U.S. patent application number 10/348365 was filed with the patent office on 2003-07-17 for orbit transfer vehicle with support services.
Invention is credited to Fleeter, Richard, Gloyer, Paul, Goldstein, David, McDermott, Scott A..
Application Number | 20030132350 10/348365 |
Document ID | / |
Family ID | 25451378 |
Filed Date | 2003-07-17 |
United States Patent
Application |
20030132350 |
Kind Code |
A1 |
Goldstein, David ; et
al. |
July 17, 2003 |
Orbit transfer vehicle with support services
Abstract
The excess space and weight capacity of a conventional launch
vehicle for a high-energy orbit, such as GEO, is used to deploy
satellites to a low-energy orbit, such as LEO. In a preferred
embodiment, an orbit-transfer vehicle provides the navigation,
propulsion, and control systems required to transport a payload
satellite from a high-energy-transfer orbit, such as GTO, to a
predetermined low-energy orbit. Upon entering the low-energy orbit,
the payload satellite is released from the orbit-transfer vehicle.
To reduce the fuel requirements for this deployment via the
orbit-transfer vehicle, a preferred embodiment includes aerobraking
to bring the satellite into a low-earth orbit. In a preferred
embodiment of this method of deployment, the provider of the
orbit-transfer vehicle identifies and secures available excess
capacity on launch vehicles, and allocates the excess capacity to
the satellites requiring low-earth orbit deployment, thereby
providing a deployment means that is virtually transparent to the
purchaser of this deployment service.
Inventors: |
Goldstein, David;
(Roslindale, MA) ; McDermott, Scott A.;
(Washington, DC) ; Gloyer, Paul; (Waveland,
MS) ; Fleeter, Richard; (Reston, VA) |
Correspondence
Address: |
ROBERT M. MCDERMOTT, ESQ.
1824 FEDERAL FARM ROAD
MONTROSS
VA
22520
US
|
Family ID: |
25451378 |
Appl. No.: |
10/348365 |
Filed: |
January 21, 2003 |
Related U.S. Patent Documents
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Application
Number |
Filing Date |
Patent Number |
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10348365 |
Jan 21, 2003 |
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09925206 |
Aug 9, 2001 |
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6561461 |
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09925206 |
Aug 9, 2001 |
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09350813 |
Jul 9, 1999 |
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6286787 |
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Current U.S.
Class: |
244/158.7 |
Current CPC
Class: |
B64G 1/007 20130101;
B64G 1/641 20130101; B64G 1/242 20130101; B64G 1/222 20130101; B64G
1/50 20130101; B64G 1/428 20130101; B64G 1/10 20130101; B64G
2001/224 20130101; B64G 1/425 20130101; B64G 1/244 20190501; B64G
1/1085 20130101 |
Class at
Publication: |
244/158.00R |
International
Class: |
B64G 001/00; B64G
001/22 |
Claims
We claim:
1. A method of deploying a payload satellite into a target orbit
having an associated target-orbit energy-level, comprising:
attaching the payload satellite to an orbit-transfer vehicle,
attaching the orbit-transfer vehicle and the payload satellite to a
launch vehicle, launching the launch vehicle, deploying the
orbit-transfer vehicle with the attached payload satellite from the
launch vehicle with an associated kinetic energy, the kinetic
energy associated with the orbit-transfer vehicle being
substantially different from the target-orbit energy-level
associated with the target orbit, and maneuvering the
orbit-transfer vehicle into the target orbit via a change in the
kinetic energy associated with the orbit-transfer vehicle, to an
energy level corresponding to the target-orbit energy-level,
wherein the maneuvering of the orbit-transfer vehicle is effected
via a plurality of functional components, and at least a subset of
the functional components are also used by the payload
satellite.
2. The method of claim 1, wherein the subset of functional
components used by the payload satellite includes at least one of:
a communications device, a power regulation device, an energy
storage device, a heating system, a navigation device, and an
attitude control device.
3. The method of claim 1, further including detaching at least some
components of the orbit-transfer vehicle from the payload satellite
when the payload satellite is maneuvered to the target orbit.
4. The method of claim 1, further including aerobraking the payload
satellite to effect at least a portion of the change in the kinetic
energy associated with the payload satellite.
