Satellite on-orbit refueling: a cost effectiveness analysis Hibbard, Rustie L. 1996-09

Satellite on-orbit refueling: a cost effectiveness analysis Hibbard, Rustie L. 1996-09
Calhoun: The NPS Institutional Archive
Theses and Dissertations
Thesis Collection
1996-09
Satellite on-orbit refueling: a cost effectiveness analysis
Hibbard, Rustie L.
Monterey, California. Naval Postgraduate School
http://hdl.handle.net/10945/32247
NAVAL POSTGRADUATE SCHOOL
MONTEREY, CALIFORNIA
THESIS
SATELLITE ON-ORBIT REFUELING: A COST
EFFECTIVENESS ANALYSIS
by
Rustie L. Hibbard
September, 1996
Thesis Advisor:
Second Reader:
Dan C. Boger
Bill Clifton
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September, 1996
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Master's Thesis
TITLE AND SUBTITLE Satellite On-Orbit Refueling: A Cost
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PERFORMING
ORGANIZATION
REPORT NUMBER
Effectiveness Analysis
6.
7.
AUTHOR(S) Rustie L. Hibbard
PERFORMING ORGANIZATION NAME(S) AND ADDRESS(ES)
Naval Postgraduate School
Monterey CA 93943-5000
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11. SUPPLEMENTARY NOTES The views expressed in this thesis are those of the author and do not reflect the
official policy or position of the Department of Defense or the U.S. Government.
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Approved for public release; distribution is unlimited.
13. ABSTRACT (maximum 200 words)
With the ever-shrinking military budget constraints facing military and civilian contractors, the
ability to extend the operational life of any system for minimal cost compared to a replacement is
desirable. This fact has never been more true than in today's space industry. This thesis addresses
the possibility of extending satellite life through the use of on-orbit refueling. Through compilation
and analysis of satellite operational life span data, it is shown that maneuvering fuel depletion has a
significant impact on satellite operations in geosynchronous orbit. If these satellites could be
refueled economically this would prove not only cost-effective but also improve satellite tactical
employment for space support to the warfighter. Through the manipulation of satellite data,
launch/design cost, on-orbit refueling vehicle design/construction costs and on-orbit operational
requirements, it can be shown that on-orbit refueling can be done cost effectively. Single versus
multiple satellite refueling operations were evaluated to determine the concept's viability.
15. NUMBER OF
PAGES
14. SUBJECT TERMS Satellite, Refueling, Cost Analysis
100
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Unclassified
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Unclassified
19. SECURITY CLASSIFICATION OF ABSTRACT
Unclassified
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ABSTRACT
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Approved for public release; distribution is unlimited.
SATELLITE ON-ORBIT REFUELING: A COST EFFECTIVENESS
ANALYSIS
Rustie L. Hibbard
Commander, United States Navy
B.S., Jacksonville University, 1979
MASTER OF SCIENCE IN SYSTEMS TECHNOLOGY
(SPACE SYSTEMS OPERATIONS)
from the
NAVAL POSTGRADUATE SCHOOL
September 1996
Author.· _ _ __
Approved by·
iii
iv
ABSTRACT
With the ever-shrinking military budget constraints facing military and civilian
contractors, the ability to extend the operational life of any system for minimal cost
compared to a replacement is desirable. This fact has never been more true than in
today's space industry. This thesis addresses the possibility of extending satellite life
through the use of on-orbit refueling. Through compilation and analysis of satellite
operational life span data, it is shown that maneuvering fuel depletion has a significant
impact on satellite operations in geosynchronous orbit. If these satellites could be
refueled economically this would prove not only cost-effective but also improve satellite
tactical employment for space support to the warfighter.
Through the manipulation of
satellite data, launch/design cost, on-orbit refueling vehicle design/construction costs and
on-orbit operational requirements, it can be shown that on-orbit refueling can be done
cost effectively.
Single versus multiple satellite refueling operations were evaluated to
determine the concept's viability.
v
vi
TABLE OF CONTENTS
I. INTRODUCTION .......................................................................................................... 1
II. SATElLITE ON-ORBIT SERVICIN"G ........................................................................ 5
ill. SATElLITE DATA ANALYSIS .............................................................................. 23
N. OORDESIGN ........................................................................................................... 29
V. OOR COST FEASffiiLITY ASSESSMENT ............................................................. 45
VI. CONCLUSION .......................................................................................................... 61
APPENDIX A. SATElLITE DATA SUMMARY ......................................................... 71
APPENDIX B. DSCS-Ilffi SATELLITE DATA. ............................................................ 77
APPENDIX C. OOR NON-RECURRIN"G COST ESTIMATES ..................................... 81
LIST OF REFERENCES ................................................................................................. 89
IN"ITIAL DISTRffiUTION LIST ..................................................................................... 91
vii
I. INTRODUCTION
A.
WHY SATELLITE REFUELING?
With ever-shrinking budget constraints facing the military and civilian
contractors, the ability to get the most for your dollar has become a major factor in all
programs. The ability to extend the operational life of any system, for minimal cost when
compared to a complete system replacement, would obviously make that system much
more desirable as well as salable. This fact has never been more true than in today's
space industry. Budget constraints have become the driving force, resulting in the
continual search for more efficiency and flexibility from new satellites and systems.
If a satellite design life of five years could be extended to seven years, over a
twenty year period a savings of more than one satellite and its associated launch cost
could be realized. Currently, in the opinion of many, fuel usage drives the operational
restrictions placed on satellites, with fuel being closely managed in order to utilize the
complete satellite design life. A fully functional satellite, which has depleted its
maneuvering/station keeping fuel to reserve levels, is no longer usable. The reserve fuel
must be used to boost the satellite into a super-synchronous orbit in order to vacate the
geo-synchronous slot for a replacement satellite. A second option is to use all fuel for on
station maneuvering and allow the satellite to drift toward the nearest "dead point" which
further complicates the space debris problem. If the satellite actual life exceeds design
life, the operational limiting factor could very well be onboard maneuvering fuel. The
ability to replenish satellite maneuvering fuel on-orbit could result in a significant
satellite operational life extension. With little or no fuel budgeted for contingency
1
operations, the need to maneuver a communications or intelligence satellite to cover
evolving regional conflicts (such as Iraq/Kuwait or Chinaffaiwan) could directly and
disastrously affect the initial design fuel budget. Once a satellite/fuel load is placed on
orbit, tactical repositioning would directly impact satellite life expectancy through the
diversion of fuel budgeted for normal station keeping operations to satellite
maneuvers/repositioning. Fuel considerations/limitations may also preclude development
of operational concepts necessary to meet lower priority tasking/requirements. This
"husbanding" of limited onboard fuel assets for strategic missions could negatively
impact "space support to the war fighter". Again, the ability to replenish satellite
maneuvering fuel on-orbit could result in significant operational flexibility, as well as
increased operational life expectancy.
There have been numerous proposals to develop a "satellite launch on command"
capability to cover regional conflicts such as those previously mentioned. During
operation Desert Storm there were grossly insufficient regional communications
channels/capabilities, with "FLASH" message delivery often taking several days. The
deployment of on-orbit units such as a "duty" communications or intelligence satellite
which could be maneuvered to cover the latest global "hotspot" could compensate for this
shortfall. However, the satellite must be able to maneuver freely without concern for
onboard fuel. This concept would be feasible if a satellite was refuelable "on-orbit".
The Westar 6 communications satellite, launched 4 February, 1984, suffered a
PAM-D upper stage booster malfunction, stranding the satellite in a useless orbit.
The
apogee kick motor and onboard thrusters were used to boost the satellite to a higher orbit
2
for eventual retrieval by STS 51 A. Total rescue cost exceeded $10.5 million.
On 4
April, 1983, the launch ofTDRS 1 experienced a second stage IUS malfunction, which
required the use of 30 onboard thruster burns, consuming 370kg of maneuvering fuel, to
obtain desired orbital positioning. Most recently, a GPS Block III satellite experienced
booster malfunction, which left the satellite in an orbit too low to fulfill mission
requirements. The onboard fuel was sufficient to boost the satellite to the required
operational orbit, however, the maneuver would consume the fuel budget for the entire
satellite life span. The satellite was thus necessarily boosted into super-sychnronous
orbit, fully functional but operationally worthless. Had these satellites been on-orbit
refueling (OOR) capable, a refueling mission after initial satellite altitude repositioning
could have restored the maneuvering fuel reserves and saved much of the cost of the
replacement satellite/ associated launch or retrieval efforts. [Ref 7 ,Comsats]
B.
TACTICAL APPLICATIONS
As airborne refueling revolutionized tactical and strategic aviation, an on-orbit
satellite refueling capability could result in a similar expansion in mission scope and
flexibility in space. The OOR capability would allow operational necessity vice fuel
considerations to drive mission tasking. No longer must each satellite repositioning be
weighed against the tactical or strategic "benefit" which often falls short when
considering a limited maneuvering fuel budget. [Ref l,p.9-14]
An OOR capability could allow for an actual decrease in initial onboard fuel
budget, thus allowing for increased payload/mission capability. Current satellite design
requires an onboard fuel capability sufficient to meet the design life expectancy.
3
However, care must be taken to not oversupply onboard fuel, to preclude the satellite
reaching its end of life with several hundred pounds of now useless on-board fuel. With
launch costs up to $10,000/pound to geosynchronous orbit (about 35,000km.),
elimination of excess onboard weight is critical. Engineers must also consider the fact
that many satellites exceed their scheduled design life, and hence may require additional
on-board fuel if this proves to be the case. Engineers and designers must carefully
balance all these factors and then hope for the best. I can think of nothing more
frustrating than being forced to discard a fully functional satellite due to station-keeping
fuel depletion. Although you may be gambling on a successful refueling mission, if the
initial fuel budget is sufficient to meet design life, should actual satellite life exceed
design life, OOR capability could solve the initial design dilemma. The tradeoffs would
involve the actual weight of the docking/refueling apparatus versus the launch
cost/weight penalty. However, if the weight increase would not dictate a shift to a larger
payload capable launch vehicle, the impact would be minimal.
The scope of this evaluation will be limited primarily to satellites in
geosynchronous orbits (GEO).
Due to the associated system redundancy required for
the safety of manned expeditions and the associated expense, this evaluation will be
limited to unmanned vehicles. Specific refueling vehicle design will not be addressed.
The points which must first be addressed are:
-Is on-orbit refueling (OOR) technologically feasible?
- Is fuel actually a limiting factor in GEO satellite operations?
-Is OOR cost effective?
4
II: SATELLITE ON-ORBIT SERVICING
A.
BACKGROUND
On-orbit satellite servicing is not a new idea. The concept was recently explored
in 1984, when NASA first discussed use of the space shuttle to retrieve, refuel and repair
imaging reconnaissance satellites in order to extend their operational life spans. [Ref 2]
This concept was first successfully demonstrated in April, 1984 during the recovery and
repair of the Solar Maximum satellite. This shuttle mission was the first to use a direct
insertion technique, which resulted in a shuttle apogee of 250nm, necessary to reach the
265nm altitude of Solar Max.[Ref 3, p.42-44] The successful rendezvous with the
satellite allowed astronauts, using extra-vehicular activity (EVA) suits and the shuttle
manipulator arm, to successfully retrieve the 4,500lb satellite (after one initial failure) for
repair in the shuttle bay. Replacing a failed General Electric attitude control box, the
coronagraph's main electronic control box, and installing a vent port baffle to prevent
plasma entry into satellite electronics took the two astronauts approximately six hours.
The repair was made possible due to the Goddard/Fairchild multi-mission spacecraft
modular design employed on Solar Max. [Ref 4, p.18]
NASA's success with Solar Max led to the scheduled on-orbit attempt by shuttle
mission 51-I to repair the $85-million Hughes/Navy Leasat 3 satellite. [Ref 5,p.48]
At
an altitude of 242nm, the 7.5ton Leasat 3 failed to activate after its initial deployment on
April12, 1985 by STS-51D. A previous attempt by the mission 51-D crew to deploy the
manual arming lever, using the shuttle manipulator arm, was unsuccessful. After some
initial difficulty in retrieval, the satellite's sequencer was disabled and the booster motor
5
safety pinned. Two small panels were removed and a spin bypass unit was installed to
allow Leasat 3 to process coirimands directly from the ground. After connecting a battery
powered control box, the satellite's 7 .5ft omni antenna was deployed, which concluded
the initial EVA at 7.62 hours. The second EVA of 2.45 hours consisted of the installation
of temperature sensors on the motor nozzles, removal of previously installed safety pins
and the activation of two 13 hour timers which precluded the processing of ground
signals for 13 hours, in order to allow for safe withdrawal by the shuttle prior to satellite
activation. [Ref 6, p.21-23] NASA received $8.5 million for conducting the successful
repair effort. Compared to the initial satellite cost of $85 million (plus associated launch
costs) coupled with a replacement satellite/launch costs, the repair was truly a bargain.
[Ref 7, Comsats]
The most recent and probably most famous instance of on-orbit servicing was
conducted by STS-61, to repair the $1.5 billion Hubble Space Telescope (HST). After
initial launch in 1990, scientist discovered the HST had several problems, the most
significant being the inability to focus (due to improperly ground optics) as well as a
"jitter" problem related to the solar arrays. The very rapid temperature change during
day/night transitions resulted in array deflections, which although extremely minute,
directly impacted HST operations. The original arrays were replaced with a shielded, 9
coil spring mounting array, with an onboard braking control to eliminate solar induced
array movement. Servicing also included: the installation of corrective optics space
telescope axial replacement (Costar) to correct HST's vision flaws, swapping a secondgeneration wide field camera, replacement of a failed relay box in the Goddard High
6
Resolution Spectrograph (GHRS), installation of a coprocessor module to add computer
memory, replacement of three failed gyroscopic units, and change out of the
magnetometer. [Ref 8, p.28-29] Although the HST was designed with multiple
replaceable parts for periodic on orbit repair, many of the scheduled repair operations
involved units or access panels which were not designed for on-orbit servicing. [Ref 9,
p.14-16] Servicing efforts proved a resounding success, further justifying the on-orbit
servicing satellite design concept.
