
developed parts of the world, the operators of traditional phone networks 
are confronted with serious systemic problems that severely constrain the 
continued growth of their communications utilities. 
A spacecraft design that enables the compact nesting of multiple spacecraft 
in the same launch vehicle has partially been addressed by Mark G. 
Rochefort. In International Application Number PCT/US88/02365, Rochefort 
considered using a substantially cup-shaped configuration for stored 
spacecraft to allow a large number of identically shaped spacecraft to be 
stored together in the stowage compartment of a launch vehicle. This 
approach did not anticipate the large number of antennas necessary for a 
satellite in a large communications system. Rochefort did not consider the 
advantage of storing solar arrays near the antenna arrays. He also did not 
consider unfurling a solar array away from the antenna arrays. Rochefort's 
stowage of satellites in a cup-shaped manner also unnecessarily exposes 
the surfaces of a satellite to damage during storage, launch and placement 
into orbit. 
No system that is currently available to the general public is capable of 
taking advantage of the enormous augmentation of communications capacity 
that could be achieved if the traditional centralized grid of terrestrial 
switches, wires, fibers, and microwave repeaters could be completely 
bypassed. Public phone companies are not presently able to sell continuous 
global service to their customers who wish to use phones that are not 
hard-wired to the land-based network. Some commercial spacecraft now in 
service help to relay some portion of the total call traffic, but all 
these calls must still pass through the conventional land-based system. 
The problem of providing an economically feasible network for voice, data, 
and video which can be used by subscribers all over the world has 
presented a major challenge to the communications business. 
The development of a constellation of reliable, high gain satellites which 
can communicate directly to terrestrial terminals without routing calls 
through land-based networks would constitute a major technological advance 
and would satisfy a long felt need within the electronics and telephone 
industries. 
SUMMARY OF THE INVENTION 
A Modular Communication Satellite for a Satellite Communication System is 
disclosed. The preferred embodiment of the invention includes a foldable, 
high-gain, electronically steered antenna array which is always pointed 
toward the Earth, and which resembles an oblate flower. A group of 
polygonal antenna panels are attached to each other and to a primary bus 
structure by antenna deployment hinges. The upper portion of the satellite 
incorporates an array of individual intersatellite antennas, which are 
always pointed tangentially to the Earth and towards the intersatellite 
antenna arrays of other satellites within the satellite constellation. An 
Astromast.TM. boom is mounted between the space-facing surface of the 
primary bus structure and an assembly of solar array storage booms. The 
Astromast.TM. boom can expand and rotate amorphous silicon solar arrays 
which are unfurled from within the solar array storage booms. The 
amorphous silicon solar arrays gather solar radiation to provide power the 
satellite, and also provide thermal control for the satellite by shielding 
the satellite from solar radiation. The satellite is capable of being 
nested or stacked in a compact arrangement that fits within a payload bay 
of a launch vehicle. 
An appreciation of other aims and objectives of the present invention and a 
more complete and comprehensive understanding of this invention may be 
achieved by studying the following description of a preferred embodiment 
and by referring to the accompanying drawings.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT 
FIG. 1 is an illustration of the Callingsat 10 shown as it would appear in 
its fully opened and deployed position in low Earth orbit. An Earth-facing 
antenna array 12 is used to provide reception of 30 GHz uplinks 14 and 
also to provide transmission of 20 GHz downlinks 16. The Earth-facing 
antenna array 12 comprises individual electronically steered phased-array 
antennas 18 located on eight mobile, fixed terminal satellite link M/FTSL 
antenna panel sets 20 and on the Earth-facing surface of the primary bus 
structure 22. Each M/FTSL antenna panel set 20 has three adjoining antenna 
facet panels 24. The M/FTSL panels are deployed at angles with respect to 
the Earth's surface that limit the required steering angle from the 
satellite to the portion of the Earth surface served by this antenna. Four 
intersatellite link ISL antenna arrays 26 are located on the space-facing 
surface of the primary bus structure 22. Four individual ISL antennas 28 
make up each of the ISL antenna arrays 26. The ISL antennas 28 are able to 
receive and to transmit 60 GHz intersatellite links 30. The 60 GHz 
intersatellite links 30 provide communication among the constellation of 
Callingsats. 