5. The method of claim 1, further including applying thrust to the
payload satellite to effect at least a portion of the change in the
kinetic energy associated with the payload satellite.
6. A method of facilitating the deployment of a payload satellite
into a target orbit having an associated target-orbit energy-level,
comprising: identifying an excess capacity on a scheduled launch
vehicle having an associated transfer orbit that has an associated
orbit-transfer energy level that is substantially different than
the target-orbit energy-level, facilitating an attachment of the
payload satellite to an orbit-transfer vehicle, facilitating an
attachment of the orbit-transfer vehicle with payload satellite to
the launch vehicle, facilitating a deployment of the orbit-transfer
vehicle into the transfer orbit via the launch vehicle, the
orbit-transfer vehicle thereby having a kinetic energy
corresponding to the orbit-transfer energy level, facilitating a
maneuvering of the orbit-transfer vehicle to the target orbit via a
substantial change in the kinetic energy of the orbit-transfer
vehicle, from the orbit-transfer energy level to the target-orbit
energy-level, and facilitating a sharing of functional components
used for maneuvering the orbit-transfer vehicle with the payload
satellite.
7. The method of claim 6, wherein the functional components used by
the payload satellite includes at least one of: a communications
device, a power regulation device, an energy storage device, a
navigation device, and an attitude control device.
8. The method of claim 6, further including facilitating an
aerobraking of the orbit-transfer vehicle so as to achieve the
target orbit.
9. An orbit-transfer vehicle comprising: a coupling that
facilitates an attachment of a payload satellite to the
orbit-transfer vehicle, a propulsion system that is configured to
facilitate maneuvering the payload satellite from a first orbit to
a target orbit, and a spacecraft kernel that is designed
independent of the orbit-transfer vehicle, and independent of the
payload satellite, and is configured to provide an integrated
communications system for use by the orbit-transfer vehicle.
10. The orbit-transfer vehicle of claim 9, wherein the spacecraft
kernel is further configured to provide regulated power to the
orbit-transfer vehicle.
11. The orbit-transfer vehicle of claim 10, wherein the spacecraft
kernel is further configured to provide regulated power to the
payload satellite.
12. The orbit-transfer vehicle of claim 9, wherein the spacecraft
kernel is further configured to provide communications to the
orbit-transfer vehicle.
13. The orbit-transfer vehicle of claim 12, wherein the spacecraft
kernel is further configured to provide regulated power to the
payload satellite.
14. The orbit-transfer vehicle of claim 9, further including a
second coupling that facilitates attaching the orbit-transfer
vehicle to a launch vehicle, the launch vehicle being configured to
deploy the orbit-transfer vehicle to the first orbit.
15. The orbit-transfer vehicle of claim 9, wherein the kernel is
configured to remain with the payload satellite when the payload
satellite is placed in the target orbit.
Description
CROSS-REFERENCE TO RELATED APPLICATIONS
[0001] This is a Divisional of U.S. patent application Ser. No.
09/925,206, filed Aug. 9, 2001, which is a Continuation-In-Part of
U.S. patent application Ser. No. 09/350,813, filed Jul. 9,
1999.
BACKGROUND OF THE INVENTION
[0002] 1. Field of the Invention
[0003] This invention relates to the field of aerospace, and
generally to the deployment of small satellites via an orbit
transfer vehicle.
[0004] 2. Description of Related Art
[0005] Satellites, because of their unobstructed fields of view of
large areas of the earth, are often the preferred technical
solutions to a variety of communications and monitoring problems as
well as space and earth science applications. The high cost of
satellite deployment, however, often precludes their use from a
cost-efficiency viewpoint. In order to distribute the high costs of
deployment among a large number of applications or users, the
majority of deployed satellites are those that handle a multitude
of tasks, or a multitude of customers for the same task. To
minimize the loss of available access time to or from the satellite
from or to the location on earth being serviced by a satellite,
many satellites are placed in geosynchronous orbit. A
geosynchronous orbit tracks the revolution of the earth, so that
the satellite appears to be fixed over the same area of the earth,
thus providing continual access to that area. Due to the physics
involved, a geosynchronous orbit is approximately 36,000 kilometers
above the earth. To provide reliable communications over this long
distance, a geosynchronous satellite requires highly sensitive
receivers and/or highly powerful transmitters. Because of the
aforementioned economic and technical requirements, geosynchronous
satellites generally weigh a ton or more, and cost (in 1999)
hundreds of millions of dollars to deploy to the selected
geosynchronous orbit.
