TRIANA CONCEPT DESIGN STUDY
AA 4831 SPACECRAFT SYSTEMS II
DR. I. MICHAEL ROSS
DESIGN TEAM
LT Rick Behrens
LT Christian Dunbar
LT Michael Larios
Capt Darin Powers
10 JUNE 1998
2
TABLE OF CONTENTS
I.
INTRODUCTION ..................................................................................................................... 7
II.
MISSION DESIGN .................................................................................................................. 9
A.
LAUNCH VEHICLE .................................................................................................................. 9
1. Launch Vehicle Selection .................................................................................................. 9
2. Launch Vehicle Integration .............................................................................................. 10
B. ORBITOLOGY....................................................................................................................... 11
1. Initial Conditions .............................................................................................................. 11
2. L1 Transfer Trajectory ..................................................................................................... 11
3. Halo Orbit ......................................................................................................................... 12
V REQUIREMENTS ........................................................................................................................ 13
1. First Estimate ................................................................................................................... 13
2. SWINGBY Estimate ......................................................................................................... 14
V Budget ................................................................................................................................ 15
4. Propellant Budget ............................................................................................................ 15
5. Kick Motor Sizing ............................................................................................................. 15
D. MISSION TIMELINE ............................................................................................................... 16
III.
SPACECRAFT DESCRIPTION............................................................................................. 19
A.
1.
2.
3.
4.
5.
6.
7.
8.
9.
B.
1.
2.
3.
C.
1.
2.
3.
4.
5.
STRUCTURE AND MECHANISMS SUBSYSTEM ......................................................................... 19
Initial Design .................................................................................................................... 19
Design Approach ............................................................................................................. 19
Requirements .................................................................................................................. 20
Primary Load Bearing Structure ...................................................................................... 21
Main Payload Deck & Bay Dividers ................................................................................. 22
Outer Drum ...................................................................................................................... 22
Mechanisms..................................................................................................................... 23
Equipment Placement ..................................................................................................... 23
Structure Weight Estimate ............................................................................................... 23
THERMAL CONTROL SUBSYSTEM ......................................................................................... 24
Requirements .................................................................................................................. 24
Final Design ..................................................................................................................... 25
Design Approach ............................................................................................................. 25
ELECTRICAL POWER SUBSYSTEM......................................................................................... 29
Requirements .................................................................................................................. 29
Final Design ..................................................................................................................... 29
Design Approach ............................................................................................................. 30
EPS Purpose ................................................................................................................... 30
Power Sources ................................................................................................................ 30
a) Phase 1: Launch and Transit ..................................................................................................... 31
b) Phase 2: On-orbit ....................................................................................................................... 34
c) Power Distribution, Regulation and Control ............................................................................... 40
D.
COMMUNICATIONS SUBSYSTEM............................................................................................ 41
Requirements .................................................................................................................. 41
Final Design ..................................................................................................................... 41
Evolution of Design .......................................................................................................... 42
Link Analysis .................................................................................................................... 43
E. PROPULSION SUBSYSTEM ........................................................................................... 48
1. Requirements .................................................................................................................. 48
2. Components .................................................................................................................... 48
3. Subsystem Operation ...................................................................................................... 51
F.
ATTITUDE CONTROL SUBSYSTEM ......................................................................................... 52
1.
2.
3.
4.
3
1. Requirements .................................................................................................................. 52
2. Final Design ..................................................................................................................... 52
3. Design Approach ............................................................................................................. 53
a)
b)
c)
d)
e)
G.
H.
IV.
Normal Operation ...................................................................................................................... 53
Total Moment of Inertia ............................................................................................................. 53
Characterizing Roll Motion. ....................................................................................................... 54
Momentum and Torque Requirements. ..................................................................................... 54
Reaction Wheel Size ................................................................................................................. 56
COMMAND AND DATA HANDLING SUBSYSTEM ....................................................................... 56
FLIGHT AND MISSION SOFTWARE ......................................................................................... 57
1. Final Design ..................................................................................................................... 57
2. Requirements .................................................................................................................. 57
3. Evolution of Design .......................................................................................................... 57
MISSION PAYLOADS ........................................................................................................... 59
A.
EARTH SENSOR................................................................................................................... 59
1. Requirements .................................................................................................................. 59
2. Final Design ..................................................................................................................... 59
3. Design Approach ............................................................................................................. 61
B. ALTERNATE SENSORS ......................................................................................................... 64
1. Requirements .................................................................................................................. 64
2. Recommendations ........................................................................................................... 64
a) Geocorona ................................................................................................................................ 64
b) Solar Wind................................................................................................................................. 64
V.
COST ..................................................................................................................................... 65
A.
B.
C.
REQUIREMENT .................................................................................................................... 65
ESTIMATED TOTAL COST ..................................................................................................... 65
COSTING APPROACH ........................................................................................................... 66
a)
b)
c)
d)
e)
f)
VI.
Parametric Cost Model Description ........................................................................................... 66
The Triana Cost Estimation ....................................................................................................... 67
Space Segment Costs............................................................................................................... 68
Launch Segment Costs ............................................................................................................. 70
Ground Segment Costs ............................................................................................................. 71
Operations and Support Costs .................................................................................................. 71
APPENDIX ............................................................................................................................ 73
A. SPACE SHUTTLE PRICING ALGORITHM .................................................................................. 73
V BUDGET ................................................................................................................................... 74
C. SWINGBY MISSION FILE .................................................................................................... 76
D. PROPELLANT MASS ESTIMATION .......................................................................................... 77
E. KICK MOTOR COMPARISON .................................................................................................. 78
F.
STRENGTH AND RIGIDITY ESTIMATIONS ................................................................................ 79
G. STRUCTURE SUBSYSTEM W EIGHT ESTIMATE ........................................................................ 82
H. PROPELLANT TANK SIZING ................................................................................................... 84
VII. REFERENCES ...................................................................................................................... 85
4
ACKNOWLEDGEMENTS
The Naval Postgraduate School TRIANA Design Team would like to thank the following
individuals and corporations for their willing and energetic assistance.
Ball Aerospace & Technologies Corp
Aerospace Systems Division
Lockheed Martin
Athena Government Marketing
Mr. William P. Carter, Director
Lockheed Martin
NASA Civil Space Programs
Mr. Tim Malloney
NASA Goddard Space Flight Center
Mr. Jay Smith
Ms. Susie Smith
Mr. Tom Stengle
Mr. Dave Weidow
Satellite Power Corporation
Swales Composite Structures, Inc.
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6
I.
INTRODUCTION
This design project was completed as part of the AA 4831, Spacecraft Systems II, course
at the Naval Postgraduate School, which is one if the capstone courses in the Space Systems
Operations curriculum. The stated goal of the course is to complete the systems analysis and
design of a generic spacecraft, and thereafter apply the "tools of the trade" to a selective project.
The TRIANA Earth Observing mission was selected as that project because it represents a
relevant, real-world mission that was within scope of the course.
TRIANA's prime mission is to provide "real time", HDTV color images of the sunlit disk of
the Earth for educational outreach opportunities. Images from the spacecraft will then be made
available to the world through the Internet. The TRIANA concept was initially conceived by the
Vice President of the United States, the Honorable Albert Gore. NASA had posted a Request for
Information (RFI) concerning the concepts for educational outreach and the results of a
preliminary feasibility study. This information provided the initial requirements used in the study
and bounded many of the subsequent system trades.
TRIANA Requirements include:

100 kg (dry), 175W Spacecraft

Launched to the Solar Lagrangian Point, "L-1"