An expandable Astromast.TM. boom 32 such as that produced by Astro-Spar of 
Carpenteria, Calif., is mechanically coupled to the primary bus structure 
22, and at full extension can reach approximately 12 meters in length. A 
boom crossmast 34 is mechanically coupled to the far end of the 
Astromast.TM. boom 32. A solar array storage boom 36 is attached to the 
boom crossmast 34. Two more solar array storage booms 36 are coupled to 
opposite ends of the first solar array storage boom 36. An amorphous 
silicon solar array 38 extends from each of the solar array storage booms 
36. Inflatable booms 40 extend from each end of the solar array storage 
booms 36, and are attached to the amorphous silicon solar arrays 38. A 
cantilever boom 42 is attached to the far end of each of the amorphous 
silicon solar arrays 38, between each pair of inflatable booms 40. 
Structural support for each of the thin, generally rectangular, amorphous 
silicon solar arrays 38 is provided by the framework created by each solar 
array storage boom 36, a pair of inflatable booms 40, and a cantilever 
boom 42. 
Pulse plasma thrusters 44, storage batteries 46, and shunt regulators 48 
are attached to the boom crossmast 34. Propulsion for the Callingsat 10 is 
provided by six pulse plasma thrusters 44, which in this embodiment are 
produced by Olin RRC. The pulse plasma thrusters 44 provide propulsion to 
accomplish maneuvers such as orbit insertion, drag make-up, station 
keeping, and de-orbit that is required at the end of the lifetime of a 
Callingsat 10. Each of the pulse plasma thrusters 44 is designed to 
provide 60 kN * sec of thrust. This configuration affords the advantages 
of redundant propulsion, and also insures reliable service for the 
lifetime of the satellite 10. 
FIG. 2 is a perspective view 50 of the assembled Callingsat 10, before it 
is launched and deployed in low Earth orbit. Many of the components of the 
Callingsat 10 are designed to be stored, folded, or manipulated to achieve 
as small a volume as possible. This design provides an extremely compact 
structure that can be delivered into space at the lowest cost. Four M/FTSL 
antenna panel sets 20 located at opposing corners of the primary bus 
structure 22 are folded inward to become the inner antenna arrays 52. The 
remaining four M/FTSL antenna panel sets 20 that are also attached to the 
primary bus structure 22 are also folded inward to become the outer 
antenna arrays 54. 
Two component storage compartments 56 are located on the upper surface of 
the primary bus structure 22. Two ISL antenna arrays 26 are attached to 
the exterior of the component storage compartments 56. The amorphous 
silicon solar arrays 38 and inflatable booms 40 are stored inside their 
respective solar array storage booms 36, which are folded together and are 
mounted to the primary bus structure 22 by two modular aluminum solar 
array attachment structures 58. Two additional ISL antenna arrays 26 are 
attached to the solar array attachment structures 58. The component 
storage compartments 56 are used to house much of the required internal 
systems, such as the command and data handling subsystem 60, the 
attitude/orbit determination and control subsystem 62, and the 
communications payload subsystem 64. 
FIG. 3 provides a front view 66 of the assembled Callingsat 10, which shows 
how the M/FTSL antenna panel sets 20 are folded underneath the primary bus 
structure 22. FIG. 3 also illustrates how the amorphous silicon solar 
arrays 38 are stored effectively within the solar array storage booms 36. 
The scale of the assembled Callingsat 10 is suggested by a person P 
standing approximately two meters tall alongside the Callingsat 10. 
FIG. 4 offers a detailed side view 68 of the assembled Callingsat 10. From 
this view, details of one of the three solar array storage booms 36 can be 
seen. Cantilever booms 42 extend between both ends of the solar array 
storage booms 36. Before deployment, the amorphous silicon solar arrays 38 
and the inflatable booms 40 are stored within the solar array storage 
booms 36. 
FIG. 5 is a front view 70 of an assembled Callingsat 10 before being placed 
into a launch vehicle LV. After a final inspection, the Callingsat 10 is 
placed in the payload dynamic envelope PDE within the interior LVI of the 
launch vehicle LV, as seen in the cutaway top view 74 illustrated in FIG. 
6. FIG. 7 provides a bottom cutaway view 78 of a Callingsat 10 stored 
within the payload dynamic envelope PDE in the launch vehicle LV. This 
view shows how the geometric design for the antenna facet panels 24 is 
chosen to supply a large amount of surface area for placement of antennas 
18, while simultaneously providing an extremely compact folded structure. 