[0006] As contrast to large geosynchronous satellites, the use of
small satellites at low-earth orbit (LEO) are becoming increasingly
common. Copending U.S. patent applications "SATELLITE COMMUNICATION
SYSTEM WITH A SWEEPING HIGH-GAIN ANTENNA", U.S. Ser. No.
09/045,971, filed Mar. 21, 1998 for Fleeter et al, Attorney docket
AA980217; "LOW-COST SATELLITE COMMUNICATION SYSTEM", U.S. Ser. No.
09/045,970, filed Mar. 21, 1998 for Fleeter et al, Attorney docket
AA980218; "IN SITU REMOTE SENSING", U.S. Ser. No. 09/130,854, filed
Aug. 7, 1998 for Richard Fleeter, Attorney docket AA980723; "RF
INSPECTION SATELLITE", U.S. Ser. No. 09/267,942, filed Mar. 11,
1999 for Hanson et al, Attorney docket AA990202 illustrate the use
of low cost satellites for a variety of applications, and are
incorporated by reference herein. Low-earth orbits are typically
hundreds of miles above the earth, rather than thousands of miles.
Because of their order of magnitude closer proximity to earth,
satellites in low-earth orbit require significantly less
communicating and monitoring power and sensitivity than the
satellites in geosynchronous orbit. Because they are not stationary
above any location on the earth, multiple satellites in low-earth
orbit are required to provide continuous coverage of a particular
area on earth. Because multiple satellites are required in
low-earth orbit to provide continuous coverage, a low-earth orbit
satellite system is particularly well suited to applications that
employ low cost satellites. As advances continue to be made in
electronic circuit density and efficiency, the number of
communication and monitoring applications that can be embodied in
small, low cost satellites continues to increase.
[0007] Deployment of a small, less than five hundred pound,
satellite into low-earth orbit typically costs, in 1999 dollars,
between seven and ten million dollars. Because a plurality of
satellites is required to provide continuous coverage of an area,
the overall cost of deploying constellations of low-earth orbit
satellites can often amount to hundreds of millions of dollars.
BRIEF SUMMARY OF THE INVENTION
[0008] It is an object of this invention to provide a lower cost
means for deploying a satellite into orbit. It is a further object
of this invention to provide a method for economically brokering
the deployment of a satellite. It is a further object of this
invention to provide an orbit-transfer vehicle to effectively
deploy small satellite systems. It is a further object of this
invention to reduce the cost and complexity of small satellite
systems.
[0009] A launch of geosynchronous satellites typically includes one
or two large, multi-ton, satellites that are deployed at the
geosynchronous altitude of 36,000 kilometers via a large
multi-stage rocket system, such as an Ariane system. Typically,
after allocating the available space and weight capabilities of the
rocket system to the primary payload of the one or two large
satellites, some excess space and weight allocation remains. For
example, if an Ariane launch vehicle can accommodate four tons, and
the primary payload satellites are 11/2 and 2 tons each, the launch
vehicle has an excess capacity of a half ton. Because the marginal
cost of adding one or two small satellites is minimal, this excess
space or weight capacity can be brokered for the deployment of
small satellites at substantially less cost than the primary
payload, often less than a quarter of the cost per pound charged to
the primary payload satellites.
[0010] The expressed objects of this invention, and others, are
achieved by providing a means of utilizing the excess space and
weight capacity that is typical of a launch of large geosynchronous
satellites to deploy small satellites at a low-earth orbit.