Planned five year operational mission

Total mission cost less than $50M, including launch and operations

Free worldwide, 24-hr access to the updated image via the Internet

Potential Users: General Public, Educators, Students, Non-profit Groups, Weather
Forecasters
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8
II.
A.
MISSION DESIGN
Launch Vehicle
1.
Launch Vehicle Selection
Launch vehicle selection process began with a survey of currently available U.S.
expendable launch vehicles and the Space Transportation System (i.e., Space Shuttle). Four
launch vehicles were then selected for further consideration: Space Shuttle, Lockheed Martin
Athena II (formally the LLV2), Orbital Sciences Taurus XL, and McDonnell Douglas Delta II
(7925). The Space Shuttle option is very attractive considering cost as estimated in Appendix A.
However, current NASA practices discourage Shuttle missions from performing nonessential
satellite operations, other than Get-Away-Specials, when other commercial means are available.
Also, the availability of cargo bay space in the Space Shuttle during the next two years is unlikely
with the scheduled Space Station missions. The next best option is the Athena II based upon
performance margin, payload volume, and cost. The Taurus XL is a close alternative and the
Delta II was found to be much more capable than necessary and correspondingly too expensive.
Calculations for both the Space Shuttle and Athena II options were carried through V estimates
and considered for kick motor sizing.
For launch vehicle comparison, the Spacecraft Dry Weight was taken as the design goal
of 100 kg. Estimating 50 kg for propellant yielded a Loaded Spacecraft Weight of 150 kg. Initially
from the RFI and then later validated from V calculations, the use of a Morton Thiokol STAR
30C Motor (626.1 kg) was added for a total of 776.1 kg. Finally an adapter weight of 15 kg was
estimated for a final Boosted Weight of 791.1 kg.
Table II-1 Launch Vehicle Comparison
Launch Vehicle
Space Shuttle
Athena II
Taurus XL
Delta II (7925)
Payload Volume
Dia / Height
4.6m / 18.3m
1.98m / 4.64m
1.37m / 3.3m
2.8m / 4.2m
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Cost
Payload to LEO
Margin
$10M
$24M
$26M
$45M
23090 kg
1300 kg
1389 kg
5045 kg
96%
34%
43%
84%
2.
Launch Vehicle Integration
For the Space Shuttle option, in addition to the spacecraft and kick motor, consideration
of the Space Shuttle’s cargo bay must be taken into account. A system similar to the STS PAMD system depicted below must be used consisting of a cargo bay cradle assembly, spin table and
separation system, and a special payload attach fitting (PAF) that attaches to the forward
restraints of the cradle assembly.
Figure II-1 Launch Vehicle Integration
For the Athena II option, the Model 92 Payload Fairing (1.98m x 4.64m) was estimated to
provide adequate volume for the spacecraft and kick motor. The standard Model 24 Payload
Adapter (attached to the Model 66 Payload Adapter) with a 591mm diameter bolt circle was
chosen for the STAR 30C kick motor interface. An option to consider is to have Lockheed Martin
provide the STAR 30C kick motor and interface as part of the Athena II contract. This would
increase launch costs by about $3M (cost of the kick motor alone is about $1.5M) but would
reduce design load of the spacecraft to just the spacecraft-kick motor interface.
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B.
Orbitology
1.
Initial Conditions
The initial orbital parameters depend on the launch vehicle and date of launch. For
example, the inclination relative to the ecliptic plane is directly related to the time of year and
must be accounted for in the transfer trajectory design. For the Space Shuttle case, a typical
parking orbit is taken to be a 28.5 inclined, circular orbit at 300 km altitude. The estimation for
the Athena II provided a 28.5 inclined, circular orbit at 1200 km altitude. Considering the Athena
II’s launch margin, it may be possible to use “dog legs” to make a plane change at LEO to save
energy for the kick motor. In either case, precise timing is required to ensure proper phasing of
the transfer orbit. An additional consideration is the benefit of performing a swingby of the Moon
to pick up additional energy or adjust orbital planes. A lunar swingby would require additional
phasing constraints that would also limit launch windows.
2.
L1 Transfer Trajectory
The solar Lagrangian point L1 lies on the Earth-Sun line where the forces of gravity from
the two bodies are nearly equal, approximately 1.5M km from Earth (about four times the distance
to the Moon). In Keplarian, two-body motion, a body’s angular velocity is greater the shorter the
radius of the orbit. However, the L1 point orbits the Sun with the Earth, so it’s angular velocity is
actually less than the Earth’s rather than greater. Restricted three-body approximations must be
used to design the transfer trajectory to take this and other unique properties of this point into
account. The actual interplanetary trajectory TRIANA will use to get to solar L1 will be the same
as that used by the Solar and Heliospheric Observatory (SOHO) spacecraft in July 1995.
The transfer trajectory will begin with a large tangential burn of the STAR 30C kick motor
that will place the spacecraft in a hyperbolic escape trajectory relative to the Earth. Anticipate
several mid-course corrections will be necessary to correct for kick motor dispersion errors. The
actual orbit insertion maneuvers are very complex and involve both radial and normal
components to the burns as well as tangential. The transfer will take approximately 106 days to
complete.
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3.
Halo Orbit
The Earth-Sun L1 Lagrange point is actually unstable. If TRIANA were to be precisely
stationed there, any disturbance force would cause the spacecraft to steadily drift away.
Frequent small corrections would be required to maintain a stable orbit. That is why TRIANA will
follow a “halo” orbit about the L1 point just as SOHO. The circular acceleration of the halo orbit
compensates for the gravitational difference between the Sun and Earth. However, the halo orbit
is also unstable so periodic orbit maintenance is required. The SOHO spacecraft performs orbit
maintenance maneuvers every 2 to 3 months. Following the SOHO mission profile, the halo orbit
will have a period of 180 days. This allows for smooth Sun-spacecraft-Earth velocity changes
and also avoids direct conjunction with the Sun that would make receiving telemetry from the
spacecraft very difficult. Also note that the halo principally lies in a plane that is perpendicular to
the Sun-Earth line. However, the Earth’s orbit is not perfectly circular so the position of L1
actually moves along the Sun-Earth line during the course of Earth’s obit. This relates to a third
dimension to the halo orbit along the Sun-Earth line. There is much ongoing research in the area
of restricted three body problems and halo orbit trajectory design. A depiction of the SOHO orbit
schematic is shown below.
Figure II-2
SOHO Orbit Schematic
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C.
V Requirements
1.
First Estimate
The mission calls for an interplanetary trajectory from Earth to the solar Lagrangian point
L1. For the first estimate, a patched conic approach was taken with a hyperbolic escape transfer
from Earth’s sphere of influence (SOI) and a heliocentric Hohmann transfer from Earth orbit to L1
orbit.
Vse
Vse
Vsl1
Vsa
Vsp
Vs1
Vs2
= Earth velocity wrt Sun
= L1 velocity wrt Sun
= Apogee velocity of transfer
orbit wrt Sun
= Perigee velocity of transfer
orbit wrt Sun
= Hohmann transfer insertion
velocity
= L1 insertion velocity
Vs1
Vsa
L1
Vsl1
SUN
EARTH
Vsp
Vs2
Figure II-3 Heliocentric Hohmann Transfer
The Vs1 calculated for the heliocentric Hohmann transfer insertion was used as V, the
velocity relative to the Earth at the edge of the SOI (~145 Earth radii), for the hyperbolic escape
transfer calculations.
The Vs2 calculated for L1 insertion was used in the estimation of
propellant requirements.
Within the Earth’s SOI, hyperbolic escape trajectories were considered with perigee of
the hyperbola based on typical Space Shuttle parking orbits and estimated performance of the
Athena II launch vehicle. V requirements for both launch vehicle cases were used for kick motor
sizing. See Appendix B for details.
13
V
Vei
Vep
V
Ve1
Ve1
= Initial s/c velocity wrt Earth
= Hyperbola perigee velocity
wrt Earth
= s/c escape velocity wrt
Earth
= Hyperbolic escape insertion
velocity
Vep
Vei
EARTH
Figure II-4 Hyperbolic Escape Transfer
2.
SWINGBY Estimate
The V requirements were then refined from the use of SWINGBY, Release 98.01, a
mission analysis and trajectory design computer program that was provided by the Flight
Dynamics Division of NASA’s Goddard Space Flight Center. A mission file, listed in Appendix C,
was developed to estimate the V requirements for the major trajectory events such as transfer
orbit insertion, mid-course corrections, and halo orbit insertion maneuvers. Below is a depiction
of the transfer:
Figure II-5 SWINGBY Transfer
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3.
V Budget
From
the
above
calculations
and
estimates
for
stationkeeping
based
upon
correspondence with NASA’s Goddard Space Flight Center, the V budget below was calculated:
Table II-2 V Budget
Event
Transfer Orbit Insertion
(Shuttle / Athena II)
Mid-Course Corrections
Halo Orbit Insertion
Stationkeeping
(5 yr Mission Life)
4.
Patched Conic
3201 m/s / 3006 m/s
SWINGBY
3180 m/s / 2982 m/s
60 m/s
385 m/s
20 m/s
434 m/s
20 m/s
Propellant Budget
From the above V budget, the required mass of propellant necessary for the mission life
of the spacecraft is estimated. The final dry mass of the spacecraft is taken to be the design goal
of 100 kg. Assuming a simple anhydrous hydrazine (N 2H4) propulsion system with a conservative
specific impulse of 220 seconds, the required propellant mass is calculated to be 24.04 kg.
Considering the level of confidence in the V estimations, a propellant margin of 50% was
deemed necessary to then make the design propellant mass equal to approximately 50 kg. See
Appendix D for details.
5.
Kick Motor Sizing
Again, considering the typical parking orbit of the Space Shuttle and estimated
performance of the Athena II launch vehicle with the V required for the hyperbolic escape
trajectory, a comparison of the Thiokol STAR 27E, STAR 30C, and STAR 37XFP kick motors was
done.
Loaded Spacecraft Weight was estimated as before to be 150 kg with an additional
adapter weight of 15 kg was added to the corresponding weights of the motors to calculate the
effective V capability.
See Appendix E for details. The table below lists the performance
margins of the three motors in each launch vehicle case.
15
Table II-3 Kick Motor Margin Comparison
Launch Vehicle
Space Shuttle
Athena II
V Required
3201 m/s
3006 m/s
STAR 27E
-19%
-12%
STAR 30C
15%
20%
STAR 37XFP
27%
32%
As readily seen, the STAR 27E does not produce enough energy for the mission. A
performance margin of 15% in the Space Shuttle case for the STAR 30C kick motor is deemed
more than adequate for this phase of the mission and is considered to have validated earlier
assumptions and the RFI’s initial feasibility study.
D.
MISSION TIMELINE
The major phases for the TRIANA mission are the launch, transfer orbit insertion, cruise,
and halo orbit insertion. The mission begins with the launch of either the Space Shuttle or an
Athena II launch vehicle from the Eastern Test Range. The spacecraft is initially turned OFF
throughout the launch phase. In the Space Shuttle case, the spacecraft will typically be placed
into a 28.5 inclined, circular parking orbit at approximately 300 km altitude. When scheduled,
the spacecraft will be spun to 60 rpm on it’s spin table and separation springs will release to
impart 0.76 m/s velocity with respect to the Shuttle. Once clear, the STAR 30C kick motor will fire
for approximately 51 seconds to inject the spacecraft into the transfer trajectory. In the Athena II
case, launch trajectory will be tailored to achieve a 28.5 inclined, circular orbit at approximately
1200 km altitude. Following the third stage burn, the Athena II’s Orbital Assist Module (OAM) will
orient the spacecraft in the proper direction, spin it up to 60 rpm, and then separate. The STAR
30C kick motor will then fire as in the Shuttle case.
Approximately one minute after the STAR 30C burns out, the spacecraft will separate
from the kick motor. The spacecraft will then be turned ON. On-board system checks will be
conducted, inertial position determined, and telemetry established via the low gain omni-direction
antennas. Approximately 30 minutes later, the thrusters will be fired to de-spin the spacecraft
and re-orient to the cruise attitude with the aid of the reaction wheel system. The solar panels will
16
then be deployed. Expect that the first major mid-course correction will occur approximately 1.5
days later to account for launch vehicle/kick motor dispersion errors. Halo orbit insertion will
occur approximately 105 days later. While during the cruise phase, the payload and high gain
antenna will be checked out and operated.
Once established in the Halo orbit, mission operations will commence. Anticipate that
periodic orbit maintenance and momentum dumping of the reaction wheel system with the
thrusters will be necessary to account for secular disturbance torque accumulation and inherent
halo orbit instability. Design life is for 5 years, though with propellant margins and hopefully good
health of the spacecraft, a considerably prolonged actual mission life will be possible.
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III.
A.
SPACECRAFT DESCRIPTION
Structure and Mechanisms Subsystem
1.
Initial Design
The initial design for the structure subsystem consists of a cylindrical kick motor adapter
attached to a conical support structure that supports both the propellant tanks and main payload
deck. The outer drum of the spacecraft is a 1m x 1m octagonal cylinder. Four solar arrays attach
to the base of the outer drum and are supported by spars connected to the adapter-conical
support structure interface.
Graphite epoxy honeycomb sandwich is used for most of the
spacecraft structure based upon its light weight, and high strength/stiffness properties.
Figure III-1 TRIANA Cut-Away
2.
Design Approach
A survey of recent spacecraft missions was made and used to make early decisions
regarding spacecraft form.
The Lunar Prospector mission was found to be very similar to
TRIANA in scope and its spacecraft design was used as a starting point.
Also, the NRAD
Clementine mission was found to be very useful in early concept exploration. Consideration of
the estimated size of the payload telescope, required antenna diameter, and estimated volume of
propellant tanks initiated the design process. Consideration of standard kick motor interfaces and
launch vehicle payload fairing envelopes helped bound the dimensions of the thrust tube and load
bearing structures.
Multiple iterations were accomplished as more knowledge of power and
propulsion requirements became known.
Extensive use of AutoCAD enabled more precise
19
design of dimensions and configuration layout. Mass and moments of inertia of individual and
collective components were also estimated with AutoCAD.
3.
Requirements
The structure must maintain TRIANA’s critical dimensions and alignments during all
phases of operation including ground handling, test, storage, launch, and on-orbit operations.
The table below summarizes the loading and frequency environment for both launch vehicles:
Table III-1 Launch Vehicle Constraints
Launch Vehicle
Space Shuttle
Athena II
Load Envelope
Axial
/
Lateral
+6.7 g
/
+5.9 g
-2.0 to +8.0 g /
 2.5 g
Fundamental Frequency
Axial
/
Lateral
13 Hz
/
13 Hz
30 Hz
/
12 Hz
(45-70 Hz)
The payload envelope of the Model 92 Payload Fairing with the Model 24 and 66 Payload
Adapters for the Athena II was taken as the most limiting case. As depicted below, TRIANA with
the STAR 30C kick motor attached has plenty of clearance.
Figure III-2 Payload Fairing Clearance
20
Additional constraints that need to be addressed in future iterations of design besides
load factors and low frequency vibration are the spectral density of random vibration, acoustics,
shock, thermal stress, pressure venting, and contamination effects.
The material properties of composites vary depending on the fiber fraction, fiber type
selected, textile weave type, and individual material properties of the fibers and matrix materials.
Typical properties of graphite epoxy composites are listed below:
Table III-2 Typical Composite Properties
Property
Elastic Modulus (Gpa)
Tensile Strength (MPa)
Compressive Strength (MPa)
Fine Grained Graphite
10-15
40-60
110-200
Unidirectional Fibers
120-150
600-700
500-800
3-D Fibers
40-100
200-350
150-200
Initial analysis considered only compressive/tensile loads for strength and natural
frequency for rigidity. From commercial market survey of composite honeycomb panels and
specifications of recent spacecraft such as Lunar Prospector, assumed core thickness of 1.1 cm
and face sheets of 1 mm.
See Appendix F for details.
Even using the most conservative
estimates for Elastic Modulus and Strength, the required thickness of the face sheets is only
0.03554 mm. This is largely due to the small mass of the spacecraft. The real driver then is the
availability, cost, and ease of manufacture/fabrication of the designed structures.
4.
Primary Load Bearing Structure
Kick motor adapter is 14 cm height x 61 cm diameter, cylindrical tube made of 1.3 cm
thick graphite-epoxy honeycomb. An additional 1 cm lip is added for integration with the kick
motor separation system.
Load bearing structure is 48.7 cm height x 61 cm inner diameter and 100 cm outer
diameter conical tube.
Four circular cutouts are made for support of the propulsion tanks.
Material is also made of 1.3 cm thick graphite-epoxy honeycomb.
21
Figure III-3 Primary Load Bearing Structure
5.
Main Payload Deck & Bay Dividers
The main payload deck is a 100 cm diameter, circular panel that is attached to the top of
the load bearing structure and supports the majority of the internal equipment. The bay is further
divided by four 50 cm x 50 cm square panels that together with the deck provide approximately
2.785 m2 of mounting area. The deck and dividers are made of graphite-epoxy honeycomb.
Figure III-4 Main Payload Deck & Bay Dividers
6.
Outer Drum
The outer drum is a 1 meter height x 1 meter diameter, octagonal, cylindrical drum. The
base of the drum is attached to the kick motor adapter at eight points with four spars that provide
additional support to the solar arrays and thruster modules. The mid-section of the drum is
attached to the top of the load bearing structure at eight points near the center of each
rectangular panel. The top panel of the drum is a 1 meter diameter, octagonal panel with a 22 cm
diameter cutout for the payload telescope. It is attached to the drum at eight points and provides
stability for the telescope and support for the antennas and thruster modules.
22
Figure III-5 Outer Drum
7.
Mechanisms
The only mechanisms are the solar panel hinges.
Estimate that a hook latch
arrangement with redundant torsion springs could be utilized to keep the solar panels in the
stowed position. The solar panels would then be deployed with pyrotechnic release devices.
Further analysis on the loads and torsion of the hinges and joints is necessary.
8.
Equipment Placement
Initial placement of internal equipment was based roughly on balancing center of mass of
the spacecraft.
Additionally, we endeavored to mount subsystem components in the same
payload bay for ease of access during testing and integration. Additional analysis of the thermal
requirements of each component is necessary.
Figure III-6 Equipment Placement
9.
Structure Weight Estimate
To estimate weight, a figure of merit was calculated assuming 1 mm graphite-epoxy face
sheets on 11 mm honeycomb core. Approximate densities for the materials were obtained from
Swales Composite Structures of Newark, DE and are 1937.59 kg/m 3 for the face sheets and
23
32.036 kg/m3 for the honeycomb core making the figure of merit equal to 325 kg/m 3. This very
conservative estimate includes the weights of adhesive, fittings, and potting. See Appendix G for
details. The table below lists structure subsystem weights by component.
Table III-3 Structure Subsystem Weights
Component
Kick Motor Adapter
Load Bearing Structure
Main Payload Deck
Payload Bay Dividers
Outer Drum
Top Panel
Support Spars
Weight
1.27 kg
4.02 kg
3.31 kg
4.07 kg
14.18 kg
3.33 kg
0.4 kg
Total
30.58 kg
Generally, the structure subsystem is over-designed for strength and stiffness. Estimate
that the areas of greatest concern will be the stress about the fittings and attachment points.
Finite element analysis is necessary to properly determine shearing energy and failure modes.
There is great potential for weight and size reduction of the spacecraft.
B.
THERMAL CONTROL SUBSYSTEM
1.
Requirements
The most restrictive temperature ranges are listed below. Since the propellant tanks will
be controlled separately using MLI, the overall temperature bounds are from 5 to 20 °C.
Table III-4 Temperature Ranges by component.
Component
Temperature Range °C
Source
Electronics
0 to 35
SMAD p. 410
Sensor
0 to 20
Kodak
Batteries
5 to 20
SMAD p. 410
Propellant
7 to 53
SMAD p. 410
24
2.
Final Design
TRIANA will use a passive thermal control subsystem consisting both of gold multilayer
insulation and optical solar reflectors (OSR’s) on the body of the spacecraft and localized use of
multi-layer insulation (MLI) for components with higher temperature limits. Below is a plot of one
cycle of the spacecraft’s equilibrium temperature using the combination of gold MLI and OSR.
Figure III-7 TRIANA's Temperature Profile
Since TRIANA will be continuously exposed to solar radiation, the spacecraft will always
be at equilibrium temperature. For that reason, a mixture of gold MLI and quartz-over-silver
OSR’s were used for thermal control.
The propulsion end (bottom) of the spacecraft would
absorb the majority of the solar radiation while the sensor/coms end (top) would receive none.
The table below depicts the placement of the thermal control materials on the spacecraft.
Table III-5 Thermal control composition.
3.
Design Approach
Spacecraft equilibrium temperatures were determined by estimating the solar flux at L1
and applying it to the appropriate surface area of the spacecraft. Since TRIANA would circle L1
in three dimensions, it is necessary to consider how solar incidence angle varies over the period
of the halo orbit. The angle the incident energy makes with the halo itself is negligible (0.13°).
Instead, spacecraft pointing causes the largest variation in incidence angle. The below table lists
25
pointing and incidence angles for a complete halo orbit (180 days). The spacecraft’s maximum
pointing angle is 22.44° when it crosses the major axis of the ellipse.
The solar angle of
incidence (on the sides of the spacecraft) at that point is 67.43° which also accounts for the angle
the sun makes with the halo at that point in the orbit. Likewise, the minimum pointing angle
occurs when the spacecraft is at the top of the halo (positive z direction) because the halo is
inclined with its top closer toward the Sun. The minimum pointing angle is 5.31°, thus the angle
of incidence on the side of the spacecraft body is 84.56°.
Table III-6 Pointing and incidence angles for one halo orbit.
Day
0
45
90
135
180
Pointing Angle
5.31
22.44
-6.87
-22.44
5.31
AOI
84.56
67.43
83.26
67.69
84.56
Equilibrium temperature was determined for each angle of incidence after estimating
internal power dissipation to be 87 watts (50% efficiency) and solar flux at L1 to be 1418 w/m 2.
The following equation was used to calculate equilibrium temperature for the first iteration.
Equation III-1