FIG. 8 provides a side cutaway view 80 of a launch vehicle LV with a 
payload of Callingsats 10. A number of Callingsats 10 are located within 
the payload dynamic envelope PDE area of the interior LVI of the launch 
vehicle LV. The advanced design techniques employed by the present 
invention allow a large quantity of Callingsats 10 to be placed within the 
payload dynamic envelope PDE, which insures that each Callingsat 10 is 
deployed at minimum cost. To provide a sense of scale, a person P 
approximately two meters tall is shown standing near the launch vehicle LV 
before it is launched. 
FIG. 9 is an illustration 82 of a Callingsat 10 being deployed into a low 
Earth orbit, after it has been launched from the launch vehicle LV. The 
launch vehicle LV is capable of deploying Callingsats 10 either one at a 
time or in multiple groups. The attitude/orbit determination and control 
subsystem 62 located within the Callingsat 10 determines its current 
location, and compares it with the intended orbit location above the 
Earth. The attitude/orbit determination and control subsystem 62 then 
provides the control logic necessary to activate the pulse plasma 
thrusters 44 to guide the spacecraft 10 into its correct orbit location. 
As the guided Callingsat 10 approaches correct orbit location above the 
Earth E, the attitude/orbit determination and control subsystem 62 
provides the logic necessary to begin deployment of the Astromast.TM. boom 
32 and the solar array storage booms 36. FIG. 10 is an illustration 84 
that shows how the solar array storage booms 36 are disconnected from the 
solar array attachment structures 58, using spring-loaded, pyrotechnic, or 
other suitable means to promote detachment. Upon detachment, the 
Astromast.TM. boom 32 rotates axially about its couplings with the primary 
bus structure 22 and the boom crossmast 34, while simultaneously expanding 
lengthwise to move the solar array storage booms 36 away from the primary 
bus structure 22. 
FIG. 11 provides a view 86 of a deployed Callingsat 10 in which the 
Astromast.TM. boom 32 has expanded to distance the solar array storage 
booms 36 from the primary bus structure 22. The attitude/orbit 
determination and control subsystem 62 provides the adjustment for pitch 
and roll reaction to the inertial moment created by the distance of the 
solar array storage booms 36 from the primary bus structure 22. 
The Astromast.TM. boom 32 continues to extend to its full length. The 
satellite antennas and electronics operate most efficiently at low 
temperatures. They can most effectively dissipate the heat generated in 
their operation by radiating this heat via the back side of the M/FTSL 
antenna panel sets 20 into "cold space" which has an ambient temperature 
of approximately 4 degrees Kelvin. If the radiation of the suns is allowed 
to impinge directly on the back surface of the M/FTSL antenna panel sets, 
this increases the effective ambient temperature, which would result in 
less efficient thermal dissipation, a higher operating temperature, and 
less efficiency. The present design uses the amorphous silicon solar 
arrays 38 as a sun shade to effectively shield the Callingsat 10 from the 
sun and thus to reduce the effective ambient temperature of the space into 
which the Callingsat 10 dissipates heat. This is done by using amorphous 
silicon solar arrays 38 that when fully extended on the Astromast.TM. boom 
32 between the Callingsat 10 and the suns and oriented perpendicular to 
the suns, casts a shadow that completely covers the Callingsat 10, 
including all antenna facet panels 24. The amorphous silicon solar array 
position that provides maximum shading also provides the most efficient 
solar energy generation, since the array surface is maintained 
perpendicular to the sun's rays. The extension of the Astromast.TM. boom 
32 provides sufficient distance between the primary bus structure 22 and 
the amorphous silicon solar arrays 38, which, in turn, furnishes the 
required radiation shielding for the antennas 18 and the ISL antennas 28. 
In this embodiment, the Astromast.TM. boom 32 is approximately twelve 
meters long when fully extended. 
FIG. 12 is a illustration 88 which depicts the extension of the outer 
antenna arrays 54. The outer antenna arrays 54 are made up of four outer 
antenna arrays 54a, 54b, 54c, and 54d. They are located at four opposing 
corners of the primary bus structure 22. The outer antenna arrays 54 are 
each made up of an inner antenna panel 92, a central antenna panel 94, and 
an outer antenna panel 96. Each outer antenna panel 96 coupled in series 
to a pair of antenna deployment hinges 90, a central antenna panel 94, 
another pair of antenna deployment hinges 90, an inner antenna panel 92, 
another pair of antenna deployment hinges 90, and to the primary bus 
structure 22. When the outer antennas 54 are expanded, the outer antenna 
panels 96, the central antenna panels 94, and the inner antenna panels 92 
are all unfolded, using the antenna deployment hinges 90. 