Specifically, this invention provides a method of deployment of
small satellite systems to a target orbit from a
geosynchronous-transfer launch vehicle. In a preferred embodiment,
an orbit-transfer vehicle provides the navigation, propulsion, and
control systems required to transport a payload satellite from a
geosynchronous-transfer orbit (GTO) to a predetermined low-earth
orbit (LEO). Upon entering low-earth orbit, the payload satellite
is deployed from the orbit-transfer vehicle. To reduce the cost and
complexity of the payload satellite, the orbit-transfer vehicle is
configured to provide common functional services, such as
communications and power regulation, to the payload satellite
during the transport, and/or after deployment. To reduce the fuel
requirements for this deployment via the orbit-transfer vehicle, a
preferred embodiment includes aerobraking to bring the satellite
into a low-earth orbit. In a preferred embodiment of this method of
deployment, the provider of the orbit-transfer vehicle identifies
and secures available excess capacity on launch vehicles, and
allocates the excess capacity to the satellites requiring
deployment, thereby providing a deployment means that is virtually
transparent to the purchaser of this deployment service.
BRIEF DESCRIPTION OF THE DRAWINGS
[0011] The invention is explained in further detail, and by way of
example, with reference to the accompanying drawings wherein:
[0012] FIG. 1 illustrates a conventional deployment of a
geosynchronous satellite via a geosynchronous-transfer launch
vehicle.
[0013] FIG. 2 illustrates an example deployment of a low-earth
orbit (LEO) satellite using a geosynchronous-transfer launch
vehicle in accordance with this invention.
[0014] FIG. 3 illustrates an example method of facilitating the use
of excess capacity of a launch vehicle for deploying a satellite in
accordance with this invention.
[0015] FIG. 4 illustrates an example orbit-transfer vehicle in
accordance with this invention.
[0016] FIG. 5 illustrates an alternative example orbit-transfer
vehicle in accordance with this invention.
DETAILED DESCRIPTION OF THE INVENTION
[0017] For ease of reference and understanding, this invention is
presented using the paradigm of a deployment of a small satellite
system to Low-Earth Orbit (LEO), via a Geosynchoronous Transfer
Orbit (GTO), although one of ordinary skill in the art will
recognize that the principles of this invention are not limited to
a GTO-to-LEO transfer.
[0018] FIG. 1 illustrates a conventional deployment of a
geosynchronous satellite via a geosynchronous-transfer launch
vehicle. Note that the figures are presented herein for
illustration purposes. Although the relative size of the
illustrated orbits are approximately to scale relative to the
illustrated size of the earth 130, the specific details of the
orbits should not be interpreted as being representative. A
geosynchronous-transfer launch vehicle (not shown) containing one
or more satellites for deployment to geosynchronous-earth orbit
(GEO) 150 is launched 101 from the earth 130. The
geosynchronous-transfer launch vehicle is configured to achieve a
geosynchronous-transfer orbit 120 whose apogee 121 is tangent to
the geosynchronous orbit 150. Some time after achieving the
geosynchronous-transfer orbit 120, the one or more geosynchronous
satellites (not shown) are released from the
geosynchronous-transfer launch vehicle. All items that are released
by the geosynchronous-transfer launch vehicle have the same
velocity as the geosynchronous-transfer launch vehicle when they
are released, and therefore continue to travel in the
geosynchronous-transfer orbit 120. Each geosynchronous satellite
contains a means for applying thrust 102, preferably at or near
apogee 121, to enter an intended tangential geosynchronous orbit
150.
[0019] FIG. 2 illustrates an example method of deploying a
low-earth orbit (LEO) satellite using a geosynchronous-transfer
launch vehicle in accordance with this invention. As noted above, a
geosynchronous orbit 150 is located 36,000 kilometers above the
earth 130. A low-earth orbit 110, on the other hand, is located
well within 1500 kilometers of the earth 130. The Space Shuttle,
for example, typically orbits the earth at an altitude between 160
and 600 kilometers.
[0020] In accordance with this invention, an orbit-transfer vehicle
(not shown) containing the low-earth orbit satellite (not shown) is
launched 101 from the earth 130 via a conventional
geosynchronous-transfer launch vehicle (not shown), such as an
Ariane. The geosynchronous-transfer launch vehicle releases the
orbit-transfer vehicle into geosynchronous-transfer orbit 120. Some
time after release, preferably at the apogee 121 of the
geosynchronous-transfer orbit 120, the orbit-transfer vehicle fires
103 its integral propulsion device (not shown) to force itself, and
the attached low-earth orbit satellite, out of the
geosynchronous-transfer orbit 120, and begins a forced descent
toward the earth 130. Another firing 104 of the propulsion device
can thereafter be used to place the orbit-transfer vehicle directly
into a low-earth orbit 110, but such a direct
geosynchronous-transfer orbit to low-earth orbit transfer will
require a substantial amount of fuel to reduce the kinetic energy
of the orbit-transfer vehicle sufficiently to remain in the
low-earth orbit 110.