1
4
Gs  s Acos     s A cos 
  s Acos90   BOTTO M 
TOP
SIDE
TE  

  IR ATOP   IR ASIDE   IR ABOTTO M




Where TOP, SIDE, and BOTTOM denote a position on the spacecraft over which the
material would cover. This assumed that each position would be covered by only one material;
however, this was insufficient for controlling spacecraft temperature within the specified limits.
The figure below is a temperature plot of one of the early iterations. Changing the material type
for each part of the structure caused large, discrete changes in the temperature profile. To
provide finer control over equilibrium temperature, it became necessary to redesign the model to
accommodate multiple materials on a given structural component. Then smaller adjustments
could be made to the amount of each material used.
26
Equilibrium Temperature Profile
Top: OSR; Bottom : Gold; Sides: Gold
Equilibrium Temp.
(Celcius)
50.0
40.0
30.0
20.0
10.0
0.0
0
50
100
Tim e (Days)
150
Figure III-8 Equilibrium Temperature for Single Material Faces
Equilibrium Temperature Profile
T o p :OSR ( .5) / Go ld ( .5) ; B o t t o m OSR ( .5) / Go ld ( .5) ;
Sid es:OSR ( .5) / Go ld ( .5)
30.0
Equilibrium
Temp. (Celcius)
20.0
10.0
0.0
-10.0 0
50
100
150
-20.0
-30.0
Tim e (Days)
Figure III-9 Temperature Profile for OSR/Gold Combination
Equilibrium Temperature Profile
T o p :B lk( .6 ) / W ht ( .4 ) ; B o t t o m:B lk( .6 ) / W ht ( .4 ) ; Sid es:B lk( .5) / W ht ( .5)
Equilibrium Temp.
(Celcius)
25.0
20.0
15.0
10.0
5.0
0.0
0
50
100
Tim e (Days)
150
Figure III-10 Temperature Profile for Black/White Mixture
27
Equilibrium Temperature Profile
30.0
Top:Gold(1.0); Bottom :Gold(.6)/White(.4); Sides:Black(1.0)
Equilibrium Temp.
(Celcius)
25.0
20.0
15.0
10.0
5.0
0.0
-5.0 0
50
100
150
Tim e (Days)
Figure III-11 Temperature Profile for Gold MLI/White Epoxy/Black Paint Combination
Equilibrium Temperature Profile
T o p :A lum( .6 ) / B lk( .4 ) ; B o t t o m:A lum( 1.0 ) ; Sid es:A lum( .4 5) / B lk( .55)
Equilibrium Temp.
(Celcius)
25.0
20.0
15.0
10.0
5.0
0.0
0
50
100
Tim e (Days)
150
Figure III-12 Temperature Profile for Aluminum/Black Paint Mixture
Several iterations were attempted using different materials and by changing the
combination ratios of each. Most of the combination could suit the specification by adjusting the
combination ratios.
The Gold/White/Black represents a nonhomoginous combination where
incidence angle causes large variations in temperature. An improvement was made by replacing
gold with aluminum which conforms to the temperature limits but leaves no margin for changes in
solar flux throughout the mission. The two best combinations are the black paint and white
enamel combination and the gold MLI and OSR combination. The gold/OSR combination was
chosen as the optimum because its standard deviation was 1.65°C vise 2.9°C for the black/white
combination.
28
C.
ELECTRICAL POWER SUBSYSTEM
1.
Requirements
Total spacecraft power is not to exceed 175W.
2.

Final Design
Source: Four identical GaAs solar arrays on graphite epoxy substrate.

Physical dimensions: Each solar panel is 0.355m x 1m, with a surface area of
0.355m2. Total solar panel surface area = 4*0.355 =1.42m 2

Weight: Each solar panel weighs approximately 1kg, for a total array weight of
4kg.

Power: Per solar panel BOL power = 67.30W, per solar panel EOL power =
54.88W, with total array BOL power = 269.20W, and total array EOL power =
219.50W.