The concept of on-orbit refueling was successfully demonstrated on shuttle
mission 41-G, by CDR David Leestma and Kathryn Sullivan. This proof of concept,
using the Orbital Refueling System (ORS), demonstrated the capability to refuel satellites
currently on-orbit which have not been specifically modified for refueling operations.
This process involved special penetration of the fueling system. [Ref9, p.15]
Although the use of shuttle manned EVA evolutions to conduct on-orbit servicing
has proven successful in LEO, shuttle operational limitations preclude such operations
above 400nm. [Ref 10, p. 179] Satellites which operate in :MEO or GEO with typical
altitudes of as high as 22,000nm are not accessible to shuttle flights at this time.
However, as successful as NASA has been in conducting on-orbit satellite repairs, the
presence of manned evolutions significantly increases the cost.
However, modular
replacement or refueling operations using unmanned vehicles requires the satellite to be
designed with this eventuality in mind. Several on-orbit service vehicle (OSV) design
options have been evaluated, with the most significant being the Orbital Maneuvering
7
Vehicle (OMV), designed for NASA by the TRW Space and Support Group. Its primary
missions include:
- Spacecraft retrieval, reboost, deboost or viewing
- Spacecraft on-orbit servicing, including refueling and component
replacement
- Space station construction and logistics support
- Large observatory service (HST) from either space station or shuttle
-Experiment carrier for sub-satellite missions [Ref 1, p.32-33]
NASA plans call for the OMV to be deployed via the space shuttle and later
retrieved for return to earth for periodic servicing. The OMV is 15 feet in diameter and is
56 inches in length (see Figures 1, 2 and 3). It incorporates a fully modular
configuration which allows on-orbit replenishment of fuels as well as replacement of
modular units (ORUs). The OMV was designed to service satellites in LEO, polar orbits
(inclinations above 57 degrees) which are not accessible by shuttle operations. [Ref
11,p.29-33]
Although intended for use in LEO operations, the OMV unmanned servicing
vehicle concept can be applied to satellite refueling operations in GEO.
automation maneuvers must be precise and assured.
However,
The two major limiting factors of
on-orbit refueling are satellite rendezvous/docking and fuel (fluid) transfers.
8
HIGH GAIN ANTENNA AI
LOW GAIN ANTENNA A
PAN-TILT-ZOOM
ORU1A
POWER
DISTRIBUTION
ORU8
DATA
HANDLING
REMOTE GRAPPLE
DOCKING MECHANISM
>
ORU 18
POWER DISTRIBUTION
ORU4B
ATTITUDE
CONTROL
NAVIGATION
HiGH GAIN ANTENNA B/
LOW GAIN ANTENNA 8
COMMUNICATIONS
Figure 1. OMV Service side for ORU replacement.
9
+Z
t
RCS AV
THRUSTERS
MANIFOLDED
HYDRAZINE
TANKS 141
SOLAR ARRAY
14 PANELS)
0
+X
DOCK INO
MECHANISM
THERMAL COVER
' - - - - - - - - . . . - - - - - - - - ' P R O P U L S I O N MODULE
SHORT RANGE VEHICLE
Figure 2. OMV propulsion side - showing replaceable propulsion module
10
Figure 3. OMV shown docked with satellite.
A semi-autonomous navigation and docking capability has been developed.
Using optical reference patterns and a computer vision system to determine relative
position and attitude, semi-autonomous docking has been evaluated using both
11
L___ --------------~--
"passive"and "active" targets. Active targets have visual reference "cuing markers"
installed prior to launch, while passive targets do not. Only the primary spacecraft is
under active control, with the docking target completely independent. Cuing markers
consist of geometric patterns which provide orientation and/or distance reference using
the geometric patterns (see Figure 4). Nesting a series of these optical patterns provides
a means by which an autonomous cross-correlator guided craft can determine its range
and orientation during approach and docking maneuvers (see Figure 5). At the furthest
distance the largest pattern is · used as a reference. Upon closer approach, the correlation
pattern grows larger and larger in the field of view until it actually reaches a range where
the complete target is no longer visible. At this point a smaller nested array, located at
the center of the first pattern is discernable, and the system begins to process the second
pattern for range and orientation data. The simple task of recognizing a single visual
pattern, in a cluttered environment is well within the capability of an optical cross
correlator. This single-function vision device can accurately provide the necessary
recognition and spatial orientation necessary for semi-autonomous navigation, landing
and docking in three dimensional space, without natural landmarks. The singlefunctional optical cross- correlator, using video input from a simple imaging camera and
optical correlation-plane output coupled with standard star tracker software, provides
sufficient information for spacecraft navigation and docking maneuvers in space. [Ref 11,
p.5049-55]
12
The insert shows the four-peak pattern in
the simulated correlation plane. Example ¢
shown of attitude determination.
:... :.~.;:
:_,:··
..
·.
·... ...
..
....
. -. ·.
The out-of-plane rotation of the
..
.
. .
.
.
.
'
.
.
¢:1
.
.
array of sector-star targets results
in the foreshortened triangles in
the simulated correlation.
The four-star target pattern may
be used for range determination. ¢
The star-target distribution shown
is closer than the example shown
at the top of the page.
Figure 4. Four-sector star target visual patterns used for attitude determination.
13
Figure 5. Simulated docking port with three sector-star targets attached.
The following conditions are required:
- The spacecraft must be a rigid body with three pairs of gas jet thrusters
mounted along the principle axes for control of spacecraft translational and
rotational motion.
- A pinhole camera is rigidly mounted on the spacecraft. The cuing marks
located on the docking target platform are always visible to the spacecraft.
14
Computer-vision based methods have several advantages over the use of sensors
such as laser, infrared, radar, GPS or INS. Specifically, the estimation accuracy of the
relative position and vehicle orientation improves as the range between the two vehicles
decreases. Thus, the control accuracy of the system control loops improves
proportionally (see Figure 6 and 7). Hence, the computer-vision based control and
docking system is well suited for precise maneuvers required for autonomous satellite
docking. [Ref 12, p.649]
Currently, Russia conducts resupply missions to the MIR space station using the
Progress spacecraft, which employs the Kurs automatic rendezvous and docking system.
A back-up remote control docking capability has been developed, which although not
autonomous, does not require manned participation on-orbit. The TV-aided system
enables ground-based controllers to remotely fly the spacecraft for rendezvous and
docking.
A television camera provides live images to the ground-based cosmonaut,
who will dock the spacecraft using two control sticks, much as if he were actually
onboard. A successful demonstration of this system was conducted in 1993 by a
cosmonaut onboard the MIR space station. [Ref 13,p.70] Hence, docking unmanned
refueling missions should not pose a technical problem.
There are three major methods of on-orbit propellant transfer: direct fluid transfer,
tank to vehicle transfer, and propulsion module to vehicle transfer. Direct fluid transfer,
as implied by the name, involves the transfer of fuel from the servicing vehicle directly to
the satellite tank. Tank to vehicle transfer involves the transfer of full fuel tanks to the
satellite as orbital replacement units (ORU). Propulsion module transfer involves the
15
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Figure 6. Graphs show correlation to controlled X, Y, and Z axis motion and associated
error as range decreases.
16
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17
installation of a complete propulsion system to include the propellant tanks, main
engines, thrusters, fluid lines and management systems and controls. Trade studies and
evaluations revealed that the direct fuel transfer was the most cost effective and feasible
technique, and has the least impact on spacecraft design. Additionally, this method
allows full utilization of onboard fuel, while requiring the smallest number of interfaces.
Disadvantages involve safety concerns related to actual displacement of propellant from
one tank to another. The primary focus of this study involved fueling in space for long
duration missions, such as manned Mars exploration. Hence, concerns of lengthy fuel
transfer durations and the associated large pumping capability required would not apply
to satellite refueling, due to the much smaller relative fuel quantities required for satellite
stationkeeping/maneuvering operations. [Ref 14, p.1423-33]
The transfer of fuel is complicated by many factors, the most significant involving
a means of pumping fuel in a near weightless environment and the necessity to vent waste
gases from the receiving tank as it fills, without venting fluids. Fuel cannot be gravity fed
for obvious reason. The most promising on-orbit servicing method for direct fuel transfer
under these conditions involves the use of a screen-channel liquid acquisition device
(LAD). Designs for screen LADs are usually conduits, with walls made of porous, fine
mesh screen, which are routed around the tank perimeter and manifold at the tank outlet.
[Ref 15, p.1099-1106]
The capability for fuel/fluid transfer on-orbit was successfully demonstrated on
shuttle mission 41-G during astronaut EVAs. However, EVAs are an expensive option
and do not meet stated goals of autonomous operations. NASA conducted Fluid
18
Acquisition and Resupply Experiments (FARE) in shuttle middeck experiments to
demonstrate LAD techniques for transferring liquids in zero gravity (see Figure 8, 9 and
10). The first experiment occurred on STS 53, launched December 2, 1992, with the
second on STS 57 in June of 1993. The objective was to demonstrate tank refilling, low
gravity propellant center of gravity control, and expulsion efficiency. A fluid expulsion
UpptJr Mor· .;le-,;=~'-A---11
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Figure 8. FARE test configuration in shuttle bay.
19
Ad'''" '"'"
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Figure 9. FARE 1 receiver tank configuration.
20
Evacu<1ted !"ill- Tank Nearing Comrlcie rill
Evacu<Jted Fill- Initial Fillino
Cloudy
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Cloudy Liquid/Bubble
Mixture Which
Cleared As Tank
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Initial Phase of Tank Expulsion
Following Evacuated Fill
Continued Tank Expulsion - Bubbles Began to Fill The
Tank As Gas-Free Expuls1on Con:,nuec
Expulsion Nearing Completion
~
Expuls,on to Gas lngesr,on
Bubbles Fill
Most of ll1e TJnk
Rem<~inina L1qu1d
is Collected Around
Screen Channels
Gz:;)
~~·.::·.·.een
::.ri~~~ ..-.~!
Figure I0. Evacuated Fill/Expulsion Test Results.
21
c:nd
!he
T311~
efficiency of at least 98% was obtained and final fill levels of greater than 95% were
routinely achieved without the venting of liquid overboard, thus validating capabilities to
refuel spacecraft on-orbit. [Ref 16, p.1-12]
Although the problems presented by autonomous docking and on-orbit fluid
transfer are technoligically challenging, they remain well within the range of current
mechanical and scientific capabilities. However, production and integration costs may
dictate actual system applications in order to assure maximum cost effectiveness. The
most obvious question at this juncture remains, does satellite maneuvering fuel actually
impact satellite life span and operations to the point that on-orbit refueling is necessary?
22
III. SATELLITE DATA ANALYSIS
A.
SATELLITE LIFE-SPAN ANALYSIS
Prior to conducting a cost effectiveness analysis of on-orbit satellite refueling, the
data concerning a satellite's actual life and its design life must be examined. Without
addressing the tactical maneuvering of satellites and associated fuel considerations,
satellite life expectancy is an important factor in determining if the need to refuel exists.
The satellite data listed in Appendix A- Part 1 (page 1 and 2), was compiled from Jane's
Spacecraft 1984-1996, lnteravia Space Directory 1986-1996, as well as inputs from
various contractors such as TRW, Hughes, etc. The data set consists of U.S. satellites,
launched to geosynchronous orbit, in the last 20 years. Analysis of satellites which have
reached geosynchronous orbit and full operational status (i.e., they have not experienced
launch-related failures) follows:
Sample Size
57 Satellites
Mean Design Life
8.11 years
StdDev
Mean Actual Life
StdDev
1.94 years
11.41 years
3.08 years
Table 3-1
With the satellite actual life exceeding design life by an average of 3.3 years, this
data supports the hypothesis that satellites typically exceed design life expectations.
However, these statistics are somewhat misleading, as 28 of the 57 subject satellites are
still operational (SOPER). This computation assumes each satellite, even if still
operational, has reached end-of-life (EOL).
(GOES 5/6 experienced imagery
failure (the primary mission), but continued service in a data relay mode until fuel
depletion. The primary mission role was used in assessing mean satellite life.)
23
Evaluating only the satellites which have actually reached EOL shows:
Sample Size
Mean Design Life
7.66 years
29 Satellites
StdDev
Mean Actual Life
StdDev
2.09 years
10.43 years
3.56 years
Table 3-2
With the satellite actual life exceeding design life by an average of 2. 77 years or 26.55
percent, this data is .6 years less than the total sample average life delta of 3.3 years, but
within the standard deviation.
Evaluating only the SOPER satellites yields:
Sample Size
Mean Design Life
28 Satellites
8.46 years
StdDev
Mean Actual Life
StdDev
1.69 years
12.43 years
2.10years
Table 3-3
Examination of the SOPER satellite subset reveals that satellite actual life
currently exceeds design life by 3.97 years or 31.93 percent.
The SOPER satellite subset
life delta is greater than both the entire satellite sample delta (3.08 years) and the EOL
satellite delta (2.77 years). As these satellites are still operational, each passing year will
increase the life delta until all satellites have reached EOL. Even without considering
these additional active years, there is significant evidence that satellite actual life
consistently exceeds satellite design life, regardless of sample data chosen.
However,
for the purpose of this cost effectiveness study, the average satellite life delta of three
years will be used.
24
B.