[0021] In accordance with an aspect of this invention, aerobraking
is used to facilitate the geosynchronous-transfer orbit to
low-earth orbit transfer. Aerobraking uses the friction of the
earth's atmosphere to reduce the kinetic energy of a spacecraft.
Illustrated in FIG. 2, the firing 103 of the propulsion device
provides a thrust to force the orbit-transfer vehicle close to the
earth 130, at 113. In a preferred embodiment, the firing 103 of the
propulsion device is controlled to provide a perigee 113 of under
two hundred kilometers above the earth 130. At a nominal 150
kilometer altitude, the atmosphere of the earth is sufficiently
dense so as to impart a frictional force that reduces the kinetic
energy of the orbit-transfer vehicle, and its attached low-earth
orbit satellite. This reduction in kinetic energy results in an
apogee 114 of the orbit-transfer vehicle that is less than its
original geosynchronous-transfer orbit apogee 121. Upon achieving
apogee 114, the orbit-transfer vehicle is drawn toward the earth
130 again, and reaches perigee 115, also nominally 150 kilometers
from the earth 130. The friction of the earth's atmosphere at this
low altitude again reduces the kinetic energy of the orbit-transfer
vehicle, and the resultant apogee 116 is less than the prior apogee
114. Subsequent traversals 117 of the earth's atmosphere will
continue to reduce the kinetic energy of the orbit-transfer
vehicle, further lowering each subsequent apogee 118. Drogue
devices may be attached to the orbit-transfer vehicle to further
increase the efficiency of the aerobraking process.
[0022] When sufficient kinetic energy is removed from the
orbit-transfer vehicle, the decreasing apogee 111 of the
orbit-transfer vehicle substantially approaches the altitude of the
intended low-earth orbit 110. At this apogee 111, the
orbit-transfer vehicle effects the firing 105 of the integral
propulsion device to force an ascent of the orbit-transfer vehicle
so as to raise the perigee of the orbit-transfer vehicle, and
attached low-earth orbit satellite, beyond the earth's atmosphere,
thereby preventing further decreases of apogee. The preferred
firing 105 places the orbit transfer vehicle, and attached
low-earth orbit satellite, into a symmetric low-earth orbit 110,
with a perigee that is substantially equal to the apogee 111. In a
preferred embodiment of a GTO to LEO transfer, over one hundred
aerobraking orbits are made before firing the integral propulsion
device at 105. Typically, the orbit-transfer vehicle releases the
low-earth orbit satellite at this low-earth orbit 110, and
thereafter the low-earth orbit satellite operates independently, as
it would have, had it been launched directly from the earth 130 to
the low-earth orbit 110. That is, although the low-earth orbit
satellite may travel hundreds of thousands of miles to reach an
orbit 110 that is only a few hundred miles above the earth's
surface, its operation is substantially independent of this rather
circuitous deployment scheme. Optionally, as discussed further
below, some or all of the components used by the orbit-transfer
vehicle may remain with the satellite, to provide services, such as
communications and power regulation, to the satellite after
deployment.