Storage: A single secondary nickel cadmium (NiCd) battery consisting of 24 sealed
rechargeable cells, provides 28VDC, at a depth-of-discharge of less than 80%, a
capacity of 7.5 A-hr, to a 87W load for a period of 1.5 hours.

Physical Dimensions: Estimated to be 42 cm long, by 9 cm wide, by 11 cm high.

Weight: Estimated to weigh less than 9kg.

Commercial product used for design calculations: Sanyo Canada Inc., High
Capacity, sealed and rechargeable, CADNICA battery model number KR-7000F.
( www.sanyo.ca )

Distribution: A dual bus using redundant, shielded cabling and mechanical relays, is
used for power distribution and switching.

Regulation and Control: A +28V standard bus direct energy transfer (DET) employing
array shunt regulators.

Weight of the Power Distribution, Regulation and Control systems is estimated at
7kg, which is 7% of spacecraft design goal weight (.07*100kg) = 7kg.

Total EPS weight is 20kg.
29
3.
Design Approach
Development of the Triana EPS concept was performed iteratively. Most trades centered
on the quantity, placement and size of the solar arrays. Battery type and quantity was also a
major element of concept development. Ultimately the Team decided on the use of four solar
arrays a NiCd battery. Discussion of the considerations that produced this design are described
below.
4.
EPS Purpose
The primary function of the Electrical Power Subsystem (EPS) is to control and distribute
a predictable and continuous source of electrical power to the spacecraft’s subsystems
throughout the spacecraft’s lifetime. In order to ensure completeness of design it is useful to
subdivide the EPS into the four sub-components as indicated below.
Electrical Power
Subsystem
Power
Source
Energy
Storage
Power
Distribution
Power
Regulation &
Control
NiCd Battery
NiCd Battery
28V bus
Unregulated
Solar array
Figure III-13 Spacecraft's Power Subsystem Functional Breakdown (SMAD, p. 391)
5.
Power Sources
Design of the EPS is begun with an identification of the spacecraft power requirements.
Top-level power requirements are driven by mission profile, orbital parameters, desired mission
life, spacecraft configuration and payload power needs. It is important to further divide power
requirements in accordance with the mission profile. In the case of Triana there are two mission
30
phases each having distinct power requirements. These two phases are, launch and transit to L1, and the at L-1 (on-orbit) operational phase.
a)
Phase 1: Launch and Transit
Launch and transit power requirements are driven by the need to communicate
and control the spacecraft bus while making the trip to L-1. During the 3 month transit the primary
energy users will be the propulsion subsystem, the attitude control subsystem, and the
communications subsystem. Throughout this time the spacecraft will be operating on a
combination of battery and solar power. Initial systems startup and power required for solar array
deployment will be provided by pre-charged secondary batteries (charged prior to launch). A
detailed mission timeline is provided in Chapter I Section E Mission Timeline, a general Phase 1
timeline with associated power requirements is presented below.
Launch & PMF, then
PKM separation
Despin & Reorient
ACS
Comms
C&DH
Deploy Arrays
Total Power =
40.0W
17.5W
9.5W
10.0W
77.0W
On-orbit 175W
.25hr
.5hr
S/C Turn on & Checkout
Comms
17.5W
C&DH
9.5W
Star Trackers
20.0W
Total Power =
47.0W
Coasting, Transfer Firing, Orbit Injection
ACS
40.0W
Comms
17.5W
C&DH
9.5W
Total Power =
77.0W
Figure III-14 Phase 1 Estimated Power Requirements
The decision to use secondary batteries was based on a desire to optimize
battery utility by incorporating a level of fault response flexibility. While on-orbit at L-1 the
spacecraft will not experience eclipse, and therefore batteries are not required for spacecraft
operation. However, the flexibility provided by having secondary batteries dramatically increases
the operator’s courses of action should the solar array go into a fault mode. The down side is that
31
use of pre-charged secondary batteries may prohibit (due to safety concerns) the spacecraft from
being launched from the Shuttle.
Design of the spacecraft battery began with a comparison of nickel hydrogen
(NiH2) and NiCd. Initially the Team was attracted to NiH 2 due to its high specific energy density
(45-60W-hr/kg vs. 25-30W-hr/kg for NiCd) and its successful use of on spacecraft such as Lunar
Prospector, Clementine and IntellSat (SMAD pp. 403). Despite this benefit and given the Triana's
low power requirement the Team choose to take advantage of the numerous commercially
available rechargeable NiCd batteries. The risk of using a non-space qualified battery can be
mitigated through careful vendor selection and by verifying cell seals by x-raying each cell prior to
use. This technique has been employed on numerous low cost, small satellites to include the
Naval Postgraduate School's Petite Amateur Satellite (PANSAT).
In order to perform as accurate of an estimation as possible,
calculation of
battery capacity, battery weight and number of cells was performed using NiCd cells sold by
Sanyo Canada Inc ( www.sanyo.ca ) as examples. Given the small size and weight of the NiCd
batteries the decision was made to include a significant amount of margin in the design of the
spacecraft battery. Therefore the design parameters were to provide for 50% of full system power
(87W at 28VDC), for a duration of 1.5 hours, at a depth-of-discharge of not to exceed 80%. The
choice to use 80% DOD (trading life cycle for DOD) is validated by the fact that once on-orbit the
battery is not required (an 80% DOD limits battery cycle life to less than 1000 cycles, SMAD pp.
404). This trade is justified by the on-orbit availability of solar power and the projected limited (if
any) use of batteries at L-1.
Table III-7 Sanyo Canada Inc. High-Capacity, NiCd Rechargeable Cells
Model
KR-7000F
Voltage (V)
1.2
Capacity (A-hr)
7.5
Diameter (mm)
33.0
1) Determine the number of cells required for a 28V battery

# of cells/battery = 28V/(1.2V/cell) = 24 cells
32
Height (mm)
91.0
2) Characterize battery capacity (A-hr) over a range of powers and DODs

Battery requirements: Provide 87W for 1.5 hours at less than 80% DOD

Capacity (A-hr) = (Power * Time)/(# of cells * Cell Voltage * DOD)

As the chart below demonstrates use of the KR-7000F provides a 25% capacity margin
at 80% DOD and 87W. Where capacity margin = [1 - (5.66/7.5)] A-hr
Table III-8 Battery capacity in A-hr, as a function of power and DOD (using the Sanyo Model KR-7000F).
Required
Power
W
80
81
82
83
84
85
86
87
88
89
90
DOD
0.6
6.94
7.03
7.12
7.20
7.29
7.38
7.47
7.55
7.64
7.73
7.81
0.7
5.95
6.03
6.10
6.18
6.25
6.32
6.40
6.47
6.55
6.62
6.70
0.8
5.21
5.27
5.34
5.40
5.47
5.53
5.60
5.66
5.73
5.79
5.86
0.9
4.63
4.69
4.75
4.80
4.86
4.92
4.98
5.03
5.09
5.15
5.21
3) The resulting physical characteristics of a generic rectangular packing of the cells to form a
battery are:

Assuming 2 rows of 12 cells each and 2cm added to each dimension to account for the
container Length = 39.6+2 = 41cm, Width = 6.6+2 = 9cm, Height = 9.1+2 = 11cm.

Volume = 41*9*11 = 4059cm 3

Weight is estimated using a gram/volume factor derived from the measurement a
standard commercial NiCd D-cell.
i)
D-cell characteristics:
(1) Weight = 150g, Height = 5.8cm, Diameter = 3.2cm, Volume = 93.3cm 3
(2) Resulting gram/volume factor = 1.61g/cm 3
ii)
Estimation of Triana battery weight = 4059cm 3 * 1.61g/cm3 = 6535g = 6.5kg
iii) Addition of 30% overhead for wiring and packaging = 8.45kg (less than 9kg).
33
Battery recharging will be performed throughout the transit to L-1. Once deployed the
solar panels will provide power for both the bus and battery recharging. During transit spacecraft
attitude will be maintained so that minimum sun incident angle is acquired ( 
 0 degs).When
maneuvers are required the spacecraft will transition to battery power, execute the maneuver,
then reorient for solar radiation collection. It is important to note that with the solar panels
deployed full system power will be available. This allows controllers to perform system checks to
include verification of the mission sensor.
b)
Phase 2: On-orbit
Once in position at L-1 the spacecraft will never go into eclipse and therefore
power requirements can be met by solar arrays, allowing the batteries to be used as a backup
power source until they surpass their cycle life. Initially the Team considered collecting solar
radiation using a single array placed on the top (northern, positive z-axis) of the spacecraft. Also
at this stage of the design process the negative axis or south face of the spacecraft was going to
have a boom mounted 1-meter parabolic high gain communications antenna. When the Team
decided to body mount the high gain antenna the single solar array was divided into two panels
with one on each of the north and south faces. Finally as presented in earlier portions of this
report the Team decided to create four separate panels. Each panel is outwardly facing and body
mounted (attached at the aft end of the spacecraft) in such away that when deployed the solar
panels look like flower pedals around the spacecraft body.
Single Panel with
Single Axis (EastWest) Driver
Two Panels with
Single Axis (EastWest) Driver
Four Fixed Panels
Figure III-15 Evolution of Solar Array Design
34
In the original plan the solar array would have had a single axis, bi-directional
drive (east-west) that would maintain the solar array perpendicular to incident solar rays. A single
axis drive was chosen to increase reliability, to decrease weight and cost and it was decided that
a dual spin (north-south and east-west) drive was not required. Justification for not using a northsouth capable drive is derived from the geometry of the halo orbit about L-1. Optimal solar energy
transfer occurs when the sun incident angle is 0 degrees and maintaining a 0 degree incident
angle is a function of the satellite’s position relative to the sun. Once the Team began reviewing
the sun angle and the orbit geometry it was further decided that a single axis driver was not
required. Therefore, the final concept was to fix the solar array position relative to the spacecraft.
Using this concept the maximum sun angle cosine loss was calculated to be 8.62%. A factor
which the Team considered acceptable and did not warrant the extra complexity, cost and weight
required for a single or dual axis driver. The table below provides data representative of the
saving in weight and power by doing away with a single axis driver.
Table III-9 Single Axis Solar Array Driver (from Ball Aerospace & Technologies Corporation)
(www.ball.com/aerospace/ptmech.html )
Parameter
Drive application
Speed (RPM)
Design life (years)
Supported loads (lbs)
Step size
Input power (W)
Position accuracy (deg)
Temperature (deg C)
Length (in)
Diameter (in)
Weight (lb)
Model: EMS 221
Solar Array
0.004 - 0.04
5
94
0.15
1.4
+/- 1
-25 to +60
7.5
7.0
9.5 (4.275kg)
The following figures and calculations describe the solar array driver trade.
35
6
1.5x10 km
A view of the Triana orbit as seen from the sun.
Earth
Halo orbit plane
B
B
L-1
8
1.4849x10 km = .9926au
120000km
B
666672km
Sun
Figure III-16 Halo Orbit Parameters
Top View of Triana at western edge of halo orbit.
Earth
B = 23.96 deg
6
1.5x10 km
-1
B
6
B = tan (666672/1.5x10 )
666672km
B
Sun
Figure III-17 Calculation of Maximum Cosine Loss Using Fixed Arrays
36
Critical to the sun angle analysis is the assumption that the sun's rays are
perpendicular to all points within the halo orbit plane. This assumption is valid because (treating
the sun as a point source) the angle formed between sun and the rays extending out to the

1
666672

)  0.2572 degrees  .
western point of the halo orbit is approximately zero,  tan (
1.4849 x10 8


Therefore the incident rays on the western edge of the halo orbit are parallel to those incident on
the center of the orbit. The calculation of the cosine losses for the minor axis extremes (northern
most and southern most halo points) are analogous to the calculation of the major axis extremes.
These losses are:

Major axis extremes cosine losses = 8.62%

Minor axis extremes cosine losses = 0.3% given:

 1 120000

 tan (1.5 x10 6 )  4.57 degrees 

cos(4.57)  0 .9968 resulting in a 1- 0.9968 = 0.003 = 0.3%
The next step in the process was to size the solar arrays to meet on-orbit power
requirements. Solar array sizing was performed using the maximum power requirements (175 W)
as indicated in the RFI. Gallium Arsenide (GaAs) was chosen due to its greater efficiency over
silicon and a graphite epoxy substrate was assumed. The final design results were:

Physical dimensions: Each solar array is 0.355m x 1m, with a surface area of
0.355m2. Total solar array surface area = 4*0.355 =1.42m 2

Weight: Each solar array weighs approximately 1kg, for a total weight of 4kg.