SATELLITE FUEL ANALYSIS
The next question which must be addressed is, did fuel play a significant role in
satellite failures? Fuel depletion is an obvious fuel impact, but other factors must be
considered. Satellites which are still operational, but are nearing maneuvering fuel
limits/depletion, often continue East/West stationkeeping but cease North/South station
keeping in order to conserve maneuvering fuel. This practice results in geosynchronous
satellites assuming inclined orbits, which impacts the satellite's area of coverage or
"footprint" on the earth. This will impact coverage in peripheral areas at the northern and
southern extremes of coverage. As the inclination increases (about 1 degree/year without
correction) the affected area increases as well. Satellites conducting fuel conservation
operations (FCO) are thus impacted by fuel limitations.
Examining the cause of failure for the 29 satellites which have reached EOL
reveals that nine satellites (31 percent) ceased operation due to maneuvering fuel
depletion. This number increases to 11 satellites (37. 93 percent) if the fuel depletion of
GOES 5/6 is considered. Examining the SOPER satellites reveals that 19 of 28 (67.85
percent) satellites are currently conducting fuel conservation operations. Comparing the
EOL and SOPER data:
Fuel hnpact
X Design Life
X Actual Life
Life Delta~
~%
EOL
7.66 years
10.43 years
2.77 years
26.55%
37.93%
SOPER
8.46 years
12.43 years
3.97 years
31.21%
67.85%
Table 3-4
25
Even without the inevitable increase in the SOPER satellite life delta, with current
satellite design life ranging from 10 to 15 years (see Appendix A, p. 3-5), fuel
considerations in the future will continue to significantly impact satellite operations.
A projection for the satellites launched since 1990 shows:
X Design Life
X Actual Life
Life Delta 1::.
1::.%
Assumption
Since 1990
11.37 years
14.89 years
3.52 years
31.21%
No Increase
Since 1990
11.37 years
15.45 years
4.09 years
35.87%
Projected
Table 3-5
Both "no increase" and "projected increase" in satellite design life versus actual life
options are shown, with the projected increase based on the EOUSOPER satellite
design/observed life data. Although this projection is rather crude, even using the current
satellite life delta of 31.21 percent, this data indicates that fuel considerations are
becoming increasingly more significant.
Combining the entire satellite sample and associated fuel considerations, 30 of 57
(52.63%) satellites experienced some fuel-related operational impacts, with 20 percent
failing due to fuel depletion. A convincing argument can be made that fuel limitations
have a significant impact on satellite operations and that an on-orbit refueling capability
could play a major role in solving this problem.
There are many alternative solutions to the satellite fuel problem other than onorbit refueling. Many satellites, such as INTELSAT 706, are carrying additional fuel to
preclude a fuel depletion problem. However, there are satellites, such as GALAXY 517,
which do not carry sufficient fuel to meet expected design life.
26
UFO satellites, which
previously experienced a fuel surplus, have recently seen this benefit eliminated due to a
payload-for-fuel substitution. Additional fuel for UFO is not an option as the satellite is
currently within 50 pounds of maxing out the launch vehicle payload capability. [Ref 17]
As satellite design life continues to increase, there must be a point where it becomes
economically and physically impossible to provide sufficient onboard fuel to meet design
life. However, if tactical maneuvering of satellites is considered a viable mission
requirement, on-orbit refueling is the only logical solution. The next question is, can it be
done cost effectively?
27
28
IV: OOR DESIGN
A.
OOR SPECIFICATIONS
As this is primarily a cost effectiveness analysis of the OOR concept, the specific
design of the OOR will not be addressed. However, some general concepts and design
configurations must be identified. Simplicity of design should be incorporated whenever
feasible, utilizing as much existing proven space technology as possible. The OOR must
have a configuration which would support launch on Titan N (IUS) or comparable
launch vehicle. Using the large payload fairing limits of the Titan N, the OOR can be a
maximum width of 5.1 meters and height of 15-26 meters in the stowed configuration.
Maximum payload launch weight is 5250lbs. OOR design will be limited to 3562lbs
(2762lbs dry weight/800lbs of fuel), which will allow approximately 1700lbs of design
weight error margin when considering the 5200lb Titan N (IUS) launch capacity.
The Fuel Transfer System (FTS) design should assume the use of mono-propellant
(the primary fuel used in geo-synchronous satellites). With 800lbs of total fuel onboard,
the OOR should also use mono-propellant to preclude the necessity of two separate fuel
systems. With an anticipated total fuel load of 800lbs of mono-propellant, the transfer
system is envisioned with the ability to feed both its own thrusters and the refueling
system from any fuel tank/cell. This cross-feed design feature would preclude the OOR
from depleting maneuvering fuel with transferable fuel still onboard and vice-versa. This
would allow maximum flexibility and utilization of all onboard fuel.
However, to
preclude a compromise in fuel system integrity from depleting the entire onboard fuel
supply, each tank should be selectively isolated from the others.
29
Primary transfer is
envisioned through the docking mechanism; however, a secondary transfer system
should be available. This secondary system might consist of remote tele-robotics using
an umbilical fuel probe in the event that satellite docking is not feasible or the primary
system fails.
Docking and refueling operations may preclude optimal orientation of solar
arrays. Onboard battery power should be sufficient to complete a refueling operation
without solar array support. Upon completion of the refueling mission, the OOR can then
be reoriented to recharge batteries.
With a payload of 800lbs of fuel, in addition to the required fuel transfer system
of 300lbs, the structure of the OOR must be robust. Additionally, sufficient structural
integrity is required to support docking maneuvers and its associated structural stresses.
The command and control communications required by a remote/autonomous
docking system are considered in addition to normal TT&C operations. Sufficient
redundancy is required to ensure communications can be maintained throughout mission
life.
The OOR concept is obviously not applicable to satellites currently on-orbit, and
hence it must be designed with the "future customer" in mind. If the concept is to prove
viable, satellites must be designed to accept fuel servicing from the OOR from initial
inception. This dictates the early, standardized design of a docking and fuel transfer
mechanism and specifications which will meet the needs of most if not all satellite
designs. Clearance and configuration requirements for OOR docking must be identified
early. From the docking port into the spacecraft would be the design responsibility of
30
Inter Sat. Com. Antenna
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each individual satellite manufacturer, thus minimizing the concept's impact on
individual satellite design. It is anticipated that through early design and integration, the
docking/refueling mechanism should impose minimum impact on satellite weight
budget/cost. However, later integration could prove extremely costly, depending on the
design progress of the satellite. With these factors in mind a cost analysis to construct
the OOR follows.
B.
OOR DESIGN/COST PARAMETERS
In 1997 NASA is scheduled to fly ETS-7 to verify remote on-orbit docking and
robotic repair/servicing technology. ETS-7 consists of a chaser and target satellite
(Figure 11), which weigh 2.2t and 0.4t respectively. [Ref 18, p.1627] Using the
approximate size/configuration ofETS-7 and DSCS-lliB satellite costing data, a cost
estimate for an on-orbit refueling vehicle was computed. Adjustments in size/
configuration for the OOR were made using the Unmanned Space System Cost Model,
Seventh Edition [Ref 20] and verified using the Space Mission Analysis and Design,
Second Edition [Ref 21] for an overall system "reality check" of the design process.
Since specific system data and configuration is not available for ETS-7, DSCS-lliB was
chosen as a cost/equipment reference satellite due to its similar configuration to ETS-7,
its geosynchronous orbit, recent technologynaunch date (1994) and availability of
configuration data (Appendix B) from the Unmanned Space Vehicle Cost Model.
Adjustments to the OOR design from DSCS lliB and ETS-7 are shown in Table 4-1,
32
along with the recommended percentage of system design parameters from SMAD.
SMAD percentages are based on a historical satellite program data base.
Percentage
SMAD%
750lbs
27.50%
21.06%
5.42%
165lbs
6.05%
4.45%
33.56%
307lbs
12.9%
See
DSCS-IllB
Percentage
Structure
330 lbs
17.52%
Thermal
102lbs
Communication
632lbs
OOR
Payload
(DSCS Payload)
Below
TT&C
72lbs
3.82%
75lbs
2.75%
4.41%
ADCS
64lbs
3.39%
380 lbs
11.00%
5.99%
Power
585lbs
31.06%
585lbs
21.45%
29.90%
Propulsion
98lbs
5.20%
200lbs
07.30%
4.31%
N/A
N/A
300 lbs
11.00%
28%
Refueling Kit
(complies
(OOR Payload)
w/fuel
added)
Dry Weight
1883lbs
2762
lbs
Table4-1
NOTE: -Communications weight was adjusted downward from DSCS-IllB, as
communications is the DSCS primary mission.
- Propulsion weight was adjusted upward from DSCS- IllB, as docking
maneuvers will require a more precise system/propulsion between satellite refuels.
- Structure weight was adjusted upward, necessary to support the added weight of
fuel to the refueling system.
33
- ADCS was adjusted upward to include three axis stabilization, using momentum
wheels.
C.
CALCULATIONS FOR RECURRING COSTS
[All Equations from Unmanned Space Vehicle Cost Model, Seventh Edition
(1994) Ref 20]
The Unmanned Space Vehicle Cost Model (USCM) is published by the U.S. Air
Force Material Command, Space and Missile Systems Center, Los Angeles, California.
This manual, which was used to estimate the OOR cost, is a parametric estimating tool
based on cost estimating relationships (CER) built from a factual historical database.
Satellite systems have been broken down into costing factors/equations which can be
used for cost estimates for future satellite systems. The USCM breaks satellite program
costs into six basic satellite subsystems: Power, Structure, Attitude Determination and
Control (ADCS), Tracking, Telemetry and Control (TT&C), Propulsion, and Thermal
Control. For each of these subsystems, the USCM shows the primary costing factors
associated with it based on a satellite design/cost historical data base. Additionally, the
payload must be addressed for cost analysis. On DSCS-IIIB, the communications system
was the payload; however, on the OOR the refueling package would be considered the
payload. The cost analysis for the OOR is computed below. Specific cost multiples or
cost estimating relationships (CERs) have been derived from previous satellite programs.
These CERs can be applied to a specific example to estimate system cost, typically using
system or component weight.
All cost values computed are in thousands of dollars and
recurring cost estimates are summarized in Table 4-2.
34
1. STRUCTURE
750 lbs
Spacecraft Structure
Y=(5.838)(Xl)
Where
Xl= Structure Weight
Y= CER value for Spacecraft Structure
Therefore
Y= 4378.5
2.THERMAL
Thermal Weight
-Active Thermal Weight
- Passive Thennal Weight
- Total Thennal Weight
Y=76.171 + (12.187)(Xl) + (4.511)(X2)
Where
7llbs
94lbs
165lbs
Xl= Active Weight
X2= Passive Weight
Y= CER value for Spacecraft Structure
Therefore
Y= 1365.48
3. ADCS
ADCS
-Attitude Detennination Suite Weight
- RCS Suite Weight
- Total ADCS Weight
Y=(250.542)(Xl 0·735 )
Where
180 lbs
200 lbs
380 lbs
Xl =Attitude Determination Suite Weight
Y = CER value for ADCS (Attitude Determination)
Therefore
Y = 11,389.53
35
y =(27 .667)(XI o.619)(X2o.473)
XI= Reaction Control System Suite Weight
Where
X2 = Design Life (I 0/2 Yrs)
Y = CER value for ADCS(Reaction Control)
Therefore
Y = 2184.24 - 10 yrs/1020 - 2 years
4. ELECTRICAL POWER SYSTEM
EPS
- Number of Solar Cells
- Generation Suite Weight
3000
32
Y=(7 .894)(X I 0·588)
XI= Number of Solar Cells
Where
Y = CER value for Power Generation
y = 874.67
Therefore
EPS
- Beginning of Life Power
- Storage Suite Weil!ht
1200 Watts
135lbs
Y=(2.722)(XI 0·848)
XI= Beginning of Life Power
Where
Y = CER value for Power Storage
y = 1111.82
Therefore
585lbs
EPS Suite Weight
Y=(58.775)(XI
Where
0 713
· )
XI= PCD Suite Weight
Y = CER value for Power Conditioning and Distribution
36
Therefore
y =5523.0
5. TELE:METRY, TRACKING AND CONTROL
IT&C Transmitter (2)
-UHF
-SHF
4lbs
6lbs
Y=(76.928) + (20.435)(Xl)
Xl= Transmitter Weight
Where
Y = CER value for TT&C Transmitter
Therefore
y = 158.67 (UHF)
Y = 199.54 (SHF)
IT&C Receiver/Exciter
9lbs
Y=(47.359)(X lu )(X2°.420)
05
Xl= Receive/Exciter Suite Weight
Where
X2= Number of Receiver Boxes
Y = CER value for TT&C Receiver/Exciter
Therefore
y =718.25
IT&C Transponder (2)
18lbs
Y=(377.529)(Xl 0·281 )
Xl= Transponder Weight
Where
Y = CER value for Power Storage
Therefore
Y = 850.52
IT&C Digital Electronics
- Suite Weight
- Number of Digital Elect Boxes
- Number of Links
Y=(23.406)(Xlo.922)(X2o.6s9)(X3!.09I)
37
23lbs
5
2
XI= Digital Electronics Suite Weight
Where
X2= Number of Digital Electronic Boxes
X3= Number of Links
Y = CER value for TT&C Digital Electronics
Therefore
Y=2593.49
IT&C Analog Electronics
- Suite Weight.