[0023] Because the deployment methods in accordance with this
invention can be effected without affecting the satellite payload,
this invention provides a means for brokering lower cost services
for the deployment of low-earth orbit satellites. Generally, a
satellite is used as a component of a ground-based system, such as
a communications network, a research facility, and the like. The
user, or owner, of the satellite is not necessarily fluent in the
intricacies of rocketry and orbital mechanics. A satellite
deployment broker provides the interface services and support
between the owner of the satellite and the provider of a launch
vehicle. In accordance with the principles of this invention, a
satellite deployment broker can extend the range of potential
low-earth orbit satellite launch vehicle providers to include
geosynchronous-transfer launch vehicle providers having excess
capacity. FIG. 3 illustrates, for example, a flow diagram for
allocating launch services for low-earth orbit satellite
deployments. At 210, the satellite(s) requirements are determined,
including the required orbit parameters, the size and weight of the
satellite, and so on. At 220, the cost of a conventional low-earth
orbit launch that satisfies the requirements are determined or
estimated. At 230, the availability of excess capacity on scheduled
geosynchronous orbit launch vehicles is determined, and a cost is
negotiated for using this excess capacity. The overall cost of
deploying the low-earth orbit satellite via a
geosynchronous-transfer launch is the cost of using the excess
launch capacity of the geosynchronous-transfer launch vehicle plus
the cost of the orbit-transfer vehicle for transporting the
satellite from the geosynchronous orbit to the low-earth orbit, as
determined at 240. If, at 250, the overall cost of the
geosynchronous-transfer launch and orbit-transfer is less than the
conventional low-earth orbit launch, the deployment is effected by
attaching 260 the satellite to the orbit-transfer vehicle and
launching 270 the orbit-transfer vehicle with satellite via the
geosynchronous-transfer launch vehicle. Thereafter, the
orbit-transfer vehicle effects the deployment 280 of the satellite
to a low-earth orbit as discussed above. If, at 250-251, the cost
of the conventional low-earth orbit launch is less expensive than
the geosynchronous-transfer launch, the satellite is deployed 290
via the conventional low-earth orbit launch. Note that a deployment
of a constellation of satellites to low-earth orbits in accordance
with this invention can involve a combination of low-earth orbit
and geosynchronous-transfer launches, depending primarily on the
availability and cost of excess capacity on scheduled
geosynchronous-transfer launch vehicles.
[0024] FIG. 4 illustrates an example orbit-transfer vehicle 300 in
accordance with this invention. The example orbit-transfer vehicle
300 includes an adapter element 310 that provides a conventional
means 315 for securing the vehicle 300 to the
geosynchronous-transfer launch vehicle (not shown), a body element
320 that provides a cavity for holding a fuel cell 340, and a
navigation and control system 330 that provides the navigation,
propulsion, and control systems required to transport the satellite
from a geosynchronous-transfer orbit to a predetermined low-earth
orbit. The example orbit-transfer vehicle 300 also includes support
system components, such as a power regulation system 340, a battery
backup system 350, communications system 360, and so on, discussed
further below.
[0025] The adapter element 310 and body 320 in a preferred
embodiment include shielding to enable the orbit-transfer vehicle
300 to withstand the heat that is induced by the atmospheric
friction during aerobraking. The elevation at perigee 113, 115, 117
determines the required degree of shielding. Correspondingly, the
elevation at perigee 113, 115, 117 determines the number of
aerobraking orbits required to provide a sufficient reduction in
the kinetic energy of the orbit-transfer vehicle 300 to achieve a
low-earth orbit, given the capacity of the fuel cell 340. Thrusters
325 on the body element 320 effect the thrust required to effect
the orbit-transfer, under the control of the navigation and control
system 330. A satellite 380 is designed to be mounted within the
body 320, and is released from the orbit-transfer vehicle 300 when
the appropriate low-earth orbit is achieved. In a preferred
embodiment of this invention, the components 330-360 remain with
the satellite 380, to provide support services to the satellite 380
when the satellite is deployed, as discussed further below.
[0026] Alternative arrangements will be evident to one of ordinary
skill in the art in view of this disclosure. For example, FIG. 5
illustrates an orbit-transfer vehicle 400 for use with a satellite
480 having an integral body with conventional means 315 for
securing the satellite 480 to a launch vehicle (not shown).
[0027] Copending U.S. patent application "A SPACECRAFT KERNEL",
Ser. No. 09/337,154, filed Jun. 21, 1999 for Richard D. Fleeter and
Scott A. McDermott, and incorporated by reference herein, teaches
the use of a kernel system that includes components common to
spacecraft operations, with a well defined interface to
mission-specific components of a satellite system. Conceptually,
the kernel system is provided for common spacecraft functions,
independent of the particular mission, and independent of the
particular spacecraft configuration. For example, recognizing that
communications equipment, power regulation equipment, battery
backup systems, and the like, are typically required in any
spacecraft, the kernel system of the referenced patent application
provides one or more of these functional items.