Power: Per solar panel power BOL = 67.30W, per solar panel power EOL = 54.88W,
with total array power BOL = 269.20W, and total array power EOL = 219.50W.
Sizing was conducted by first calculating a single array area that met the power
requirements and then dividing that area into four equally sized panels. The sizing
calculations follow.
37

 PeTe Pd Td 



Xe
Xd 

Required output power of the solar array = Psa =
Td

Pe = 175, powered required during eclipse

Pd = 175W, powered required during daylight

Te = 0, time in eclipse

Td = 1, time in daylight (the spacecraft is always in daylight)

Xe , Xd “represent the efficiency of the paths from the solar arrays through the batteries to
the individual loads and the path directly form the arrays to the loads, respectively.”
(SMAD pp. 396) For first iteration calculations the Team used values X e = 0.60 and Xd =
0.80.


therefore Psa =
Pe
= 218.75W
Xd
Power at the beginning of life = PBOL = ( G s ) I d cos 

  .19
GaAs efficiency (a selected representative value)
Table III-10 Solar Array Efficiency and Cost (from Satellite Power Corporation)
Parameters
Panel Efficiency (%)
Cost

Silicon (Si)
15.1
$9-$17 per
centimeter
Gallium Arsenide (GaAs)
19
$27-$51 per centimeter
Gs = 1418 W/m2 , solar flux. Solar flux at L-1 is greater than the solar flux at the earth
(1358W/m2) due to reduced distance and lack of the magnetosphere. Derivation of the L1 solar flux follows:

L-1 is 0.9926au = 1.4849x108km from the sun.

Sun temperature = 5800K, radius = 6.98x108m, surface area = 6.12x1018m2

Sun blackbody radiation = T4 = 64.17x106W/m2

Total power of the sun = 3.9288x1026W
38


Surface area at L-1 = 2.7708x1023m2

Gs = Flux at L-1 = Total power/Surface area = 1418W/m 2

I d  .77 inherent degradation (a selected representative value, SMAD chapter 11)



therefore PBOL= 189.58 W/m2 (note using 1358W/m 2 PBOL= 181.56 W/m2)
23.96 degrees, worst case angle of incidence
Lifetime degradation = Ld  (1  Yd )

Satellitelife
where yearly degradation =Yd = .04, selected based on a degradation in LEO of 0.0275
for GaAs and the assumption that degradation at L-1 is greater than that at LEO. (SMAD
chapter 11) Radiation at L-1 is higher than LEO therefore yearly degradation should be
increased.

satellite life = 5 years

therefore Ld = .8154

Power at end of life = PEOL = PBOL Ld = 154.58 W/m2

Area of the Solar Array = Asa = Psa / PEOL = 1.42m2

Given this area PBOL=269.20W and PEOL= 219.50W

Which equates to a 54% power margin at BOL and a 25% power margin at EOL.

Division of the single array into four equal panels results in the following:

Each panel has an area of 0.355m2

Each panel has a PBOL= 67.30 W/m2 and a PEOL= 54.88 W/m2

Note: Degradations due to dividing the solar array into four separate panels where not
consider in these calculations. These inefficiencies will only slightly reduce the already
large power margins so meeting spacecraft power will not be a problem.

Total number of cells per panel was calculated using a cell size of 2cm x 4cm (an
estimate based on a standard silicon cell size).

Cell area = 0.02m * 0.04m = .0008m 2/cell

Number of cells/panel = Panel area/Area per cell = 444 cells per panel

Total number of cells = 444 * 4 = 1776 cells
39

A survey of commercially available solar arrays was conducted. The following information
is representative of the responses the Team received.
Table III-11 Solar Array Power and Weight (from Tecstar)
Parameter
Power (Watts/m2)
Power Density (Watts/kg)
GaAs/Ge
218 (at 28 deg C)
110.6 (on graphite epoxy substrate)

Panel weight = Solar Panel PBOL /Power Density = 0.61kg per panel

Total weight for all four panels = 4 * 0.61 = 2.43kg

These values were rounded up to 1kg/panel and 4kg for all four combined.
A final note concerning the solar panels is that the commercial data indicates that it
should be possible to decrease the size of the solar array given a 19% efficiency and a power per
unit area of 218 W/m 2 for a GaAs panel. If this were the case the total array size could decrease
to at least 1m2 and the weight would decrease to approximately 2kg, (assuming a 1m 2 solar array
with PBOL=218W.)
c)
Power Distribution, Regulation and Control
Given Triana’s low power requirements the Team decided that use of a standard
28V bus would be appropriate. Power distribution aboard the spacecraft is accomplished using
redundant, shielded wire and where required mechanical relays are used as power switches.
Mechanical relays are preferred due to their proven space history and due to the increased
radiation environment at L-1. The spacecraft structure will have a single power ground point.
Control of the EPS will be performed by a direct energy transfer (DET) system employing solar
array mounted shunt regulators. First iteration design values indicate that the solar array will
provide a 25% EOL power margin. The DET sub-system will ensure that unneeded power is
dissipated and therefore bus voltage will be maintained at a predetermined level. On Triana DET
shunts are positioned at each solar panel to prevent internal power dissipation and heat
generation. Initially the design Team had calculated that the ambient temperature at L-1 would be
to low for efficient operation of all spacecraft components. As a result the original design included
40
a plan to generate heat by shunting excess power through radiators placed inside on the main
deck of the spacecraft. Of course this trade comes with the potential cost of over designing the
solar array. That is use of a 1.42m 2 aggregate array provides excess power and therefore the
system as presented is over designed. Efforts to optimize the solar array would decrease the
utility of shunting power inside the spacecraft. If however heating is required internal shunting is
certainly an option. Regardless of where shunting is performed “a shunt regulated sub-system
has advantages of: fewer parts, lower mass, and higher total efficiency at EOL.” (SMAD pp.407)
Lastly the combined weight of the components comprising the power distribution, regulation and
control sub-systems was estimated at 7% of the spacecraft (goal) dry weight, or 7kg (0.07 *
100kg).
D.
COMMUNICATIONS SUBSYSTEM
1.
Requirements
The satellite carries a 1-meter X-band antenna and is capable of downlinking data at 200
Kbits/sec; it will require approximately 3 minutes to send a full image. Three 3-meter X-band
ground stations are required to receive data.
2.
Final Design
The satellite communications subsystem is comprised of:

A single 1m diameter parabolic high gain antenna mounted on the "earth pointing"
side of the spacecraft.

Two omni-directional 12cm diameter low gain antennas mounted on opposing ends
of the spacecraft.

Two less than 5kg, and less than 17.5W transponders.

One transponder is configured for X band the other for S band.

Each transponder can transmit on any of the three antennas.

The system is designed to meet the RFI on-orbit data rate of 200kbps.

Total sub-system weight is estimated at less than 17kg.
41

Motorola’s Cassini X-band transponder a commercially available item that in general
conforms to the stated design.
(www.mot.com/GSS/SSTG/SATCOM/GSSD/Products.html )
Table III-12 Communications Subgroup
Component
High Gain Antenna
Low Gain Antennas
X-band Transponder
S-band Transponder
3.
Description
1m diameter, estimated weight <5kg, mounted on earth pointing side
2, 12cm diameter, omni-directional antennas, estimated weight <1kg each
<5kg, <17.5W, Uplink 7.1GHz, Downlink 8.5GHz
<5kg, <17.5W, Uplink 2.5GHz, Downlink 2.6GHz
Evolution of Design
Using an approach similar to the design of the EPS the Team segregated the
development of the communications subsystem into Phase 1 and Phase 2. Phase 1 corresponds
to spacecraft launch and transit and Phase 2 corresponds to on-orbit operations. The driving
requirement during Phase 1 was the Team's desire to communicate with, and command the
spacecraft throughout launch and transit. This desire coupled with restrictions imposed by the
attitude control system and EPS implied the use of a low power omni-directional antenna.
Additionally given that the sensor would not be operating the Phase 1 data rate would be small
compared to the required on-orbit rate of 200kbps. It was the 200kbps requirement that drove the
Phase 2 communications sub-system design. To achieve the required data rate without a
significant increase in power the Team choose to take advantage of the earth sensor pointing
requirements and employ a high gain directional antenna.
Another significant design
consideration was the Team's desire to provide a measure of communications robustness. Three
options were discussed in response to this requirement. These options were:
1) A non-redundant system consisting of

A high gain antenna

Two omni-directional low gain antennas

A single X-band transponder
2) A physically redundant system consisting of

A high gain antenna
42

Two omni-directional low gain antennas

Two X-band transponders
3) A spectrally diverse and physically redundant system consisting of