FIG. 12 also illustrates the expansion of the solar array storage booms 36 
as the Astromast.TM. boom 32 reaches its fully extended length. The solar 
array storage booms 36 are coupled to each other using boom hinge 
mechanisms 98. The solar array storage booms 36 use the boom hinge 
mechanisms 98 to pivot away from their parallel stored position towards a 
coaxial arrangement. 
FIG. 13 is an illustration 100 which shows full extension solar array 
storage booms 36 and the outer antenna arrays 54. The solar array storage 
booms 36 come into coaxial alignment with each other as the boom hinge 
mechanisms 98 that couple them pivot to their extended positions. Each end 
of each solar array storage boom 36 is coupled to a deployment mechanism 
108. In FIG. 13 the cantilever booms 42 can be seen as they would be 
attached to the solar array storage booms 36 in a parallel fashion between 
the deployment mechanisms 108. The boom crossmast 34 rotates to position 
the solar array storage booms 38 correctly for expansion of the amorphous 
silicon solar arrays 36. The pulse plasma thrusters 44, the storage 
batteries 46, and the shunt regulators 48 are seen as they are located on 
the boom crossmast 34. 
As the outer antenna arrays 54a, 54b, 54c, and 54d continue to unfold away 
from the primary bus structure 22, they begin to resemble the petals of an 
oblate flower. The outer antenna array sets 54 are locked in place when 
the antenna deployment hinges 90 reach the end of their designed travel. 
FIG. 13 also illustrates how the inner antenna arrays 52 begin to unfold as 
the outer antenna arrays 54 expand away from the primary bus structure 22. 
The inner antenna arrays 52 are made up of four inner antenna arrays 52a, 
52b, 52c, and 52d located at four opposing corners of the primary bus 
structure 22, in between each of the outer antenna arrays 54. The inner 
antenna arrays 52 are each made up of an inside antenna panel 102, a 
middle antenna panel 104, and an outside antenna panel 106. Each outside 
antenna panel 106 is coupled in series to a pair of antenna deployment 
hinges 90, a middle antenna panel 104, another pair of antenna deployment 
hinges 90, an inside antenna panel 102, another pair of antenna deployment 
hinges 90, and to the primary bus structure 22. The inside antenna panels 
102 are attached to the primary bus structure 22 at a different offset 
distance than the inner antenna panels 92 of the outer antenna arrays 54. 
This offset allows the inner antenna arrays 52 and the outer antenna 
arrays 54 to be advantageously stored in a minimum volume. 
The expansion of the inner antenna arrays 52 is similar to the expansion of 
the outer antenna arrays 54, and entails the unfolding of the outside 
antenna panels 106, the middle antenna panels 104, and the inside antenna 
panels 102 using the antenna deployment hinges 90. The large surface area 
provided by the inner antenna arrays 52 and the outer antenna arrays 54 
provides adequate room for the numerous antennas 18 that make up the 
Earth-facing antenna array 12, which is designed for large communications 
traffic within the Satellite Communications System. 
FIG. 14 is an illustration 110 of the inflation and expansion of the 
amorphous silicon solar arrays 38 from the solar array storage booms 36. 
The specialized structure of the amorphous silicon solar arrays 38 and the 
inflatable booms 40 enable the large, extremely light weight, amorphous 
silicon solar arrays 38 to be deployed from the Callingsat 10 economically 
and reliably. 
The amorphous silicon solar arrays 38 comprise of photovoltaic cells 112 
located on a thin film substrate 114. New polymer films, such as Mylar.TM. 
or Kapton.TM. are used for the film substrate 114, which is attached to 
the photovoltaic cells 112. Mylar.TM. and Kapton.TM. satisfy all the 
material requirements of the film substrate 114, and offer a useful 
lifetime of many years in outer space. 