- Solenoid Driver (4)
- Squib Driver (4)
1.2lbs
2.5lbs
2.5lbs
Y=(II3.777)(XI 0·519)
XI= Analog Electronics Weight
Where
Y = CER value for TT&C Analog Electronic
Solenoid Driver (qty) YI= Y(qty0·926 )
Y = (13.777)(X2°519)(Qty926)
Squib Driver (qty)
Where X2= Squib Driver Weight
Solenoid= 451.5
Squib= 660.83
IT&C Antenna(Hom & Radiator)
- Hom & Radiator
-Gain
- Wavelength
- Effective Area
4lbs
.3 db/10
.5ft
.5 sqft
Y=(II9.35I)(XI 0 ·708)(X2°·240)
XI= Antenna Weight
Where
X2= Effective Area
Y = CER value for TT&C Antenna (Hom & Radiator)
Therefore
y =269.67
38
TI'&C Antenna (Dipoles)
Y=(26.609)(XlL070)
Where
1lb
Xl= Antenna Weight
Y = CER value for TT&C Antenna (Dipoles)
Therefore
y =26.61
TI'&C Antenna (S-Band)
.67lbs
.26 db/10
.5ft
.45s
- S-Band Weight
-Gain
- Wavelength
- E ective Area
Y=(64.560)(XlL009)(X2°·315)
Where
Xl= Antenna Weight
X2= Effective Area
Y = CER value for TT&C Antenna (S-Band)
Therefore
Y=33.51
TI'&C RF Distribution
Y=(-7.386) + (29.180)(Xl) + (70.676)(X2)
2.4lb
Xl= RF Distribution Weight
Where
X2= Active (1 =Yes, 0 =No)
Y = CER value for TT&C RF Distribution
Therefore
y = 133.32
6. COMMUNICATIONS
Communications Transmitter (TWTA)
- TWTA Weight
- Output Power
-Frequency
39
14.6lbs
25 Watts
2.1 Ghz
23
-WPF
Y=(22.196)(X 1°·727 )(X2°· 28~
X1=TWTA Weight
Where
X2= WPF- Weighted Composite Variable
Y = CER value for Communications Transmitter (TWTA)
Therefore
y
=375.0
Communications Transmitter (Solid State)
- Solid State Transmitter Weight
- Output Power
- Component Quantity
Y=(338.550) + (25.557)(X1) + (9.985)(X2)
51.26lbs
40 Watts
1
X1= Solid State Transmitter Weight
Where
X2= Output Power
Y = CER value for Communications Transmitter (Solid
State)
Therefore
y =2048.0
Communications Receiver/Exciter Weight (2)
Y=( 193 .30)(X 1°·675 )
Where
30lbs
X1= Receiver/Exciter Suite Weight
Y = CER value for Communications Receiver/Exciter
Therefore
y = 1920.0
Communications Transponder Weight (2)
Y=(67.433)(X1)
Where
X1= Transponder Weight
40
30lbs
·Y =CER value for Communications Transponder
Therefore
Y = 2023.0
Communications Digital Electronics Weight
Y=(515 .079)(X 1°·379)
56.96lbs
X1= Digital Electronics Suite Weight
Where
Y = CER value for Communications Digital Electronics
Therefore
y =2383.61
Communications
- Weight of Other Antenna Components
- Weight of Hom, Dish
-Antenna Suite Weight
Y=(35.473)(X1) + (24.835)(X2)
JOlbs
5lbs
141lbs
X1= Weight of Other Antenna Components
Where
X2= Weight of Hom, UHF Dish
Y = CER value for Communications Antenna
Therefore
y = 727.26
Communications Antenna (Reflectors)
-Antenna Reflector Diameter Squared
Y=(7 5.849)(X 1°·935 )
Where
8 sgfl_
X1= Antenna Reflector Diameter Squared
Y =CER value for Communications Antenna Reflectors
Therefore
Y=530.07
Communications RF Distribution
- RF Distribution Suite Active Weight
- RF Distribution Suite Wave Guide Weil[ht
Y=(82.601)(X1) + (11.856)(X2)
41
6lbs
6 lbs
XI= Distribution Suite Active Weight
Where
X2= RF Distribution Suite Wave Guide Weight
Y = CER value for Communications RF Distribution
Therefore
Y =566.74
7. FUEL TRANSFER SYSTEM (DRY WT)
300 lbs
FTS Total Weight
Y=3(Xt22 }
X= FTS Total Weight
Where
Y = CER value for FTS
Therefore
Y = 3156.52
8. INTEGRATION ASSEMBLY AND TEST (IA&T)
IA&T
- Spacecraft Weight
- Fuel Transfer System Total Weight (Payload)
- Weigftt
Y=(4.833)(XI)
2462lbs
300 lbs
2762lbs
XI= Spacecraft Weight+ Payload Total Weight
Where
Y = CER value for IA&T
Therefore
y
=13,348.75
9. PROGRAM LEVEL
Spacecraft Vehicle Total Recurring Cost
Y=(.289)(XI)
Where
67197.69
XI= Spacecraft Total Recurring Cost
Y =CER value for Program Level
42
Therefore
y
=19420.13
10. LOOS- (3 AXIS STABILIZED SATELLITES)
2762lbs
Spacecraft Weight
Y=(2.212)(X1)
X1= Spacecraft Weight+ Payload Total Weight
Where
Y = CER value for Operations and Orbital Support
Therefore
Y=6695.72
SMAD cost modeling calls for a cost adjustment (multiplier) greater than 1.1, for
satellite designs employing new technology or advanced development concepts. This
cost multiple deals with the uncertainty of new technology and the associated integration
issues. For the purpose of this cost analysis a cost adjustment multiple of 1.3 will be
used due to the incorporation of remote docking using visual references and the fuel
transfer system in the OOR. This yields a final cost estimate of:
($86617) x (1.3) = $112.602 rounded to 113 Million
This figure shall be used as OOR recurring costs for the cost effectiveness
calculations in Chapter V. Specific non-recurring cost data for future satellite programs
is not known. Although difficult to accurately estimate, an attempt to predict OOR nonrecurring costs is summarized in Appendix C. However, it will be assumed that many of
these cost would be offset by similar costs in repalcement satellite programs.
43
RECURRING COST SUMMARY
(in Thousands of Dollars)
Structure
4378.50
Thermal
1365.48
11389.53
Attitude Determination & Control
I
I
ADCS - Attitude Determination (RCS)
2184.24
Electrical Power System - Generation
EPS - Storage
EPS - PCD
874.67
1111.82
5523.0
Telemetry, Tracking & Conunand
TT&C - Transmitter (UHF)
TT&C - Transmitter (SHF)
TT&C - Receiver/Exciter
TT&C - Transponder
TT&C - Digital Electronics
TT&C - Analog Electronics (Solenoid)
TT&C - Analog Electronics (Squib)
TT&C - Antenna Horn & Radiator
TT&C - Antenna Dipoles
TT&C - S-Band Antennas
TT&C - RF Distribution
158.67
199.54
718.12
850.52
2593.49
451.50
660.83
269.67
26.61
33.51
133.32
Conununications
Conun - Transmitter (TWTA)
Conun - Solid State
Conun - Receiver/Exciter
Conun - Transponder
Conun - Digital Electronics
Conun - Antenna
Conun - Antenna Reflectors
Conun - RF Distribution
375.00
2048.00
1920.00
2023.00
2383.61
727.26
530.07
566.74
I
I
LOOS - 3 Axis Stabilized
Fuel Transfer SJ::Stem
6695.72
3156.52
53848.94
Total Spacecraft
13348.75
T
Program Level Cost
19420.13
Total OOR Recurring Cost
86617.82
Table 4-2
44
I
I
V: OOR COST FEASIBILITY ASSESSMENT
A.
COST EFFECTIVENESS ANALYSIS
When evaluating the cost effectiveness of on-orbit refueling, it is tempting to use
the cost/year approach which compares the replacement cost of a satellite and its design
life to the OOR cost and its capability to refuel some specific number of satellites. This
approach would yield:
Cost
Life or Life Delta
CostlYear of Satellite Life
Replacement Satellite
$98.85M
11.27 yrs
$8.77M
OOR Vehicle
$113M
i Satellite(s) serviced
(varies 1-N) with a
life delta of 3yrs
N= 1
$37.67M
N=2
$18.83M
N=3
$12.55M
N=4
$9.42M
N=5
$7.53M
Table 5-l
This comparison results in an apparent break-even point for refueling/replacement at 5
satellites serviced versus one replacement. However, this comparison does not take into
account launch cost of either the replacement satellite or the OOR (the OOR is the same
approximate weight class as the average satellite launched since 1990).
When launch
costs are considered the results appear slightly different. Specifically, with each satellite
45
refueled on-orbit the associated launch cost of the necessary replacement and (hence, notlaunched) satellite is also saved. With these considerations in mind, we can use the
following equation:
Where:
Sc = Replacement Satellite Cost
Rc= OORCost
Lc = Satellite Launch Cost
Lc = OOR Launch Cost
i = Number of Satellites Refueled/
PL.a =%increase in satellite Life
Mission
The cost effective or break-even point of OOR can be determined when the value of the
left side of this equation exceeds the value of the right
The statistical data to be used for this analysis is as follows:
EQUATION RELATED DATA
TERM
VALUE
x Satellite life delta (design life versus actual life)
% =3.0/
3.0 years26%
x Satellite design
11.27
11.27years
x Satellite cost
Sr
$98.8SM
Launch cost for OOR/Satellite
Lc
$214/227M
On-Orbit Refueling Vehicle Cost
Rc
$113M
Table 5-2.
was demonstrated in Chapter ill (Table 3-4),
years
3
of
delta
life
The satellite
using the satellite data from Appendix A. The mean design life of satellites
launched/contracted since 1990 and mean satellite cost (for satellites launched/contracted
since 1990) were also derived from satellite data (See Appendix A). Satellite launch cost
was determined from the International Reference Guide To Space Launch Systems (1991
Edition) [Ref 19] and is shown in Table 5-3. The Titan launch platform with IUS was
chosen for its ability to place the OOR/satellites in geosynchronous orbit.
46
LAUNCH
VEHICLE
MAX PAYLOAD TO GEOSYNCHRONOUS
ORBIT
COST
Titan N (IUS)
5250 lbs (2380 kg)
$214M
TitanN
(Centaur)
10000 lbs (4540 kg)
$227M
Table 5-3
Utilizing the values from Table 5-2 ($214M launch cost), the cost analysis looks like
this:
(Rc + Lc)
= IJ (Sc + Lc )i X (PL~)]
(Where i varies from 1 to N)
($113M+ $214M)= LJ($98.85M + $214M)(i=ll26%)] +
[($98.85M + $214M)(i= 2>(26%)] +
[($98.85M + $214M)<i= 3>(26%)] +
[($98.85M + $214M)<i= 4l26%)] etc.
The break-even or cost effective point actually falls at i = 4.02 satellites serviced.
Using the higher launch cost of $227 for a larger satellite yields a cost effective point at i
= 3.85 satellites serviced.
B.
SENSITIVITY ANALYSIS
A sensitivity analysis on one factor at a time yields the following results:
10% INCREASE
NEW VALUE
SATELLITES TO BE SERVICED
Satellite Design Life
12.4 years24.2%
4.32 satellites
Satellite Life Delta
3.3 years29.28%
3.57 satellites
Satellite Cost
$108.73M
3.90 satellites
OORCost
$138.6M
4.16 satellites
Launch Cost
$235.4M
4.01 satellites
Table 5-4
IN:
47
NEW VALUE
SATELLITES TO BE SERVICED
Satellite Design Life
13.52 years 22.18%
4. 71 satellites
Satellite Life Delta
3.6 years31.94%
3.27 satellites
Satellite Cost
$118.62M
3.78 satellites
OORCost
151.2M
4.30 satellites
Launch Cost
$256.8M
4.00 satellites
20% INCREASE
IN:
Table 5-5
NEW VALUE
SATELLITES TO BE SERVICED
Satellite Design Life
14.65 years 20.47%
5.11 satellites
Satellite Life Delta
3.90 years34.61%
3.02 satellites
Satellite Cost
$128.51M
3.67 satellites
OORCost
$163.8M
4.44 satellites
Launch Cost
278.82M
3.99 satellites
Table 5-6
50% INCREASE
NEW VALUE
SATELLITES TO BE SERVICED
Satellite Design Life
16.91 years17.74%
5.89 satellites
Satellite Life Delta
4.5 years39.9%
2.62 satellites
Satellite Cost
$148.28M
3.47 satellites
OORCost
$189M
4.72 satellites
Launch Cost
$321M
3.98 satellites
Table 5-7
30% INCREASE
IN:
IN:
48
Initial data analysis indicates that an increase in satellite design life would greatly
impact the cost effective point of OOR, increasing the required number of satellites to be
serviced to 5.89 with a 50% increase in design life. However, we are again faced with the
dilemma of onboard fuel; specifically, can you carry enough to meet the increased design
life? This also does not address the associated cost increase necessary to increase
satellite reliability throughout design life. Increases in the satellite life delta yield
promise, in that an increase from three to four years reduces the number of satellites
serviced for cost effectiveness to 3.02 (Table 5-6). This is significant when you consider
that the satellite life delta for the SOPER satellites (Table 3-3) is currently 3.97 years and
still rising. OOR cost increases had the greatest negative impact on cost effectiveness,
raising the number of necessary satellites serviced to 4.72 with a 50% cost increase.
Satellite cost as well as launch cost increases were fairly insignificant.
C.
RISK ASSESS:MENT
Every satellite launched has an associated risk that it may not operate correctly
once placed on-orbit. Extensive testing is conducted to ensure every component will
operate and interface as designed. However, examples such as the Hubble telescope
prove that anything is possible. The risk associated with new technology is usually higher
than previously proven systems, however for the purpose of this analysis the risk of
satellite/OOR failure will be considered comparable.
Launch failure risk will be considered equal in a one-to-one launch ratio.
However, increased launches would represent a higher risk factor. The additional launch
risk can be specifically identified through launch failure/success probability analysis. The
49
specific cost of risk, although not equated to a firm dollar figure, is its impact on program
cost through insurance or an actual launch failure; either would greatly impact program
cost. Increased satellite launches provide increased opportunity to experience a launch
failure. Although insurance can mitigate launch risk, NASA, the U.S. government, and
many large manufacturers typically self-insure due to excessive coverage rates (typically
10% or more of total satellite/launch costs).
D.
FUEL TRANSFER
The OOR design/operations must take into account several factors which may not
seem readily apparent. The first is how much fuel is available to transfer to the satellite.