A high gain antenna

Two omni-directional low gain antennas

An X-band transponder

An S-band transponder
Obviously the only difference between the options is the number and frequency of the
transponders. Option 1 was ruled out since it fails to provide even a basic level of redundancy.
Option 2 while providing physical redundancy fails to provide any spectral flexibility, an option that
the Team considered valuable. Therefore, although the use of S-band is not called out in the RFI
the Team considers Option 3 to be the most favorable and therefore link analysis was performed
for both S and X band. Of further critical importance was the Team's choice to give each
transponder access to any of the three antennas. This presents the user with different
employment options based on available power and ground communications assets.
Given the decision to use Option 3 the Team progressed in the design process by using
basic link equations to characterize key parameters. A search for applicable commercial items
was performed using the key parameters. Each of the steps are described below.
4.
Link Analysis
Given that the Team was developing a concept design, Phase 1 TT&C downlink data rate
requirements were estimated to be less than 10% of the required on-orbit data rate, or
(.1)*(200kbps) = 20kbps. The uplink data rate which should consist only of command data was
estimated to be 1% of the required on-orbit data rate, or 2000bps. The Team also restricted
transponder peak power requirements to 10% of the total power requirements, or (.1)*(175W) =
17.5W. Although both the S and X band transponders can transmit via either the high or low gain
antennas it is unlikely that the high gain antenna, due to pointing requirements, would be used
during launch and transit. Therefore Phase 1 calculations were performed only using the omni-
43
directional antenna. Lastly, the Team took the liberty to model the omni-directional antennas after
an S-band antenna taken from a proposed Lunar Prospector design. Basic characteristics of this
antenna are, 12cm diameter, 20cm length, 0.3kg weight and efficiency is estimated at .6 (Lunar
Prospector, p. 44).
Phase 2 link calculations follow those conducted for Phase 1, the only difference is in the
increased data rate due to the need to transmit the sensor image. Because the omni-directional
antennas are not the primary antenna for on-orbit transmission link calculations using the omnidirectional antennas for Phase 2 operations are not presented. Link calculations and associated
notes are presented in the tables below.
A search to identify commercially available transponders that could potentially meet the
requirements uncovered Motorola’s Cassini X-band transponder. Lastly it is important to note that
due to its position relative to the sun communications with Triana may not always be possible. In
general solar noise becomes prohibitive when solar conjunction is less than 7 degrees for a noisy
sun, and less than 3 to 5 degrees for a quiet sun. As shown in the EPS solar array calculations
(see accompanying figures) when Triana is on the extreme points of its major axis the sun angle
is 23.96 degrees however when the spacecraft is at the extreme points of the minor axis the sun
angle is 4.57degrees. Therefore it is probable that when at the northern and southern points of
the orbit with a noisy sun, Triana will have intermittent communications.
Table III-13 Link Equations and Units
Item
Frequency
Transmitter Power
Transmitter Power
Transmitter Line Loss
Transmit Antenna Diameter
Transmit Antenna Efficiency
Transmit Antenna Gain
Equivalent Isotropic Radiated Power
Propagation Path Length
Space Loss
Receive Antenna Diameter
Receive Antenna Efficiency
Receive Antenna Gain
Required Receiver Power
Data Rate
Source
Input parameter
Input parameter
10log(P)
Input parameter
Input parameter
Input parameter
10*log(nt*(pi*Dt*f/c)^2)
Pt+Gt-LI
Input parameter
10*log((c/4*pi*S*f)^2)
Input parameter
Input parameter
10*log(nr*(pi*Dr*f/c)^2)
ERIP+Gr-Ls
Input parameter
44
Symbol
f
Pt
Pt
Ll
Dt
nt
Gt
EIRP
S
Ls
Dr
nr
Gr
Pr
R
Units
GHz
Watts
dBW
dB
m
dB
dBW
km
dB
m
dB
dB
bps
Desired Bit Error Rate
Eb/N0 (given BPSK)
Sky Noise Temperature
Receiver Noise Temperature
System Noise Temperature
Noise Figure
Carrier-to-Noise Density Ratio
Receiver Gr/T
Input parameter
Input parameter
Input parameter
Input parameter
Tsky+Tr
10*log(1+Tr/290)
(Eb/N0)+10logR
Gr(dB) - Tsys(dB)
BER
Eb/N0
Tsky
Tr
Ts
NF
C/N0
Gr/T
dB
K
K
K
dB
dB-Hz
dB/K
Table III-14 Phase 1 Downlink From Omni-directional Antenna to Ground Stations
Item
Units
Frequency
Transmitter Power
Transmitter Power
Transmitter Line Loss
Transmit Antenna Diameter
Transmit Antenna Efficiency
Transmit Antenna Gain
GHz
Watts
dBW
dB
m
dB
Downlink S- Downlink
band
band
2.6
8.5
17.50
17.50
12.43
12.43
1.00
1.00
0.12
0.12
0.60
0.60
8.07
18.35
Eq Isotropic Radiated Power
Propagation Path Length
Space Loss
Receive Antenna Diameter
Receive Antenna Efficiency
Receive Antenna Gain
Required Receiver Power
Data Rate
Desired Bit Error Rate
Eb/N0 (given BPSK)
Sky Noise Temperature
Receiver Noise Temperature
System Noise Temperature
Noise Figure
Carrier-to-Noise Density Ratio
Receiver Gr/T
dBW
km
dB
m
dB
dB
bps
dB
K
K
K
dB
dB-Hz
dB/K
24.57
1.50E+06
-69.03
3.00
0.60
36.02
129.62
20000.00
1.00E-05
9.50
8
290
298
3.01
52.51
11.28
34.85
1.50E+06
-79.32
3.00
0.60
46.31
160.49
20000.00
1.00E-05
9.50
8
290
298
3.01
52.51
21.57
X- Notes
10% overall power
estimated
modeled after Lunar Prospector
estimated
Distance to L1
from Triana RFI
10% of 200kbps requirement
estimated tolerable error
BPSK is assumed
20deg elev angle (quiet sun)
No cooling, room temp assumed
Table III-15 Phase 1 Uplink From Ground Stations to Omni-directional Antennas
Item
Frequency
Transmitter Power
Transmitter Power
Transmitter Line Loss
Transmit Antenna Diameter
Transmit Antenna Efficiency
Transmit Antenna Gain
Eq Isotropic Radiated Power
Propagation Path Length
Units
GHz
Watts
dBW
dB
m
dB
dBW
km
Uplink
S-band
2.5
100.00
20.00
1.00
0.12
0.60
7.72
106.72
1.50E+06
45
Uplink
X-band
7.1
100.00
20.00
1.00
0.12
0.60
16.79
115.79
1.50E+06
Notes
estimate
estimated
modeled after Lunar Prospector
estimated
Distance to L1
Space Loss
Receive Antenna Diameter
Receive Antenna Efficiency
Receive Antenna Gain
Required Receiver Power
Data Rate
Desired Bit Error Rate
Eb/N0 (given BPSK)
Sky Noise Temperature
Receiver Noise Temperature
System Noise Temperature
Noise Figure
Carrier-to-Noise Density Ratio
Receiver Gr/T
dB
m
dB
dB
bps
dB
K
K
K
dB
dB-Hz
dB/K
-68.69
3.00
0.60
35.68
211.10
2000.00
1.00E-05
9.50
290
287
577
2.99
42.51
8.07
-77.76
3.00
0.60
44.75
238.30
2000.00
1.00E-05
9.50
290
287
577
2.99
42.51
17.14
from Triana RFI
1% of 200kbps requirement
estimated tolerable error
Ambient temp of earth
Ambient temp of spacecraft
Table III-16 Phase 2 Downlink From High Gain Antennas to Ground Stations
Item
Units
Frequency
Transmitter Power
Transmitter Power
Transmitter Line Loss
Transmit Antenna Diameter
Transmit Antenna Efficiency
Transmit Antenna Gain
Eq Isotropic Radiated
Power
Propagation Path Length
Space Loss
Receive Antenna Diameter
Receive Antenna Efficiency
Receive Antenna Gain
Required Receiver Power
Data Rate
Desired Bit Error Rate
Eb/N0 (given BPSK)
Sky Noise Temperature
Receiver Noise
Temperature
System Noise Temperature
Noise Figure
Carrier-to-Noise Density
Ratio
Receiver Gr/T
GHz
Watts
dBW
dB
m
dB
dBW
km
dB
m
dB
dB
bps
dB
K
K
Downlink
S-band
2.6
17.50
12.43
1.00
1.00
0.60
26.48
42.98
1.50E+06
-69.03
3.00
0.60
36.02
148.04
200000.00
1.00E-05
9.50
8
290
K
dB
dB-Hz
298
3.01
62.51
dB/K
11.28
46
Downlink
X-band
8.5
17.50
12.43
1.00
1.00
0.60
36.77
53.27
Notes
10% overall power
estimated
from Triana RFI
estimated
1.50E+06 Distance to L1
-79.32
3.00 from Triana RFI
0.60
46.31
178.91
200000.00 full data rate
1.00E-05 estimated tolerable error
9.50 BPSK is assumed
8 20deg elev angle (quiet sun)
290 No cooling, room temp
assumed
298
3.01
62.51
21.57
Table III-17 Phase 2 Uplink From Ground Stations to High Gain Antenna
Item
Units
Uplink
X-band
7.1
100.00
20.00
1.00
1.00
0.60
35.21
134.21
Notes
GHz
Watts
dBW
dB
m
dB
dBW
Uplink
S-band
2.5
100.00
20.00
1.00
1.00
0.60
26.14
125.14
Frequency
Transmitter Power
Transmitter Power
Transmitter Line Loss
Transmit Antenna Diameter
Transmit Antenna Efficiency
Transmit Antenna Gain
Eq Isotropic Radiated
Power
Propagation Path Length
Space Loss
Receive Antenna Diameter
Receive Antenna Efficiency
Receive Antenna Gain
Required Receiver Power
Data Rate
km
dB
m
dB
dB
bps
1.50E+06
-68.69
3.00
0.60
35.68
229.52
2000.00
1.50E+06
-77.76
3.00
0.60
44.75
256.72
2000.00
Distance to L1
Desired Bit Error Rate
Eb/N0 (given BPSK)
Sky Noise Temperature
Receiver Noise
Temperature
System Noise Temperature
Noise Figure
Carrier-to-Noise Density
Ratio
Receiver Gr/T
dB
K
K
1.00E-05
9.50
290
287
1.00E-05
9.50
290
287
K
dB
dB-Hz
577
2.99
42.51
577
2.99
42.51
dB/K
8.07
17.14
47
estimate
estimated
from Triana RFI
estimated
from Triana RFI
1% of 200kbps
requirement
estimated tolerable error
Ambient temp of earth
Ambient temp of
spacecraft
E.
PROPULSION SUBSYSTEM
1.
Requirements
The propulsion subsystem provides the spacecraft the necessary V for midcourse
corrections, halo orbit insertion and maintenance, and momentum dumping of the reaction wheel
system for the mission lifetime with substantial design margin. Several different systems were
evaluated including electric and bipropellant systems. However, due to the weight, power, and
cost constraints of the mission, a simple monopropellant hydrazine system was selected. As
calculated in the V section, a propellant mass of 50 kg is estimated for the total V budget
including a 50% margin.
2.
Components
ENGINES: The engines have to provide enough impulse to efficiently produce the
comparably larger V’s required for mid-course corrections and halo orbit insertion and the much
finer impulses of momentum dumping in all three axis. Initial design considered a combination of
multiple, small 1.12 N engines for attitude control and a single 22.24 N engine for V maneuvers.
However, analysis of stability control over estimated worst case error dispersion during prolonged
firing favored the use of larger 5.34 N (1.0 lbf) engines for both attitude control and V impulse.
Analysis also compared burn times for desired V maneuvers and minimum impulse bit.
Additional analysis would consider cost versus complexity trades, reliability issues, and failure
modes of operation.
For this design iteration, twelve Rocket Research MR-111C 1.0 lbf monopropellant
engines were selected. Specific engine characteristics are shown below.
Figure III-18 Thruster Depiction
48
Table III-18 Thruster Characteristics
MR-111C (1.0-lbf Engine)
Design Characteristics
Propellant
Thrust, Steady State
Feed Pressure
Valve Power
Weight (Engine + Valve)
Demonstrated Performance
Specific Impulse
Total Pulses
Minimum Impulse Bit
Steady State Firing
Hydrazine
1.3345 - 5.3378 N
80-400 psi
9 Watts @ 28VDC
0.33 kg
226-229 sec
5970
0.08451 Ns (20 ms)
330 sec Single Firing
14400 sec Cumulative
A simple thruster module scheme was also selected for initial design. The engines would
be mounted in four modules located on the top and base of the spacecraft structure with effective
moment arms of 50 cm in the X and Z directions. This arrangement provides attitude control in all
directions for momentum dumping and V capability in the X and Z directions. Additionally, the X
direction thrusters could be aligned slightly in the Y direction to provide some V capability,
otherwise, a rotation about the Z-axis would be necessary. The exact thruster module locations
will definitely change as better knowledge of the center of mass location and halo orbit
maintenance requirements are known.
B
X
1
2
Z
A
Y
D
3
C
Figure III-19 Thruster Module Arrangement
PROPELLANT TANKS: Design propellant mass is estimated to be 50 kg of hydrazine.
For simplicity and cost, a helium blowdown pressurization system was selected. The size of the
49
propellant tank can be estimated assuming a typical blowdown ratio of 3:1 and average hydrazine
density of 1.004 gm/cm 3. See Appendix H for details.
Survey of commercially available propellant tanks resulted in the selection of four
spherical tanks rather than a single spherical tank. Actual comparison of number of tanks and
combined weight favored the four tank arrangement and allowed for additional system reliability.
The propellant tanks selected are Pressure Systems, Inc. 80290-1 spherical
tanks. Each tank is approximately 16.36 cm in radius, made of 6Al-4V titanium, and utilizes a
rubber diaphragm pressurized by helium for positive expulsion. Additional characteristics are
shown below:
Table III-19 Propellant Tank Characteristics
Tank Characteristics
Diameter: 32.7152 cm
Material: 6AI-4V Titanium
Tank Type: Diaphragm
Operating Pressure: 396 psig
Total Volume: 17698.029 cm3
Prop Volume: 12454.168 cm3
Max. Design Wt: 2.54 kg
Min. Wall Thickness: 0.05 cm
The propellant tanks are arranged equidistantly about the spacecraft’s
supported by the main load bearing structure as depicted below.
graphite-epoxy structure securely support each tank.
Figure III-20 Propellant Tank Arrangement
50
x-axis and
Actual cut-outs from the
3.
Subsystem Operation
Key elements such as maximum propellant flow rate and operating pressure would be
calculated for selecting the proper feed lines and components such as the latch valve, cavitating
venturi, fill and drain valve, and filter. A simplified schematic of the propulsion subsystem is
depicted below.
Tank
1
Tank
2
Tank
3
P
Fill/Drain Valve
Tank
4
Pressure Transducer
Filter
Latch Valve
Solenoid
Valves
A1
A2
A3
B1
B2
B3
C1
C2
C3
D1
D2
D3
Figure III-21 Propulsion Subsystem Schematic
Operation of the propulsion subsystem is controlled by simple ON/OFF commands sent
by telemetry or calculated by the onboard flight computer. The reaction wheel system will be
responsible to orient the spacecraft in the proper direction.
Large V maneuvers would be
performed by firing all number four modules in the X direction. Assuming a large midcourse
correction of 60 m/s, approximately 7 minutes of continuous firing would be required. Since the
engine’s maximum single firing period is only 5.5 minutes, the maneuver would be broken into
two parts. For the large halo orbit insertion, a total of approximately 41 minutes of continuous
firing would be required. The nominal firing scheme is shown below:
51
Table III-20 Propulsion Subsystem Operation
Propulsion Subsystem Operation
Reorientation
X Axis
Y Axis
Z Axis
Translation
X Axis
Y Axis
Z Axis
F.
Positive
A1/B3 or C1/D3
A2/D2
A1/C3 or B1/D3
Negative
A3/B1 or C3/D1
B2/C2
A3/B1 or C3/D1
C2/D2
A1/B1/C1/D1
(A1/A3/C1/C3)
A2/B2
A3/B3/C3/D3
(B1/B3/D1/D3)
ATTITUDE CONTROL SUBSYSTEM
1.
Requirements
Table III-21 Attitude control requirements.
Disturbances
Torque (Nm)
Momentum (Nms)
Solar
Insertion Thrust
1.87E-07
0.013
1.451
4.404
Pointing Accuracy
2.91 mrad
Pointing
x
y
z
2.
Range
0°
Tolerance
±0.25°
Rate
0•
45°
9°
±0.25°
±0.25°
.5°/day
0.1°/day
Final Design
Table III-22 ADCS, www.ball.com .
Star
Tracker
Quantit
y
Accurac
y
Sun
Sensor
Quantit
y
Accurac
y
Reaction
Wheels
Quantit
y
Max
Momentum
2
10 arc
sec
1
30 arc
sec
4
8
Nms
52
Prior to kick motor firing, TRIANA would be spin-stabilized by the perigee kick motor.
During transfer, the spacecraft would be despun to deploy the solar arrays and pointed using the
thrusters. The reaction wheels would make small attitude adjustments and provide momentum
storage.
Since mid-course corrections and orbit insertion maneuvers will require multiple longduration burns of two attitude thrusters, angular momentum caused by the center-of-gravity offset
and thruster misalignment must be stored by the reaction wheels. The torque created during
these maneuvers is sufficiently large to drive the sizing requirements for the reaction wheels.
Since the total momentum during orbit insertion is too great to be stored over the entire
maneuver, momentum must be dumped after each burn.
Once on station, attitude control would be provided by the reaction wheels. A sun sensor
and orthogonal star tracker would provide attitude determination.
3.
Design Approach
a)
Normal Operation
Since it is necessary to continually point the telescope toward Earth, the pointing
angle will change in a cyclic manner. Unbiased reaction wheels are best suited to this type of
attitude control.
Roll (z-axis) and yaw (y-axis) requirements were based on the change in
pointing angle over one half the period of the halo “orbit”. In that 90 day period, the spacecraft
must roll 8.58° and yaw 23.33°. Only minor pitch adjustments would be required to maintain
spacecraft attitude.
b)
Total Moment of Inertia
Moment of inertia about each axis was estimated by assuming the body to be a
uniform cylinder of radius 0.521 m. Flat, continuous plates were assumed for the solar arrays.
Mass was distributed between the body and solar arrays based on NASA data (64 kg for the body
and 36 kg for the arrays).
53
Table III-23 Moments of Inertia.
Body
8.686
Arrays
97.882
Is y 9.676
Is z 9.676
14.130
Is x
c)
14.130
Total
2
106.568 kgm
2
23.806 kgm
2
23.806 kgm
Characterizing Roll Motion.
Roll and yaw motion were characterized to determined the torque requirements
during normal operations when the telescope must be pointed toward Earth.
The following
example is applied to roll motion. Table III-22 summarizes the remaining results.