[0028] In a preferred embodiment of this invention, the orbit
transfer vehicle includes a kernel system that provides the
communications, power management, and other common functions to the
transfer vehicle. In accordance with another aspect of this
invention, this kernel system is also configured to provide these
communications, power management and other common functions to the
payload satellite. That is, when the satellite is placed in the
target orbit, the kernel is configured to remain with the payload
satellite, regardless of whether other components of the transfer
vehicle remain with the payload satellite.
[0029] In addition to reducing the cost and complexity of the
payload satellite by providing these functions, the kernel also
serves to provide required functions before the payload satellite
is deployed. For example, during the orbit transfer, the orbit
transfer vehicle and the payload satellite will typically be
subject to substantial temperature variations, particularly if the
orbit transfer occurs over an extended time duration. By providing
power to the transfer vehicle and the payload satellite, the
temperature within each can be properly regulated via the use of
heaters or other thermal exchange devices. Similarly, by providing
battery-backup power, the transfer vehicle and payload satellite
can be powered while solar panels on these devices are obstructed
from the sun. Also, the power provided to the satellite can be used
to effect a `wake-up` call to the payload satellite when the
satellite is placed in the target orbit.
[0030] The communications functions provided by the kernel are used
in the transfer vehicle for communicating telemetry information to
an earth station, and for receiving commands from the earth
station. These commands will typically be provided to control the
navigation and control system in the transfer vehicle, but may also
be provided to control components in the payload satellite during
the transfer period. For example, the payload satellite may be
configured to collect data during the transfer period, and the
commands from the earth station may control this data collection
and/or control the reporting of the collected data to the earth
station. Similarly, the status of equipment within the payload
satellite may also be reported to the earth station via the
communications system, to facilitate the resolution of problems
before the satellite is placed in the target orbit.
[0031] Extending this concept further, the navigation and control
systems of the transfer vehicle may be configured to remain with
the payload satellite as well, to provide navigation and attitude
control to the payload satellite while it is deployed in the target
orbit. Because orbit maintenance generally requires substantially
less energy than orbit transfer, providing the fuel required for
the navigation and control of the satellite while in orbit amounts
to a relatively minor addition to the mass and/or complexity of the
transfer vehicle, and eliminates the need to include these
functions in the payload satellite as well.
[0032] In a preferred embodiment of this invention, the
orbit-transfer vehicle is configured to optionally provide these
services to a payload satellite. That is, because the
communications, power regulation, and other functions are required
to effect the maneuvering of the orbit-transfer vehicle, and the
cost of providing an interface to these functions is relatively
slight, particularly compared to the costs of providing multiple
versions of an orbit-transfer vehicle, the preferred embodiment of
the orbit-transfer vehicle includes these functions and the
interfaces that allow these functions to be used by the payload
satellite, regardless of whether the particular payload satellite
is configured to use these functions.
[0033] Note that the concept of providing functionality to the
payload satellite from the orbit transfer vehicle during and after
transit can be embodied in any orbit transfer vehicle, regardless
of whether the satellite is being transferred from a high-energy to
low-energy orbit. Conventional independent low-energy to
high-energy orbit transfer vehicles are configured to merely
provide propulsion and navigation control to place the payload
satellite into the higher-energy orbit. In accordance with this
aspect of the invention, the orbit transfer vehicle includes
components that are designed to provide services to the payload
satellite during the transfer and/or after the satellite is placed
in the target orbit, regardless of the direction of
orbit-transfer.
[0034] Similarly, it will be recognized by one of ordinary skill in
the art that the physical and/or logical partitioning of the
transfer vehicle and the payload satellite is somewhat arbitrary.
For example, the payload satellite may include the aforementioned
kernel system, and this kernel system within the payload satellite
may provide the power and communications functions to the orbit
transfer vehicle. Likewise, the transfer vehicle may be configured
to contain some of the common equipment, such as the power
regulation system, while the payload satellite is configured to
contain other common equipment, such as the communications system.
These and other system configurations and optimizations will be
evident to one of ordinary skill in the art in view of this
disclosure, and are within the spirit and scope of this
invention.
[0035] The foregoing merely illustrates the principles of the
invention. It will thus be appreciated that those skilled in the
art will be able to devise various arrangements which, although not
explicitly described or shown herein, embody the principles of the
invention and are thus within the spirit and scope of the following
claims.
* * * * *