With an anticipated total fuel load of 800lbs of mono-propellant (primary fuel used in
geosynchronous orbits), the transfer system is envisioned with the ability to cross-feed
both its own thrusters and the refueling system from any fuel tank/cell. This design
feature would preclude the OOR from running out of maneuvering fuel with transferable
fuel still onboard and vice-versa. This would allow maximum flexibility and utilization
of all onboard fuel.
The next decision is how much fuel to transfer to the satellite. The initial impulse
is to "fill-it-up" as the cost difference between lOOlbs and 200lbs of fuel is negligible.
However, when you consider distributing the 200lbs of fuel between two separate
satellites which need refueling the decision becomes more of a dilemma. Maneuvering
fuel on a dead satellite is almost as worthless as a satellite with no maneuvering fuel. An
option could be to conduct a statistical evaluation of remaining satellite life and fueling
50
for perhaps one additional year. This failure analysis would be satellite specific, requiring
failure rates of onboard systems and hence will not be addressed here.
Another operation concern is where to refuel? Should there be a malfunction
during refueling operations the geosynchronous slot could be filled with debris, thus
making it unusable. Perhaps satellites should be boosted to a higher orbit for safety
reasons. This would obviously be driven by reliability and safety design factors of the
OOR as well as fuel budget. The only remaining fuel question is, how much fuel is
required onboard the OOR to service the required number of cost effective satellites?
E.
SATELLITE REFUELING REQUIREMENTS
[All Equations from Space Mission Analysis and Design, Second Edition Ref 21]
How much fuel must be transferred to ensure the satellite will meet the required
life delta of three years? The answer to this question involves many computations based
on specific satellite/orbital parameters.
Satellite North/South drift as well as East/West
drift compensation must be considered. Satellites in geosynchronous orbit have a N/S
drift of approximately .089 degrees/year [Ref 22 p. 155]. Inclination tolerance, or how
much drift above or below the equator can be tolerated, is the driving factor in fuel budget
computations. Typically the time between maneuvers (At) is twice the time it takes the
satellite to drift from the initial orbit insertion point (X 1 ,the lower limit of satellite
inclination tolerance) to the equator or At 1+ A1i, since the drift times are equal. This is
shown in Figure 12.
51
+.1o
Equator
-.10
Figure 12. Satellite N/S Drift
To compute the time interval between N/S station keeping maneuvers (dt}:
at = total inclination tolerance/satellite drift
dt = (2)
X
(.1)/.0897°= .2229 yrs or 81.38 days
Total inclination tolerance is the angle the satellite must pass through to exceed
tolerance. Inclination tolerance is usually payload driven, due to specific pointing or
accuracy requirements. Typically the satellite is first inserted into orbit such that it is at
the lower limit of drift tolerance (in this case at -.1 degrees). This position is represented
by X 1,shown above in Figure 12. The satellite then drifts northward until it reaches the
upper limit of tolerance (X2}, when corrective action must be taken to reposition the
satellite within the payload tolerance requirements.
A velocity is applied to the satellite
using onboard thrusters to return the satellite to X 1• An additional velocity, equal to the
first but in the opposite direction, must also be applied to stop it once it gets there. The
formula for this corrective thrust or delta velocity is:
Jl.v = 2Vsin 8/2 Where e = 2(ai), ll.i =satellite inclination tolerance (.1 for
this example)
V =Velocity in Geosychonous orbit= (631.3481)(RgeoYYz
Rgeo= Earth radius+ Satellite Altitude= 6373km + 35786km =42164km
V = 3.075km/s
Jl.v = 2 (3.075) sin (.1)
ll.v = .0107km/s (10.7m/s) which must be applied every 81.38 days (for this example)
52
Total N/S stationkeeping for an additional three year lifetime must take into
account satellite N/S position at refueling. To believe the satellite would require
fueling at the southern point of inclination tolerance would be overly optimistic. The
optimum position for refueling would be at inclination of zero, or at the equator. This
position would require no inclination change for the OOR, which serves to conserve
onboard fuel assets. Using this positional data, the next satellite repositioning would be
at half the computed at, since the satellite would drift north from the refueling point
(equator) to the upper limit of satellite inclination and the equation above computes total
satellite ~t (X 1 to X2). In this case time to drift from the equator to X 2 would be 40
days, at which point a maneuver must be performed to move the satellite back to X 1•
The normal at intervals of 80 days (for this example) would then apply.
Satellite mass also plays an important part in this problem. Computing the actual
fuel required for the a v, shown above, requires a formula for the mass of propellant as a
fraction of initial satellite mass. Example continued:
Mp = Mo[1-e-<<1vt(IspxG>] Where: avis the computed velocity (from above)
: Isp is the specific impulse of the fuel used
(typically 220-240 for mono-propellant)
: G = gravitational acceleration of the earth
For the example above: Mp/Mo = [ 1-e-<10·7'<220 x9·8>] = .004951, which must be applied to
the mass of the satellite to determine how much fuel is required.
53
Table 5-8 summarizes the Mp/Mo computations for some specific inclination
tolerances, which will then be applied to example satellites.
Inclination Tolerance
at- days
avm/sec
Mp/Mo
.1
81.38
10.73
.004966
.3
244.15
32.20
.014824
.5
406.91
53.66
.024586
.7
569.67
75.14
.034251
1.0
813.83
Table 5-8
107.33
.048563
Applying the computed Mp/Mo for each inclination to satellites of various mass
results in:
Fuellbs/maneuver
Total Maneuvers/3yrs
Total Fuel
Satellite
Mp/Mo
Mass
.1° Jncl
1500lbs
.004966
7.449278
13
105
2000lbs
.004966
9.93237
13
139
2500lbs
.004966
12.41546
13
174
3000lbs
.004966
14.89856
13
209
3500lbs
.004966
17.381
13
226
Required - lbs
Table 5-9
54
Fuellbs/maneuver
Total Maneuvers/3yrs
Total Fuel
Satellite
Mp/Mo
Mass
.3° Incl
1500lbs
.014824
22.23614
4
89
2000lbs
.014824
29.64818
4
119
2500lbs
.014824
37.06023
4
148
3000lbs
.014824
44.47227
4
178
3500lbs
.014824
51.884
4
208
Total Maneuvers/3yrs
Total Fuel
Required
Table 5-10
Fuellbs/maneuver
Satellite
Mp/Mo
Mass
.5° Incl
1500lbs
.024586
36.87906
3
111
2000lbs
.024586
49.17207
3
148
2500lbs
.024586
61.46509
3
185
3000lbs
.024586
73.75811
3
222
3500lbs
.024586
86.051
3
259
Total Maneuvers/3yrs
Total Fuel
Required
Table 5-11
Fuellbs/maneuver
Satellite
Mp/Mo
Mass
.7° Incl
1500lbs
.034251
51.37688
2
103
2000lbs
.034251
68.50251
2
137
2500lbs
.034251
85.62814
2
172
3000lbs
.034251
102.7538
2
206
3500lbs
.034251
119.8785
Table 5-12
2
240
Required
55
Fuellbs/maneuver
Total Maneuvers/3yrs
Total Fuel
Satellite
Mp/Mo1
Mass
1.0° lncl
1500lbs
.048563
72.84478
1
73
2000lbs
.048563
97.12638
1
98
2500lbs
.048563
121.408
1
122
3000lbs
.048563
145.6896
1
146
3500lbs
.048563
169.971
1
170
Required
Table 5-13
Longitudinal drift is primarily caused by the oblatness near the earth's equator.
There are two stable positions (75 and 255 degrees East) and all satellites will drift to the
closest stable point. The total
be expressed as:
fl. v required to maintain longitudinal stationkeeping can
ll.v = 1.7351 sin(2(Ld- L8)) I Where: Ld =the desired longitude
Ls = the closest stable longitude
Assuming the worst case of this equation (sin function= 1), the largest ll.v
possible is 1.735m/sec per year. For the assumed satellite life delta of three years, the
totalll.v is 5.205m/sec. Converting ll.v to a percentage of satellite mass:
Mp/Mo = [1-e-<f:.vt(zzox 9·8)] = .00241, which when applied to various satellites yields:
56
Satellite Mass
Mp/Mo
lbs Fuel Required - 3yrs
1500
.00241
3.66
2000
.00241
4.83
2500
.00241
6.02
3000
.00241
7.24
3500
.00241
Table 5-14
8.44
OOR maneuvering fuel used traveling between satellites must also be considered.
Using the worst case scenario of each satellite being 180 degrees from the previous
refueling target, the total velocity required to reposition (and stop) the OOR can be
expressed as:
ll.v = 5.66(/l.~/n) m/sec Where: fl.~= Degrees oflongitude repositioning
n = Number of days required to reposition
Computing for various values of n yields:
n=Xdays
ll.v- rnlsec
Mp/Mo
30
33.96
.015628
60
16.98
.007844
90
11.32
.005237
120
8.49
.003930
180
5.66
Table 5-15
.002622
Converting this to fuel required, using Mp/Mo (the percentage of OOR mass) for
each ll.v computed in Table 5-15 yields:
57
OORMass
Mp/Mo
Fuel Required - lbs
3500
.015628 (n = 30days)
55
3500
.007844 (n = 60 days)
28
3500
.005237 (n = 90 days)
19
3500
.003930 (n = 120 days)
14
3500
.002622 (n = 180 days)
Table 5-16
10
It is apparent that as n continues to increase the required fuel consumption
decreases. Since it is unlikely that each satellite would require servicing at the same
time, planning for lower fuel consumption is viable. The time/fuel tradeoff would
depend on the urgency of the refueling mission. It is cheaper to burn fuel for
longitudinal changes than inclination changes.
However, OOR inclination changes
may be required, should a satellite be unable to be refueled at the equator. The a v
required for OOR inclination changes can be computed using the equation:
av = 2(Vi)(sin8/2)
where: Vi= velocity at geosynchronous orbit (3.07kms)
= angle of orbit inclination change required
e
Applying OOR inclination changes to various angles yields:
e - Inclination
a v required
m/sec
Mp/Mo
Angle
OORMass
(lbs)
Fuel Required
(lbs)
1
53.58
.024735
3500
75
2
107.16
.048417
3500
146
3
160.73
.071955
3500
216
4
214.30
.094911
3500
285
5
.116889
267.82
Table 5-17
3500
350
58
The inclination fuel requirement is really twice the computed value shown in
Table 5-17, as the OOR must be returned to an inclination of zero to service the next
satellite at the optimum position. Some of this fuel cost could be mitigated by servicing
the remaining satellites at the top of their inclination tolerance vice at the equator,
however, it is obvious that inclination changes are not desirable due to the excessive fuel
required.
OOR maneuvering fuel for docking must also be considered, although there is no
specific formula for this computation. Using historical data from the Challenger
rendezvous and rescue of the Solar Max satellite in 1984, 16lbs of fuel is budget for each
OOR/satellite rendezvous.
59
60
VI: CONCLUSION
A.
SUMMARY SYSTEM TRADES
There can be little argument that in geosynchronous orbit, fuel is the limiting
factor and that the technology exists to conduct on-orbit refueling. However, the cost
effectiveness of OOR is not as clear-cut to determine. Chapter N provided the number
of satellite refuelings necessary to obtain cost effectiveness. These estimates ranged
from three to five satellites, using a life delta of three years. Utilizing this data coupled
with the information in Chapter V on fuel consumption and budgets, an approximation
of fuel required for best and worst case can be compiled. The satellite weights listed in
Appendix A are satellite launch weights, fully fueled. For the purpose of this
comparison, satellites will be assumed near fuel depletion and dry weight estimates will
be used. Fuel budgets for geosynchronous satellites typically range from 600-800lbs of
fuel. [Ref 21, p.330-332] OOR repositioning is evaluated at one less than the number of
satellites to be serviced, assuming the initial OOR orbit insertion will accomplish
positioning near the first satellite to be serviced. Using the basic information from Table
5-2, fuel budget estimates for the best case refueling and yields Table 6-1.
Evidenced by the data in Table 6-1, best case numbers support the refueling of
seven satellites within the initial onboard fuel restriction of 800lbs. However, by
evaluating a satellite inclination tolerance of .5 vice 1.0 degrees would increase fuel
required for each satellite serviced by approximately 40lbs of fuel, which is shown in
Table 6-2.
61
Satellite Mass (Dry)
1500LBS
3 Sats
Serviced
Fuel
Required
4 Sats
Service
Fuel
Required
5 Sats
Serviced
Fuel
Required
6 Sats
Serviced
Fuel
Required
7 Sats
Serviced
Fuel
Required
N/S Station Keeping
(lncl Tolerance 1.0
degrees)
219
292
365
438
511
Satellite E/W
Stationkeeping
11
15
19
23
27
OOR repositioning
n = 120 days
2 Repos
28
3 Repos
32
4Repos
46
5 Repos
60
6Repos
74
Docking Maneuvers
(3)
48lbs
(4)
64lbs
(5)
80lbs
(6)
96lbs
(7)
110 lbs
Inclination Change
OOlbs
OOlbs
OOlbs
OOlbs
OOlbs
TOTAL FUEL
REQUIRED - lbs
306
407
510
617
722
Table 6-1
Note: changes in satellite life delta will alter best/worst case inclinations by
changing the total N/S stationkeeping maneuvers required throughout the chosen life
delta.
Satellite Mass
(Dry) 1500lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
TOTAL FUEL
REQUIRED -lbs
426
567
710
857
1002
Table 6-2
This reduces the number of satellites able to be serviced to five, within the
restriction of 800lbs of onboard fuel, which is still cost effective. Decreasing the time
between satellite refuelings to n =30 days (Table 5-16) increases each OOR repositioning
fuel budget by 4llbs, which yields:
62
Satellite Mass
(Dry) 1500lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
TOTAL FUEL
REQUIRED -lbs
508
690
874
1062
1248
Table 6-3
The resulting increase in OOR repositioning fuel consumption (Table 6-3) reduces
the number of satellites able to be serviced to four, and hence cost effectiveness is
questionable. However, with an inclination change of just one degree (Table 5-21)
indicates an increase in fuel consumption of at least 75lbs, or 150lbs if you return the
OOR to zero inclination. This would directly impact cost effectiveness, reducing the
number of satellites able to be serviced to three.