120,000 km
Sun
x
Earth
y
z
Pointing Angle Deviation
 2sin 1
z - axis displacement 


altitude
120,000 km 
 2sin 1 
1.6  106 km 
 8.58
Mean Rate 
d)
8.58
 0.10 / day
90 days
Momentum and Torque Requirements.
Initially, consideration was given to external disturbance torques as the driver for
reaction wheel selection. Gravity gradient and aerodynamics are neglected at L1, so the only
source of toque come from the sun. Table III-24 characterizes solar disturbances. The net
disturbance torque in this table has been doubled to account for torque caused by the solar wind.
Even so, the torque and stored momentum are very small.
54
Further design iterations revealed that the largest disturbance will come from the
attitude thrusters during the mid-course and insertion maneuvers.
The propulsion design
considered three thruster sizes. Table III-25 depicts torque and momentum requirements for the
three thrusters. The momentum values in this table are based on the duration of burn for each
thruster to produce the required delta-v (for insertion) in nine burns. The smallest thruster
introduces the least torque and momentum; however, they must be fired 26 minutes per burn to
produce the same delta-v in nine burns.
Therefore, the 0.5 and 1.0 lbf thrusters are more
suitable, and the maximum momentum value was chosen as the momentum requirement for the
reaction wheels.
Table III-24 On-station disturbance torque (solar).
Design Factors
Solar Array (North)
Lever Arm
Reflectivity
Surface Area
Solar Array (South)
Lever Arm
Reflectivity
Surface Area
1.03 m
0.6
2
0.42 m
1.00 m
0.6
2
0.42 m
Solar Force (North)
Solar Force (South)
3.10912E-06
3.10912E-06
N
N
Solar Torque (North)
Solar Torque (South)
3.20239E-06
3.10912E-06
Nm
Nm
Net Disturbance Torque
Stored Momentum over 90 days
55
1.87E-07 Nm
1.45 Nms
Table III-25 Disturbance torques caused by various thrusters during insertion and mid-course maneuvers.
Thrust
Cg Offset
Trust Difference
Insertion Torque
Stored Momentum over
Insertion Time
e)
.2 lbf
0.890
3
0.089
0.002669
.5 lbf
2.224
3
0.222
0.006672
1 lbf
4.448
3
0.445
0.013345
N
cm
N
Nm
4.164
5.805
4.404
Nms
Reaction Wheel Size
A 38% margin was added to the momentum requirements from Table III-25.
Final reaction wheel size is 8 Nms.
G.
COMMAND AND DATA HANDLING SUBSYSTEM
The command and data handling subsystem (C&DH) “provides stored and ground
command capability, science and engineering data formatting and packetizing, and controls all
spacecraft subsystems.” (Lunar Prospector pp.53) In general C&DH weight and size is
proportional to spacecraft complexity. Due to the simplicity of the Triana design the Team projects
that a single-unit system that performs both telemetry and command functions is appropriate.
Using the size, weight and power parametric estimation process described in SMAD (pp.390) the
Triana C&DH system has the following parameters:

Size: 4250cm3

Weight: 4.125kg

Power: 9.5W
Details required for a more robust description of the C&DH system are dependent on the
refinement of the other spacecraft sub-systems. Therefore no further specification of the C&DH
system will be provide for this first iteration design.
56
H.
FLIGHT AND MISSION SOFTWARE
1.

Final Design
Limited capability on-board computer, responsible for sub-system commanding (flight
software).

Processing of sensor data will be performed by ground based computers located at
the mission operations centers.