The inflatable booms 40 attached to the amorphous silicon solar arrays 38 
are expanded by gas pressure provided by inflation gas 118 from within the 
deployment mechanisms 108. The inflatable booms 40 expand away from the 
deployment mechanisms 108, and are used to unfurl the amorphous silicon 
solar arrays 38 from their stored positions within the solar array storage 
booms 36. In one embodiment, the inflatable booms 40 have pleated, 
"accordion bellows" style structures. In another embodiment, the 
inflatable booms 40 roll out to unfurl the amorphous silicon solar arrays 
38, in a similar fashion to a child's party favor. The inflatable booms 40 
provide an extremely effective, light weight means for supporting the 
light and flexible amorphous silicon solar arrays 38. 
Once the inflatable booms 40 are fully inflated to unfurl the amorphous 
silicon solar arrays 38, they are then rigidized to retain their inflated 
structure. A small amount of photocurable chemical vapor 116 is deposited 
into the inflatable booms 40 from within the deployment mechanism 108. The 
photocurable chemical vapor 116 mixes with the inflation gas 118, and 
cures on the inner surface of the inflatable booms 40 when it is exposed 
to ultraviolet radiation provided by the Sun S. As the photocurable 
chemical vapor 116 cures, it forms a rigid surface on the inner walls of 
the inflatable booms 40. The rigid inflatable booms 40 then act together 
with the solar array storage booms 36 and the cantilever booms 42 to 
provide a long lasting ultralightweight framework for the amorphous 
silicon solar arrays 38. 
Other embodiments of the present invention employ different techniques to 
provide an efficient framework for the amorphous silicon solar arrays 38. 
In one embodiment, continuous gas pressure is provided to provide 
sufficient rigidity for the inflatable booms 40. In another embodiment, 
the amorphous silicon solar arrays 38 and inflatable booms 40 are combined 
to form an inflatable pillow-like structure that can be rigidized with 
photocurable chemical vapor 116. 
FIG. 15 provides a perspective view 120 of the fully deployed amorphous 
silicon solar arrays 38 on the Callingsat 10. The amorphous silicon solar 
arrays 38 provide power for the spacecraft 10 through the photovoltaic 
cells 112. The photovoltaic energy is produced directly from incoming 
solar energy when photons are absorbed in the semiconductor substrate 
photovoltaic cells 112. Amorphous silicon is used in the preferred 
embodiment, since its irregular, non-crystalline arrangement allows highly 
efficient light absorption in an ultra-thin film. This amorphous silicon 
film is very lightweight, and is flexible, reliable, and extremely 
resistant to physical abuse. Excess power provided by the photovoltaic 
cells 112 is stored in the storage batteries 46. The amorphous silicon 
solar arrays 38 also function as a solar heat shield for the Callingsat 
10, and provide an extremely lightweight and effective active thermal 
control device. 
Many components in an orbiting satellite function properly only if they are 
maintained within a rather narrow temperature range. The temperatures of 
satellite components are influenced by the net thermal energy exchange 
between the satellite and its thermal environment, which is influenced by 
the magnitude and distribution of radiation input from the Sun S and the 
Earth E. 
The objective of satellite thermal control design is to provide the proper 
heat transfer between all satellite elements so that temperature sensitive 
components remain within their specified temperature limits. Techniques 
used for satellite thermal control can be passive or active. Passive 
techniques include thermal coatings, thermal insulation and heat sinks. 
Active thermal control techniques include heat pipes, louvers, heat 
shields, and electrical heaters. 
A satellite heat pipe is a thermal device that can provide a significant 
transfer of thermal energy between to regions on the satellite. A heat 
pipe comprises a closed cylinder whose inner surfaces are lined with a 
wick that provides a capillary effect. Heat in the warm portion of the 
pipe vaporizes a working fluid. The resulting pressure difference drive 
the vapor to the cooler end of the tube, where the vapor condenses and 
releases its latent heat of vaporization. The loss of liquid in the warmer 
area creates a capillary pressure that promotes movement of liquid from 
the cooler region to the warmer region, thus creating the continuous cycle 
required for heat transfer. 
FIG. 16 is a side view 122 of a fully extended and deployed Callingsat 10. 
This depiction shows how the advanced design of the Callingsat 10 provides 
a large Earth-facing antenna array 12 to provide a large volume of 30 GHz 
uplinks 14 and 20 GHz downlinks 16, and ISL antenna arrays 26 to provide a 
large volume of 60 GHz intersatellite links 30. This view also illustrates 
how the expandable Astromast.TM. boom 32 is extended to position the 
amorphous silicon solar arrays 38 correctly. 