Re-evaluating the problem using a satellite mass of 2500lbs with optimum
inclination tolerance of 1 degree yields:
Satellite Mass
(Dry) 2500lbs
3 Sats
Serviced
Fuel
Required
4Sats
Serviced
Fuel
Required
5 Sats
Serviced
Fuel
Required
6 Sats
Serviced
Fuel
Required
7 Sats
Serviced
Fuel
Required
Satellite N/S
Stationkeeping
(lncl Tolerance
1 degree)
366
488
610
732
854
Satellite E/W
Stationkeeping
19
25
31
38
44
OOR repositioning
n = 120 days
24
2Repos
42
3 Repos
56
4 Repos
70
Repos
84
6 Repos
Docking Maneuvers
48
64
80
96
112
Inclination Change
00
00
00
00
00
TOTAL FUEL
REQUIRED - lbs
461
619
777
936
1034
Table 6-4
63
Evidenced by the data in Table 6-4, the refueling of five satellites can be
conducted within the onboard fuel restriction of 800lbs, which still meets cost
effectiveness criteria.. However, by evaluating a satellite inclination tolerance of .5
degrees (worst case) fuel required for each satellite serviced would increase by
approximately 65lbs, yields:
Satellite Mass
(Dry) 2500lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
TOTAL FUEL
REQUIRED- lbs
656
879
1104
1326
1489
Table 6-5
This reduces the number of satellites able to be serviced to three, within the restriction of
800lbs of onboard fuel, which is not cost effective.
Decreasing the time between satellite refuelings to n = 30 days (Table 5-16)
increases each OOR repositioning fuel budget by 41lbs, which yields:
Satellite Mass
(Dry) 2500lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
TOTAL FUEL
REQUIRED - lbs
738
1002
1266
1531
1735
Table 6-6
The resulting increase in OOR repositioning fuel consumption (Table 6-6) does
not reduce the number of satellites able to be serviced below three but cost effectiveness
is certainly not going to increase.
Re-evaluating the problem using a satellite mass of 3500lbs with optimum
inclination tolerance of 1 degree yields:
64
Satellite Mass
(Dry) 3500lbs
3 Sats
Serviced
Fuel
Required
4Sats
Serviced
Fuel
Required
Satellite N/S
Stationkeeping
(lncl Tolerance
1 degree)
510
Satellite E/W
Stationkeeping
5 Sats
Serviced
Fuel
Required
6 Sats
Serviced
Fuel
Required
7 Sats
Serviced
Fuel
Required
680
850
1020
1190
27
36
45
54
63
OOR repositioning
n = 120 days
28
2Repos
42
3 Repos
56
4Repos
70
Repos
84
6Repos
Docking Maneuvers
48
64
80
96
112
Inclination Change
OOlbs
OOlbs
OOlbs
OOlbs
OOlbs
TOTAL FUEL
REQUIRED -lbs
613
822
1031
1240
1449
Table 6-7
Evidenced by the data in Table 6-7, three satellites can be refueled within the
800lb OOR fuel restriction, which does not meet cost effectiveness criteria. Again
evaluating a satellite inclination tolerance of .5 degrees (worst case) would increase fuel
required for each satellite serviced by approximately 90lbs, yielding:
Satellite Mass 3500lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
TOTAL FUEL- lbs
883
1182
Table 6-8
1481
1780
2079
This reduces the number of satellites able to be serviced to two, within the
restriction of 800lbs of onboard fuel, which is not cost effective. Decreasing the time
65
between satellite refuelings to n = 30 days (Table 5-16) increases each OOR repositioning
fuel budget by 41lbs, which yields:
Satellite Mass
(Dry) 3500lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
TOTAL FUEL
REQUIRED - lbs
965
1305
1645
1985
2325
Table 6-9
The resulting increase in OOR repositioning fuel consumption (Table 6-9) does not
reduce the number of satellites able to be serviced below two, but cost effectiveness is
certainly not going to improve.
However, actual on-orbit refueling targets will probably consist of a cross section
of satellite sizes, instead of all of one size as examined in the examples above. Reevaluating the problem using a cross section of satellite sizes 1500lbs, 2500lbs, and
3500lbs with optimum inclination tolerance of 1 degree yields the results shown in Table
6-10.
Evidenced by the data in Table 6-10, five satellites can be refueled within the
800lb OOR fuel restriction, which meets cost effectiveness criteria. Again evaluating a
satellite inclination tolerance of .5 degrees (worst case) would increase fuel required for
each satellite serviced by approximately 40lbs, 65lbs, and 90lbs, respectively, and is
shown in Table 6-11.
66
5 Sats
Serviced
Fuel
Required
6 Sats
Serviced
Fuel
Required
7 Sats
Serviced
Fuel
Required
438
560
682
982
20
40
52
61
65
OOR repositioning
n = 120 days
28
2 Repos
42
3 Repos
56
4Repos
70
Repos
84
6 Repos
Docking Maneuvers
48
64
80
96
112
Inclination Change
OOlbs
OOlbs
OOlbs
OOlbs
OOlbs
TOTAL FUEL
REQUIRED - lbs
461
584
748
909
1243
Satellite Mass
(Dry) 2 of each
rotation order- 1500,
2500, & 3500lbs
3 Sats
Serviced
Fuel
Required
4 Sats
Serviced
Fuel
Required
Satellite N/S
Stationkeeping
(lncl Tolerance
1 degree)
365
Satellite E/W
Stationkeeping
Table 6-10
II Satellite Mass 3500lbs
II TOTAL FUEL -lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
656
819
Table 6-11
1048
1299
1673
This reduces the number of satellites able to be serviced to three, within the
restriction of 800lbs of onboard fuel, which is not cost effective. Decreasing the time
between satellite refuelings to n =30 days (Table 5-16) increases each OOR repositioning
fuel budget by 41lbs, which yields:
67
Satellite Mass
(Dry) 3500lbs
3 Sats
4 Sats
5 Sats
6 Sats
7 Sats
TOTAL FUEL
REQUIRED - lbs
738
942
1221
1504
1919
Table 6-12
The resulting increase in OOR repositioning fuel consumption (Table 6-12) does not
reduce the number of satellites able to be serviced below three, but cost effectiveness is
certainly not going to improve.
Obviously, as satellite mass increases the cost effectiveness of on-orbit refueling
decreases. However, the initial design limitation of 800lbs is not carved in stone. With
an increase to 1250lbs of fuel, the cost effectiveness for five satellites can be maintained
throughout all examples with the exception of 3500lb satellites computed in Tables 67/8/9. Launch capability of the Delta IV-IUS is 5200lbs which would allow for an
increased fuel payload. Consulting the Appendix A satellite data reveals that only 20
satellites exceed 3000lbs fueled, hence limiting the possibility of the latter fuel
computational restrictions shown in Tables 6-7/8/9. By increasing the OOR fuel
payload, the impact on the OOR cost would be minimal, with OOR structure and the fuel
transfer package being the most obvious areas for cost increases (i.e., a larger/heavier
structure in order to support the additional fuel weight and a heavier fuel transfer package
for the additional fuel tanks and piping required). These OOR cost areas (discussed in
Chapter IV) do not carry a significant cost multiple and hence would not greatly impact
OOR cost. Additionally, factors such as time between satellite refueling and inclination
changes can be managed to reduce fuel impact.
68
SOME FINAL THOUGHTS
B.
Planning and cost analysis was done assuming the OOR was a "throw away" or
one time only use vehicle. If the OOR was constructed using modular/ORU fuel cells and
could be refueled in space for additional missions, this would greatly improve OOR cost
effectiveness. This "refueling of the refueler" would probably have to occur in low earth
orbit (LEO), perhaps as a space station mission. Replacement fuel cells could be
launched onboard shuttle flights as space available cargo. thus saving launch costs for
future OOR missions and further enhancing cost effectiveness. However, the de-orbit to
LEO would have negative impact on OOR fuel. This impact could possibly be limited by
the Reusable Orbital Transfer Vehicle concept which proposes use of the Earth's
atmosphere to slow and capture the spacecraft. thus obtaining low earth orbit after initial
de-orbit. This concept would require further analysis which exceeds the scope of this
paper.
Satellite on-orbit refueling is both cost effective and tactically significant. As
satellite program costs continue to increase and operations and research budgets continue
to decrease the cost savings and operational flexibility provided by on-orbit refueling
cannot be ignored.
69
70
APPENDIX A:
SATELLITE DATA SUMMARY
[Ref? and 22]
71
SATELLITE
-.J
IV
COMSTAR 01
COMSTAR 03
COMSTAR 02
FLTSATCOM 1
FLTSATCOM 2
WESTAR 3
FLTSATCOM 3
GOES4
FLTSATCOM 4
SBS 1
INTELSAT 502
COMSTAR 04
GOES5
INTELSAT 501
FLTSATCOM 5
SBS2
SATCOM 3R
INTELSAT 503
SBS3
SATCOM4
WESTAR 4
INTELSAT 504
WESTAR 5
INTELSAT 505
AURORA1
DSCS 111-1
DSCS 11-15
TORS F1
SATCOM 1R
GOES6
LAUNCH
SATELLITE
LAUNCHED WEIGHT KG COST$M
o5t13n6
05/29n6
07/22n6
o21o9n8
05/o4n9
08t1on9
01/17/80
09/09/80
10/30/80
11/15/80
12106/80
02121/81
05/22181
05/23/81
08/06/81
09/24/81
11/20/81
12115/81
01/11/82
01/15/82
02126/82
03/05/82
06/08/82
1516
1516
1516
1005
1005
574
1005
900
1005
546
1928
1516
900
1928
1005
546
1078
1928
546
21.25
21.25
21.25
25.66
50
33
21.25
33
50
33
50
1078
572
1928
1100
50
33
50
1928
33
09/28/82
10/28/82
10/30/82
10/30/82
04/04/83
1084
1042
590
2200
04/11/83
04/26/83
1120
900
150
100
50
LAUNCH
VEHICLE
DESIGN vs
~CTUAL LIFE
REMARKS
I
FUEL DEPLETED 1984
7\8.3
7\10
FUEL DEPLETED 1986
7\17
FUEL DEPLETED 1993
5\18 SOPER FUEL CONSERVE OPS(FCO) MAR 95
5\15
FCO EOL 1992
7\11
EOL 1990
5\15
FCO
5\2 FAILED
EOL 1988 SUPERSYNC BOOST
ATLAS CENTAUR
5\15.3
FCO
DELTA
7\9.4
FUEL DEPLETED 1990
ATLAS CENTAUR 7\16 SOPER
FCODEC 1988
ATLAS CENTAUR 7\15 SOPEF
THOR DELTA
5\3-8
MAGE FAIL '84>RELAY MSN, FUEL '89
ATLAS CENTAUR 7\15 SOPER
FCOMAY 1988
ATLAS CENTAUR
5\5
DAMAGED ON LAUNCH EOL 1986
DELTA
7\14.7 SOPER
FCOOCT 1989
DELTA
10\10
EOL 1991
ATLAS CENTAUR 7\15 SOPER
FCO MAY 1989
SHUTILE
7\13.5
FUEL DEPLETED JUN 1995
DELTA
10\9
EOL 1991
DELTA
10\9.5
FAILED NOV 1991
ATLAS CENTAUR
7\13.5
FCOAPR 1989, DIED NOV 1995
DELTA
10\10
FUEL DEPLETED MAY 1992
ATLAS CENTAUR 7\14 SOPER
FCOAPR 1989
THOR DELTA
10\14 SOPER
OK
TITAN
10\14
FCOMAR 1995
TITAN
10\14
FCOMAR 1995
SHUTILE
10\12 SOPER JS FAIL·REQ USE OF 370KG MAN FUI
DELTA
10\
EOL
THOR DELTA
5\6-9
IMAG FAIL '89\COMM RELAY '92 FUEL
ATLAS CENTAUR
ATLAS CENTAUR
ATLAS CENTAUR
ATLAS CENTAUR
ATLAS CENTAUR
DELTA
ATLAS CENTAUR
~·
SATELLITE
-:J
w
INTELSAT 506
GALAXY 1
TELESTAR 3A
SATCOM 2R
GALAXY2
INTELSAT 508
SPACENET 1
INTELSAT 510
TELESTAR 3C
LEASAT 2
SBS4
GALAXY3
LEASAT 1
WESTAR 6
SPACENET2
LEASAT3
GSTAR 1
TELESTAR 3D
INTELSAT 511
ASC 1
INTELSAT 512
DSCS 111-2
DSCS 111-3
INTELSAT 507
SATCOM K2
SATCOM K1
GSTAR2
FLTSATCOM 7
GOES?
SATELLITE
LAUNCH
LAUNCHED WEIGHT KG COST$M
05/19/83
06/28/83
07/29/83
09/08/83
09/22183
03/05/84
05/02184
06/09/84
08/30/84
08/30/84
08/30/84
09/21/84
11/08/84
02104/84
11/10/84
04/12185
05/08/85
06/17/85
06/29/85
08/27/85
09/28/85
10/03/85
10/03/85
10/19/85
11/27/85
01/12186
03/28/86
12104/86
02126/87
LAUNCH
VEHICLE
DESIGNvs
!ACTUAL LIFE
ATLAS CENTAUR 7\13 SOPER
1928
33
10\SOPER3
DELTA
1200
33.33
10\13 SOPER
3423
45.66
DELTA
10\11.5
DELTA
1120
33.33
DELTA
10\10.5
1200
ARIANE
9\11
1928
33
ARIANE
10\12 SOPER
1195
75
CENTAUR 9\12 SOPER
2013
. --------- -·--------------------------- - · - ATLAS
3423
6894
1117
1200
6894
1195
6894
1200
3423
2013
1250
2013
1170
1170
1928
1900
1900
1270
1100
835
45.66
85
50
33.33
85
75
75
85
33.33
45.66
33.33
150
150
33
38
38
33.33
125
92
SHUTILE
SHUTILE
SHUTILE
DELTA
SHUTILE
ARIANE
SHUTILE
ARIANE
SHUTILE
ATLAS CENTAUR
SHUTILE
ATLAS CENTAUR
SHUTILE
SHUTILE
ARIANE
SHUTILE
SHUTILE
ARIANE
ATLAS CENTAUR
DELTA
---·-··
..
- ---
.. ----- -·· --
FCO SEP 1993
------EOL 1994
FCO
----EOL 1995
CHINASAT FCO
---- -- -FCOAUG
1992
. -- --------- ------.
FCOJAN 1995
FUEL DEPLETED MAY 1994 ---.FCO
---~
1-o\f2-s6PER
7\11.5
10\11.5
10\11
7\8
REMARKS
1991
---- ---- ·-
--------- FCOJAN
---
---1
FUEL DEPLETED OCT 1~-~?___
FAILED SEP 1992
BOOST FAIL REOD REPAiR- - -i
- -- I
OK
!
-- ----· - --·
OK
-----------------·. --FCO
----- ·-·-··--FCO
----.
FCOAUG 1994 -·-···· -···
10\12 SOPER
7\10.3 SOPER
10\11 SOPER
10\11 SOPER
9\11 SOPER
10\9
--~Q_L~-~ 1994?£_~<:)
FCOAUG 1994 ----. ----9\11 SOPER
10\11 SOPER
-------10\11 SOPER
-··--- ...
FCOAUG 1990
7\11 SOPER
10\11 SOPER
OK
10\10 SOPER
OK
---------FCOOCT 1993
10\10 SOPER
----FCO
5\9.1 SOPER
FUEL DUE TO NUM OPMOVES 1995
5\8
----=~ :-J
LAUNCH
REMARKS
-- -···-·· ----
SPACE NET3R
iiNTELSAT 513
PAS 1
GSTAR 3
SBS5
TORS F3
INTELSAT 515
TORS F4
OSCS 111-4
FLTSATCOM 8
INTELSAT 602
-...]
~
···-- ------ ------ ------ ------ 03/11/88
05/17/88
06/15/88
09/08/88
09/08/88
09/29/88
01/27/89
03/13/89
09/04/89
09/25/89
10/27/89
1250
2013
1220
1250
1241
2200
2013
2200
1170
1100
4600
SATELLITES LAUNCHED SINCE 1990
LEASAT 5
6894
INTELSAT 603
4600
INTELSAT 604
4600
GALAX Y6
01/09/90
1212
SBS6
03/14/90
2478
SATCOM C1
06/23/90
1169
GSTAR 4
10/12/90
1295
SPACE NET4
10/12/90
728
AUROR A2
11/20/90
1338
TDRSF 5
11/20/90
2200
INTELSAT 605
04/13/91
4600
INTELSAT 601
05/29/91
4600
GALAX Y5
08/02/91
1412
GALAX Y?
08/14/91
2986
SATCOM C4
10/29/91
1402
SATCOM C3
03/14/92
1375
75
40
33.33
50
250
100
150
140
ARIANE
ARIANE
ARIANE4
ARIANE
ARIANE
SHUTTLE
ARIANE
SHUTTLE
TITAN
ATLAS CENTAUR
ARIANE
-------]
.
10\8 SOPER
----FCOA~~-i995 ___ .:~~9\8 SOPER
11\8 SOPER
FUEL FOR 13.5 YEARS
i
10\8 SOPER BOOST FAIL- USED 80% MAN FUEL I
10\8 SOPER
OK
10\8 SOPER
__ ______ FCO ___ ...
9\7 SOPER
FCOMAY 1997
---------- --10\8 SOPER
------- --- - -- -- -- ----10\6 SOPER
7\6.3 SOPER
--13\7 SOPER
OK
I
-------
85
140
60
75
88
30
140
140
SHUTTLE
TITAN
TITAN
ARIANE4
ARIANE4
ARIANE
ARIANE
DELTA2
DELTA
SHUTTLE
ARIANE
ARIANE
ATLAS 1
ARIANE4
DELTA
ARIANE 4
------7\6 SOPER
OK
13\6 SOPER NEW KICK MTR INSTALL STS49 1992
13\6 SOPER
OK
10\6 SOPER
OK
10\6 SOPER ONBOARD FUEL FOR 15.6 YEARS
10\6 SOPER
OK
10\6 SOPER
------- ------10\5 SOPER
OK
12\5 SOPER
10\5 SOPER
OK
15\6 SOPER
OK
13\5 SOPER
OK
10\4 SOPER
ONLY FUEL FOR 9 YEARS
15\4 SOPER
ONLY FUEL FOR 12 YEARS
12\4 SOPER
OK
---12\4 SOPER
~---·-·--
--
--
-·-
--
-
SATELLITE
--J
U1
INTELSA TK
INTELSAT 701
INTELSAT 702
INTELSAT 703
INTELSAT 704
INTELSAT 705
INTELSAT 706
INTELSAT 707
INTELSAT 708
INTELSAT 709
PAS4
PAS3R
GALAXY 3R
GALAXY 9
TDRS6
TORS?
TELSTAR 402R
UHF 1 (UFO)
UHF2
UHF3
UHF4
UHF5
LAUNCH SATELLITE
LAUNCHED WEIGHT KG COST$ M
UH~
08/01192
08/31/92
09/10/92
01/13/93
03/25/93
09/03/93
10/22/93
06/17/94
06/24/94
01/01/95
03/01/95
03/08/95
05/01/95
05/31/95
09/01/95
10/01/95
10/22/95
12/15/95
01/01/96
03/01/96
05/24/96
06/01/96
12/01/96
2836
3610
3610
95
78.85
78.85
78.85
78.85
78.85
3610
3610
3610
4643
4643
4643
4643
2985
2985
3380
140
90
75
75
2200
2200
3000
3000
3000
3000
3000
3000
3083.7027027
LAUN'CHED MEAN
LAUNCH
VEHICLE
DESIGN vs
!ACTUAL LIFE
ATLAS CENTAUR 10\6 SOPER
ARIANE4 4
65% at 1Oyrs
ARIANE44
65% at 1Oyrs
ATLAS II
65% at 1Oyrs
ATLAS II
65% at 1Oyrs
ATLAS II
65% at 1Oyrs
ARIANE4 4
65% at 1Oyrs
LONG MARCH3B 65% at 1Oyrs
ARIANE44
65% at 1Oyrs
ARIANE4 4
65% at 1Oyrs
11
ATLAS IIA
10
10
SHUTTLE
SHUTTLE
ARIANE
ATLAS CENTAUR
ATLAS CENTAUR
ATLAS CENTAUR
ATLAS CENTAUR
ATLAS CENTAUR
ATLAS CENTAUR
172.2
64.4
60.1
87.2
79.3
65.4
89.8270833333
REMARKS
OK
..
FUEL FOR 15YRS
LAUNCH FAILURE
-----------
--------- ----------
-·------·-
-- ------------ --------
10
10
12
10-15YRS
10-15YRS
10-15YRS
10-15YRS
10-15YRS
10-15YRS
----···-·----
----- --
PARTIAL LAUNCH FAILURE - ORBIT
--------··· -··
.
---------
.
----
...
-
-·
--------
CONTRACTS
LAMSC/M~AI
_l_
12/01/90
l
1
100
l
l
FOR COST ANALYSIS ONLY
I
SATELLITE
-..]
0'\
INMARSAT F1
INMARSATF2
INMARSAT F3
INMARSAT F4
ASIASAT (GE)
INTELSAT 801
INTELSAT 802
INTELSAT 806
TDRS8
TDRS9
TDRS10
UHF7
UHF8
UHF9
UHF10
CONTRACT MEAN
LAUNCH SATELLITE
LAUNCHED WEIGHT KG COST$M
02101/91
02101/91
02101/91
02101/91
09/01/92
09/01/92
09/01/92
03/01/94
02101/96
02101/96
02101/96
LAUNCH
VEHICLE
80
80
- - r------------·- -80
80
133
82.5
82.5
82.3
160.5
160.5
160.5
94.4
121
166
135 - -
--
3000
3000
3000
__ 3000
--
-
DESIGN vs
!ACTUAL LIFE
---- -·-
J
REMARKS
-·---~---------
--- ··--------
···-
--··
- ....... ···-··
------------- --------------· --· ········-------~----~
------------------- .
-
----------
------------
---------------.
---------
10-15YRS
10-15YRS
10-15YRS
10-15YRS
--------------- ------------------ --------------.
112.3875
!LAUNCHED/CONTRACT MEAN FOR SATELLITES SINCE 1990
98.85125
-------m----- ---·-- ]
APPENDIXB:
DSCS IIIB SATELLITE DATA
[Ref20]
77
SATEWTE NAME:
DSCS IIIB
GENERA L
BLOCK l OESIGNA TEO UNfTS:
SKETCH
Block B. Units B4 through B7
VSER{CONTRA cnNQ;
AF SMC
CONTRACTOR:
CO!il'RA.CT COMPLETTON:
General Electric
TYPE OF CONTRACT:
FPIF
PRECECESSCR VEHICLES:
DSCS Ill -A1, A2. and A3
LAUNCH· WEIGHT (U3):
DRY WEIGHT (LB):
1883.70
LAUNCH VEHICLE;
FIRST LAUNCH DATE:
LAST LAUNCH DATE:
Atlas II, Shuttle. Dual
Compatible, Centaur & IUS
Classified
May 1994
DESIGN UFE
(YR):
NO. OF NEW UNITS:
NO. OF OUAL UNITS
NO. OF PFIOO UNITS
10
0
0
4
ORBIT PARAMETERS
APOClEE (NMI);
PERIGEE (NMI):
INCUNATION (OECi):
Synchronous
Synchronous
0.1
MISSION DESCRIP TION
The DSCS Ill was developed for the Air Force by GE. Its mission is to provide uninterrupt
ed secure strategic
and tactical voice and data transmission, military command and control, and
ground mobile communica tions.
This is achieved by antijam abilities and high frequency wideband communica
tions. Block B consists of 11
satellites. some of which have already been delivered. Block B satellites are
covered in two data packages.
Block B1 consists of satellites B4-B7 and Block B2 consists of satellites B8-B14.
These satellites have some solid
state amplifiers replacing TWTAs, a new X-band downlink capability for the
AFSATCOM transponder, and
improved security equipment.
DESCRIB E All CAVEAT S
The average recurring cost (and resulting first unit cost) for DSCS IIIB was significantl
y higher than that of
DSCS lilA. The common belief is that significant overruns were incurred on the DSCS
iliA contract, particular1y by
the subcontractors operating under fixed price contracts. As a result. the DSCS
IIIB recurring costs are more
representative of the "true" recurring costs of the DSCS Ill program. Therefore.
the data point for DSCS IIIB units
4 through 7 was used only in developing recurring CERs. Furthermore. due
to the similarity of the two DSCS IIIB
data points. in several cases these blocks were combined to form one data point
and a new first unit cost was
calculated.
MAIN SUBSYS TEM/CO MPONEN T COST DRIVERS
srnuCTURE WEIGHT (LB):
STRUCTURE
MATERIAL TYPE
330 48
Aluminum. magnesium. beryllium. magnesium thorium
THERMAL
THERMAL WEIGHT (LB)
AVERAGE TEMP RA.NGE (DEG FAHRENHEfT)
TYPE OF THERMAl.
CONTROL
10193
70
Passivejsemi·passive surface coating: single and multr·layer insulation: mirrored
surface:
passive conduction. Active;'semi·active· heaters and radiators.
78
SATEWllO NAME:
DSCS 1118
_(f
MAIN SUBSYSTEM COMPONENT COST DRIVERS (Cont'd)
·f
ATTITUDE DETERMINATION & CONTROL SYSTEM
162.28
NO. Of TANKS: 4
POINTlNG ACCURACY (DEC):
98.83 (included in ACS weight)
~
J
ACS WEICHT (t.B):
RCS WEIGHT (t.B):
0.10
Non-scanning Earth Sensors (2); Rate Gyro Assembly (1); Sun Sensors {2)
Fuel-1388.7 5; Oxidizer- 4165.25
Reaction Wheels
SENSOR TYPES ~NCLUDE NO. OF EACH TYPE):
TANK VOLUME (CU IH):
TORQUE METHODS:
ELECTRICAL POWER SYSTEM
EPS WEIGHT (t.B):
SOL POWER (WATTS):
SOLAR ARRAY AREA (SO Fl):
585.59
Summer solstice: 1310;
Autumnal Equinox: 1397
126
8ATTERY TYPE
(AND NO.)