FIG. 17 is a from view 124 of a fully deployed Callingsat 10. This view 
portrays the large amorphous silicon solar arrays 38 and shows how they 
provide thermal heat shielding for the Callingsat 10. The amorphous 
silicon solar arrays 38 supply a large surface area of photovoltaic cells 
112 which collect solar radiation and provide power for the Callingsat 10. 
FIG. 18 is an illustration 126 of the fully deployed Earth-facing antenna 
array 12 used in the present invention. The antennas 18 located on the 
inner antenna arrays 52, the outer antenna arrays 54, and the primary bus 
structure 22 consist of uplink antennas 18a and downlink antennas 18b. The 
uplink antennas 18a are used to receive 30 GHz uplinks 14, and the 
downlink antennas 18b are used to transmit 20 GHz downlinks 16. 
FIG. 19 is an alternate embodiment 128 of the Earth-facing antenna array 
12. This embodiment features a different arrangement of uplink antennas 
18a and downlink antennas 18b that are located on the inner antenna arrays 
52, the outer antenna arrays 54, and the primary bus structure 22. FIG. 20 
illustrates another embodiment 130 for the Earth-facing antenna array 12. 
The pattern shown in FIG. 20 utilizes uplink antennas 18a and downlink 
antennas 18b that reside on the inner antenna arrays 52, the outer antenna 
arrays 54, and the primary bus structure 22. 
Although the present invention has been described in detail with reference 
to a particular preferred embodiment, persons possessing ordinary skill in 
the art to which this invention pertains will appreciate that various 
modifications and enhancements may be made without departing from the 
spirit and scope of the claims that follow. 
LIST OF REFERENCE CHARACTERS 
10 Callingsat satellite 
12 Earth-facing antenna array 
14 30 Ghz uplinks 
16 20 GHz downlinks 
18 Antennas 
18a Uplink antennas 
18b Downlink antennas 
20 M/FTSL antenna panel set 
22 Primary bus structure 
24 Antenna facet panels 
26 ISL antenna array 
28 Individual ISL antenna 
30 60 GHz intersatellite links 
32 Astromast.TM. boom 
34 Boom crossmast 
36 Solar array storage boom 
38 Amorphous silicon solar array 
40 Inflatable boom 
42 Cantilever boom 
44 Pulse plasma thrusters 
46 Storage batteries 
48 Shunt regulators 
50 Perspective view of assembled Callingsat 
52 Inner antenna array 
52A First inner antenna array 
52B Second inner antenna array 
52C Third inner antenna array 
52D Fourth inner antenna array 
54 Outer antenna array 
54A First outer antenna array 
54B Second outer antenna array 
54C Third outer antenna array 
54D Fourth outer antenna array 
56 Component storage compartment 
58 Solar array attachment structure 
60 Command and data handling subsystem 
62 Attitude/orbit determination and control subsystem 
64 Communications payload subsystem 
66 Front view of assembled Callingsat 
68 Side view of assembled Callingsat 
70 Front view of Callingsat before launch 
74 Top cutaway view of Callingsat in launch vehicle 
80 Bottom cutaway view of Callingsat within launch vehicle 
82 Illustration of Callingsat deployed in low Earth orbit 
84 Illustration of deployed Callingsat with detached solar array storage 
booms 
86 Illustration of deployed Callingsat with extended Astromast.TM. boom 
88 Depiction of expanding solar array boom and extension of outer antenna 
arrays 
90 Antenna deployment hinges 
92 Inner antenna panel on outer array 
94 Central antenna panel on outer array 
96 Outer antenna panel on outer array 
98 Boom hinge mechanism 
100 Depiction of extended solar array boom and extension of inner antenna 
arrays 
102 Inside antenna panel on inner array 
104 Middle antenna panel on inner array 
106 Outside antenna panel on inner array 
108 Deployment mechanism 
110 Depiction of inflating solar arrays 
112 Photovoltaic cells 
114 Film substrate 
116 Photocurable chemical vapor 
118 Inflation gas 
120 Fully deployed solar array 
122 Side view of fully deployed Callingsat 
124 Front view of fully deployed Callingsat 
126 Illustration of fully deployed earth-facing antenna arrays 
128 Alternate embodiment of earth facing antenna arrays 
130 Another embodiment of earth facing antenna arrays 
E Earth 
P Person 
LV Launch vehicle 
LVI Launch vehicle interior 
PDE Payload dynamic envelope 
S Sun