WEIGHT OF ONE BATTERY (LB}:
DISTRIBUTION WEICHT (LB):
PCE SUIT£ WEICHT
(L8):
SATTERY CAPACfTY
(AMP-HR):
NiCd (3)
45.16
211.43
113.90
35.0 per battery
I
TOTAL IMPULSE:
NO. OF SOLAR CELLS:
GENERAnON WEIGHT (t.B):
31.72
APOGEE KICK MOTOR
I
AKM DRY WEIGHT (1.8):
STABIUZATION loiETHOD:
TELEMETRY, TRACKING, AND COMMAND SYSTEM
TUC WEICHT (t.B):
POWER REQUIRED
RF' POWER OUTP!JT
(WATTS):
(WATTS):
TWTA OR SOUO STATE
AMPS:
Solid State
9.00
RECEIVER FREQUENCY
(MHZ):
TRANSMITTER WEICHT
(L8):
TlUNSMfrrER
FREQUENCY (MHZ):
TRANSPONDER
WEICHT (LB):
TRANSPONDER FREQUENCY (MHz):
7600
14.16
7600
18.04
Receive A: 1807.764; 8: 1823.779
Transmit A: 2257.5; 8: 2277.5
DICITAL ELECT WEICHT
(L8):
ANALOG ELECT
ANTENNA WEICHT (l8):
WEICHT (L8):
ANTENNA APERTURE
(INCHES):
ANTENNA CAIN (DECIBELS):
Not specified
65%: -7.5; 25%: -4.5
70.99
23.00
N/A
TR.ANSMfTTER OIJTPUT
POWER (WATTS):
TRANSPONDER OlfTPVT
POWER (WATTS):
0.7
2.0
COMMWEIOHT
(L8):
I
POWER REQUIRED
(WATTS):
632.40
RECEIVER FREQUENCY (MH:z:):
j
0.67
I
DICITAL ELECT OUTPUT
POWER (WATTS):
I
RECEIVER WEICHT (LB):
DATA RATE <••1•1:
Command, Real Time: 1: Telemetry, Real Time: 1
COMMUNICATION
RF POWER OUTPUT (WATTS):
I
TWTA OR SOLID STATE AMPS:
I
RECEIVER WEICHT (L8):
TVVTA & Solid State
51.26
Freq Gen 5.00; SCT Converter. not specified; Freq Synth. not specified; SCT-SHF, 7975
to 8025; SCT-UHF, Classified
TRANSMITTER FREQUENCY (MHz):
Freq STD. 5.0; LNA, not specified; TDAL. 7900 to 8400; TDL. not specified; TVVTA-10W: 7400
to 77500;
TWTA-40W: 7250 to 7400; SCT-SHF: 8000; HESSA, not specified
TRANSiolrTTER WT
(LB):
131.25
I
TRANSPONDER WT
(L8):
NjA
l
TRANSPONDER FREQUENCY
(MHz):
N/A
I
DICITAL ELECT WEICHT (l.B)o
1
ANACOC ELECT
WEJCKT (LB):
56.96
9.42
l
AflTENHA WEICHT
(LB):
305.60
ANTENNA APERTURE {INCHES):
ECH-R. 6.5; ECH-T. 7.7; Gimballed Dish Antenna. 855; 61 MBA Receive. 45; 19 Transmit
MBA, 28; UHF Receive. not specified; UHF Transmtt. not specified
ANTENNA PEAK CAIN (DECIBELS):
ECH-R. 168 dBi; ECH-T. 17.0 dBi; Gimballed Dish Antenna. 30.2; 61 MBA Receive,
narrow coverage-29.4. earth coverage-14.4; 19 Transmit MBA, narrow
coverage-26 /26.5, earth coverage-s1615; UHF Receive & Transmit, classified
TRANSMITTER OUTPl!T POWER (WATTS):
TRANSPONDER OUTPUT POWER ('WATTS):
N/A
Freq STD. 2.02; LNA. MBA-1.5~. ECH-3.08; TDAL. Fetal-5.8; TDL, not specif~~':!.
TVVTA-10W, 10; TWTA-40W. 40; SCT-SHF, 50; HESSA, 10
I
DICITAl ELECT OUTPUT POWER (WATTSio
Not specified
I
DATA RATE
<••!•):
SCT digital processor, real time: 0.07
79
:
80
APPENDIXC:
OOR NON-RECURRING COSTS ESTIMATES
[All Equations from Unmanned Space Vehicle Cost Model, Seventh EditionRef. 20]
Satellite non-recurring cost consists of the Research, Development, Test and
Evaluation (RDT&E) which typically includes design, analysis, testing, prototypes and
qualification runs. Additionally, it also includes ground station costs. The non-recurring
cost estimate uses the same CER methodology used to estimate recurring cost in Chapter
N. and are summarized in Table C-1. Non-recurring cost estimates for the OOR are as
follows:
1. STRUCTURE
750 lbs
Spacecraft Structure
Y=(99.045)(Xl)0·789
Where
Xl= Structure Weight
Y= CER value for Spacecraft Structure
Therefore
Y= 18376.21
2.THERMAL
Thermal Weight
Y= (0.243)(Xl)o.s97 + (X2)o.9s3
Where
165lbs
Xl= Thermal Weight
X2= Satellite Weight
Y =CER value for Spacecraft Structure
Therefore
Y=12364.23
3. ADCS
81
ADCS
- Determination Suite Weight
- RCS Suite Weight
- Total ADCS WeiKht
Y=(666.439)(Xl)0·711
Where
180 lbs
200 lbs
380 lbs
Xl= Attitude Determination Suite Weight
Y =CER value for ADCS (Attitude Determination)
Therefore
Y = 26746.02
Y=(125.998)(Xl)0·733
Where
Xl= Reaction Control System Suite Weight
Y = CER value for ADCS(Reaction Control)
Therefore
y = 6123.77
4. ELECTRICAL POWER SYSTEM
EPS
- Number of Solar Cells
- Generation Suite Weight
- Beginning of Life Power
- Storage Suite Weight
-EPS Suite Weight
Y=(0.025)(Xl) + (0.024)(X2)
Where
3000
32lbs
1200 Watts
135lbs
585.59
Xl= (Generation Suite Weight)(Beginning Life Power
(BOL))
X2= Number of Solar Cells
Y =CER value for Electrical Power Generation
Therefore
Y = 1032
82
Y=(l14.127) + (2.584)(X1)
Where
X1= (Weight of One Battery)(Capacity of One Battery)
Y
Therefore
=CER value for Electrical Power Storage
y =4183.93
Y = (5.515)(X1)
Where
X1= BOL Power
Y = CER value for Power Conditioning and Distribution
Therefore
y =6618.0
5. TELEMETRY, TRACKING AND CONTROL
TT&C
- Transmitter
- Receiver/Exciter
- Digital Electronics (2 Links)
- Antenna (4 Systems)
Y =(67 .121 )(X1)
Where
JOlbs
9lbs
23lbs
4lbs
X1= Transmitter Suite Weight
Y = CER value for TT&C Transmitter
Therefore
y = 671.21
Y = (-224.351) + (116.683)(X1)
Where
X1= Receive/Exciter Suite Weight
Y
Therefore
=CER value for TT&C Receiver/Exciter
y =825.80
y = (211.243)(X1)o.7s7 (X2)o.ss3
Where
X1= Digital Electronics Suite Weight
83
X2= Number of Links
Y = CER value for TT&C Digital Electronics
Therefore
y =4500.29
Y = (-222.262) + (30.670)(Xl) + (480.840)(X2)
. Xl= Antenna Suite Weight
Where
X2= Number of Antenna Systems
Y =CER value for TT&C Antenna
Therefore
y = 1823.78
6. COMMUNICATIONS
Communications Transmitter (TWTA)
- TWTA Weight
- Solid State Transmitter
- Receiver/Exciter
- Transponder (2 units)
- Digital Electronics(5 links)
- Antenna (4 systems)
-Antenna Reflectors
Y=(524.16l)(XI)0·875
14.6lbs
51.26 lbs
30lbs
30lbs
56.96lbs
141lbs
8sqft
XI= TWTA Weight
Where
Y = CER value for Communications Transmitter (TWTA)
Therefore
y = 5473.61
Y= (0.249)(XI)u01 (X2)0·728
Where
XI= Solid State Transmitter Weight
X2= Transmitter Frequency
Y = CER value for Communications Transmitter (Solid
State)
84
Therefore
Y=5283.18
Y=(273.793)(Xl)
Where
Xl= Receiver/Exciter Suite Weight
Y =CER value for Communications Receiver/Exciter
Therefore
Y=8213.79
Y=(682.7 69)(X 1)0·463
Where
Xl= Transponder Weight
Y = CER value for Communications Transponder
Therefore
Y=3297.47
Y=(211.243)(X1)0·787 (X2)0 ·853
Where
Xl= Digital Electronics Suite Weight
X2= Number of Links
Y =CER value for Communications Digital Electronics
Therefore
Y= 20074.15
Y=(-222.262) + (30.670)(Xl) + (480.840)(X2)
Where
Xl= Antenna Suite Weight
X2= Number of Antenna Systems
Y = CER value for Communications Antenna
Therefore
Y =6025.57
Y=(1763.889)(Xl)
Where
Xl= Antenna Reflector Diameter Squared
85
Y = CER value for Communications Antenna Reflectors
Therefore
Y= 14111.11
7. INTEGRATION ASSE:MBLY AND TEST (IA&T)
IA&T
- Spacecraft Weight
- Fuel Transfer System (FTS) Total Weight
-Weight
Y= 956.384 + (0.191)(Xl)
Where
2462lbs
300 lbs
2762lbs
Xl= Spacecraft Weight+ Payload (FTS) Non-Recurring
Cost
Y = CER value for IA&T
Therefore
y =29396.41
8. PROGRAM LEVEL
Satellite Total Recurring Cost
y =(2.340)(X 1)o.8o8
Where
113000K
Xl= Spacecraft Total Non-Recurring Cost
Y = CER value for Program Level
Therefore
Y= 28320.65
9. AEROSPACE GROUND EQUIP:MENT (AGE)
Satellite Total Non-Recurring Cost
Y=(8.304)(Xl)0·638
Where
149976.41 K
Xl= Space Vehicle Total Non-Recurring Cost
Y =CER value for Aerospace Ground Equipment
Therefore
y = 16656.73
86
NON-RECURRING COST SUMMARY
(in Thousands of Dollars)
Structure
18376.21
Thermal
12364.23
Attitude Determination & Control
26746.02
ADCS - Attitude Determination
RCS
6123.77
Electrical Power Supply
EPS - Generation
EPS - Storage
EPS - PCD
1032.00
4183.93
6618.00
Telemetry, Tracking & Command
TT&C - Transmitter
TT&C - Receiver/Exciter
TT&C - Digital Electronics
TT&C - Antenna Suite
671.21
825.80
4500.29
1823.78
Communications
Comm - Transmitter (TWTA)
Comm - Solid State
Comm - Receiver/Exciter
Comm - Transponder
Comm - Digital Electronics
Comm - Antenna
Comm - Antenna Reflectors
5473.61
5283.18
8213.79
3297.47
20074.15
6025.57
14111.11
3156.52
Fuel Transfer System (EST)
148900.64
Total Spacecraft
IA&T
29396.41
Program Level Cost
28320.65
Aerospace Ground Equipment
16656.73
Total OOR Non-Recurring Cost
Table C-1
87
223198.1
88
LIST OF REFERENCES
1.
Ridolfi, D., Satellite Refueling: The Tactical Advantage, SWC Space Tactics
Bulletin, 1995.
2.
Aviation Week and Space Technology: December 17, 1984.
3. Aviation Week and Space Technology: March 26, 1984.
4.
Aviation Week and Space Technology: April16, 1984.
5. Aviation Week and Space Technology: August 19, 1985.
6.
Aviation Week and Space Technology: September 9, 1985.
7.
Wilson, A. Interavia Space Directory (1992-3 Edition): New York 1994
8.
Aviation Week and Space Technology: March 15, 1993.
9.
Aviation Week and Space Technology: December 17, 1984.
10. Jenkins, D.R., Space Shuttle: The History of Developing the National Space
Transportation System, Motorbooks International, 1992.
11. Scholl, S.M., Optical Processing for Semiautonomous Terminal Navigation and
Docking, Applied Optics, September 1993.
12. Chi-Chang J.H./McClamroch H.N., Autonomous Spacecraft Docking using a
Computer Vision System, Proceedings of the 31st Conference on Decision and
Control, December 1992.
13. Aviation Week and Space Technology: September 13, 1993.
14. Clark, M.M./Hanes, D.E./Mauceri, A.J., Engineering, Construction, and Operations
in Space Ill; Proceedings of the 3rd International Conference, May 1992.
89
15.
DiStefano, E./Cady, B.C./Rangel, R.H., Methods of Filling Screen Liquid
Acquisition Devices in Low Gravity, Journal of Spacecraft and Rockets, November
1994.
16.
Dominick, Sam M.; Tegart, James R., Orbital Test Results of a Vaned Liquid
Acquisition Device: 30th Joint Propulsion Conference and Exhibit, June 1994.
17.
PHONCON LCDR Sassman SPAWAR/UFO Program: August 15, 1996.
18.
National Space Development Agency of Japan: Proceedings of the 34th SICE,
Hokkaido, JPN, July 1995.
19.
Isakowitz, S.J., International Reference Guide to Space Launch Systems (1991
Edition), AliA, Washington, D.C., 1991.
20.
Air Force Material Command, Unmanned Space Vehicle Cost Model (Seventh
Edition), August 1994, Los Angeles AFB, CA, 1994.
21.
Larson, W.J./Wertz, J.R., Space Mission Analysis and Design (Second Edition),
Microcosm Inc./Kluwer Academic Publishers, 1995.
22.
Jane's Space and Launch Vehicles (1993-1996 Editions), London, 1996.
90
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1.
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Ft. Belvoir, VA 22060-6218
2.
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Naval Postgraduate School
411 Dyer Rd.
Monterey, CA 93943-5101
2.
Dr Dan C. Boger .................................................................................................... 2
Naval Postgraduate School
411 Dyer Rd.
Monterey, CA 93943-5101
4.
Dr I. M. Ross ......................................................................................................... 3
Naval Postgraduate School
411 Dyer Rd.
Monterey, CA 93943-5101
5.
CDR. R. L. Hibbard ................................................................................................2
6305 Sweetbriar Dr.
Fredricksburg, VA 22407
6.
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SWC/Dot
FalconAFB, Colorado Springs, CO 80912-7383
91
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