Injection of sensor information onto the Internet will be conducted by the ground
station computers.
2.
Requirements
Control the spacecraft sub-systems and format sensor data for transmission and
processing.
3.
Evolution of Design
The decision to limit the capability of the spacecraft computer to sub-system controlling
(flight software) was primarily made to increase reliability and decrease cost. Furthermore,
manipulation of the sensor data is greatly enhanced by performing data processing at the ground
stations. Unlike orbiting spacecraft, ground stations have the significant advantage of being able
to upgrade hardware, and software upgrades are not limited to the amount of available processor
memory, (you just add more as you need it).
The primary function of the flight software is to react to and "distribute" the commands
sent to the spacecraft by the ground controllers. The flight software also conducts numerous
health and welfare calculations using data provided by spacecraft sub-systems. Examples of
these tasks include,

Attitude and spin rate computations

Ephemeris computations

Electrical power sub-system monitoring (battery charge, power output)

Other monitoring functions associated with spacecraft health and welfare.
57
The results of these calculations are transmitted to the ground control center by the C&DH
system.
Mission sensor data is not processed by the spacecraft computer. Instead the "raw"
data is transmitted to the ground control stations for local processing. Although the Team does
not define the information transfer protocols it is projected that the raw data will be sent to a onboard buffer, and transmitted as continuous stream to the three ground stations equally spaced
around the globe.
In an attempt to further characterize the flight and mission software the Team chose
to use ADA as the programming language. ADA was chosen due to NASA's use of the language
on other spacecraft development efforts. (Lunar Prospector pp. 56) Prior use of and familiarity
with ADA increases the potential of software reuse, (ADA reuse was performed on the Lunar
Prospector Program). Without having conducted a detailed analysis it is estimated that Triana
software will consist of approximately 100,000 lines of code (LOC) (per conversation with L.
Scaglione NASA Chair, NPS). It is predicted that the Ground Software would comprise 75% of
the total (75,000 LOC) , and the Flight Software would comprise 25% (25,000 LOC).
58
IV.
A.
MISSION PAYLOADS
EARTH SENSOR
1.
Requirements
TRIANA’s primary mission is to provide recent images of the sun-lit side of the Earth at
regular intervals displayed on the Internet. The requirements for this mission are simple. The
image must be “high definition” quality at a resolution of approximately 7 km. It must also be full
color (24-bit red, green and blue). Finally, a new image of the Earth must appear on the Internet
“every few minutes”. Intitially, this was taken to mean approximately three minutes, but this is a
surprisingly short interval since very little will change in a 7 km image in that interval, but it would
be achievable under current bandwidth constraints.
The refresh rates appear to be flexible, so
the final design reflects a more reasonable rate.
2.
Final Design
TRIANA will employ three high-resolution charged coupled devices (CCD’s) in a redgreen-blue (RGB) color configuration. The optics will consist of a 22 cm schmidt-casegrain
telescope with a 2.74 meter focal length. This will allow the spacecraft to achieve its primary
mission of acquiring 7 km (30 arc minute) resolution Earth images.
Table IV-1 Earth sensor specifications.
Optics
Aperture
Focal Length
F-number
GSD
IFOV
FOV
0.22
2.74
12.3
7
4.37E-06
0.00896
Sensor
Pixels
Pixel Size
Pixel Depth
Sepectral Response
2048 x 2048
9.0
8
0.35 - 1.00
Weight
Size
Length
Diameter
Power
m
m
km
rad
rad
pixels
µm
bit
µm
20 kg
1.01 m
0.22 m
24 W
59
Space-qualified cameras of similar specifications have already been designed and
implemented in other space missions. To provide a simple, inexpensive design, the telescope
would be similar to that of the narrow-angle sensor for the Mars Orbital Camera currently flying on
the Mars Global Surveyor (MGS). The figure below illustrates the telescope design for the MGS
mission. The aperture is twice that needed for TRIANA (40 cm), but it is conceivable to use an
identical, scaled design for this mission.
In order to accommodate the three CDD arrays,
however, dichroic beam splitters must be placed at the end of the optical path to direct the
appropriate spectral range to the proper sensor.
Figure IV-1 The Mars Orbital Camera for the Mars Global Surveyor Mission,
www.msss.com/mars/global_surveyor/mgs.html
The 2048 X 2048 24-bit image has a raw file size of 12.6 MB.
No compression is
required to transmit an image of that size every three minutes over the 180 Kbps bandwidth
allotted by the communications payload. It is important to note, however, that the disk of the
Earth captured by the sensor occupies only 60% of the square image (2.6M pixels). Therefore, it
is possible to provide some processing that would transmit only pixels illuminated by the disk of
the Earth, but even this is not necessary to achieve the over-conservative requirement of 3
minutes per image. It is also more likely that a refresh interval of 10 to 15 minutes is better suited
60
to internet viewing because terrestrial activity at this resolution is not noticeable in less than 10
minutes. Extending the interval also prevents unnecessary data exchange on the Internet.
Figure IV-2 The Mars Orbital Camera prior to vibration testing
3.
Design Approach
The TRIANA telescope would employ schmidt-casegrain reflective optics with a 0.7° field
of view (FOV) similar to that of the Mars Orbital Camera. The 22 centimeter aperture would
provide a resolution of 7 km. Optical specifications were determined by the following equations
(SMAD, p.145).
Equation IV-1 Instantaneous Field of View
GSD
h
WhereGSD is the ground sample distance (resolution)
IFOV   
andh is the distance from the Earth to the sensor
 10
(1.6
km).
6
61
Equation IV-2 Field of View
FOV  N
Where N is the number of pixels in one dimension on a sensor
and is the instantaneous field of view (IFOV).
Equation IV-3 Telescope Aperture

D  1.22 
 
Where is the center wavelength and
 is the instantaneous field of view (IFOV).
Equation IV-4 Telescope Focal Length
dh
GSD
Where d is the photosite dimension (one pixel),
f 
h is the distance from the Earth to the sensor
 10
(1.6
km),
6
and GSD is the ground sample distance (resolution).
Equation IV-5 F-Number
f
D
Wheref is the focal length and
D is the aperture.
F
Two designs were considered for the focal plane. The first was a three-plane design
where dichroic beam splitters would separate the image into red, green, and blue bands – each
band collected by a separate 2048 x 2048 CCD array. The second alternative was a single plane
design where the CCD would have the red, green, and blue bands integrated together on one
array through microfiltration techniques. The second design would be simpler and could provide
savings in cost, weight and power consumption provided a space-qualified product of the required
size is currently available. The single plane design is the preferred choice, however, a spacequalified chip is not available at this time. The single-chip technology is patented by Eastman
Kodak Company and used in commercial digital camera production.
become available, it must be three times
If a single-chip should
the size of the three-chip design to match the
62
resolution. The savings comes in less redundancy in optics and processing since the three-chip
design requires beam-spitting optics and processing for each chip.
In either case, the image would be eight bits per channel, and the overall image size
would be 12.6 MB per image - uncompressed. The stated requirement for transmission is to
update the image of the Earth on the Internet every 3 minutes. It would take 70 seconds to
transmit uncompressed image over the 180 Kbps bandwidth allotted by the communications
payload. Allowing for propagation time, 105 seconds would be available for ground processing
and dissemination. This may not be enough time to manage a 12.6 MB image file. It is unlikely
that the casual user will be able to view a new 12.6 MB file every three minutes (nor should they
want to). Two options are available to fix this problem.
First, most users (including most educational institutions) have little need for a new image
of the Earth every three minutes. An object on the Earth would have to travel 140 km/hr (86 mph)
to move one adjacent pixel in that three minute interval. The object would also have to be 14 km
in size just to notice it. It is more likely that a refresh interval of 10 to 15 minutes is better suited
to Internet viewing.
2048 pi els x
Ea th ril l min
u tesa
2.6 mill i opix nls.e
4.2 mill i opix nls e
are u n in
u s e
e ryv imag
e
e .
d
Figure IV-3 A typical image of the Earth from TRIANA
Second, the disk of the Earth occupies only 62% of the complete 12.6 MB image.
Therefore, the user is only interested in 7.8 MB of the 12.6 MB file. It is possible to use encoding
schemes either in the spacecraft or at a ground station to transmit only the pixels illuminated by
the Earth. Both GIF and JPEG encoding schemes would be well suited to this operation.
63
Exact weight and size values were not available at this time. Weight and size of the
camera were scaled from the Mars Orbital Camera. Power consumption was based on typical
values from similar imaging systems.
B.
ALTERNATE SENSORS
1.
Requirements
In a sense, once TRIANA is on orbit is on orbit, most of its mission is already over since a
large part of the educational value will be in the design and construction of the spacecraft and
mission architecture. Placing slightly more than a video camera one million miles from the Earth
provides little additional educational value to the scientific community. To justify the expense for
science education, other scientific sensors should be considered as secondary payloads.
2.
Recommendations
a)
Geocorona
The Geocorona photometers are designed to measure the Hydrogen Lymanalpha (0.12 µm) emission of the neutral atmosphere. Two geocorona sensors exist currently.
Placing a third sensor from a different perspective (as TRIANA provides) would allow for threedimensional analysis of the Earth’s corona. It could be possible to use TRIANA’s telescope to
provide a complete UV image of the geocorona while adding little weight to the spacecraft.
b)
Solar Wind
Solar wind sensors provide indications of solar activity by measuring the flow of
charged particles in space. Most solar wind sensors currently in operation collect solar wind data
near Earth to monitor the condition of the Earth’s exoatmosphere. Having such a sensor between
the sun and the Earth could provide warning of the effects of solar anomalies. Currently the
Advanced Composition Explorer (ACE) has been providing solar wind data at L1 for less than a
year and will continue to do so during TRIANA’s lifetime. A solar wind instrument on TRIANA
could provide information supplemental to that provided by ACE.
64
V.
A.
COST
REQUIREMENT
Total mission cost less than $50 million, including launch and operations.
B.
ESTIMATED TOTAL COST
Use of a parametric model resulted in an estimated cost as presented below.
Table V-1 Total cost.
Segment
Space Segment
Launch Segment
Ground Segment
Operations & Support
Total (FY92$K)
Triana Costs
(FY92$K)
$332,540.50
$24,000.00
$30,100.00
$11,125
$397,765.50
Obviously this cost projection is not within the RFI requirement of $50M and the Team
does realize that this projection ($400M) is a significant overstatement of potential Triana costs.
Detailed refinement of this estimate is not within the scope of this concept study, however during
the conduct of the concept development the Team gained considerable insight into the actual
hardware costs of particular items. Not surprisingly the actual hardware costs of particular items
were relatively low, however the additional overhead associated with system integration often
resulted in a 100 to 300% increase in costs. This statement is not intended to degrade the
required element of system integration it is indeed a critical component of design and
manufacture, however it is intended to highlight the high costs associated with program level
overhead.
A method that could be used to reduce that overhead is to allow a University or other
educational institution to partner with NASA and industry on the actual design, construction and
launch of the spacecraft and ground operations centers. The primary trade in using this approach
is that labor costs are traded for schedule. That is students and faculty are the primary designers
and constructors of the spacecraft, and they receive technical guidance and oversight from NASA
65
and industry representatives. Given that the academic environment is focused on learning and
not production it is anticipated that this method would lengthen the Triana development
schedule., however it would also better meet the education goals set forth in the RFI. Regardless
of who is chosen as the prime developer the following cost analysis is presented as a point of
departure for future cost discussion and it is by no means intended for use in any formal design
proposal.
C.
COSTING APPROACH
During the concept development process the Triana Concept Team considered cost as a
design variable. That is while searching for and selecting components to use within the design,
the Team chose the lowest cost technologies which met the performance requirements. Given
that the goal was to produce a first iteration design the Team did not actively identify and then
conduct in-depth cost comparisons for each proposed design solution. Instead, as previously
described the Team made "cost conscious" design decisions and included cost by completing an
overall system level cost estimate.
a)
Parametric Cost Model Description
To frame the system level cost estimate the Triana Concept Team employed the
parametric cost model as described in SMAD, which is a derivative of the 5 th edition of the USAF
Unmanned Spacecraft Model. (SMAD Chapter 20) Fundamental to the employment of a
parametric cost model is the use of cost estimating relationships (CER). CERs are mathematical
functions that determine component cost from historically derived parameters. Through analysis
and case study research of previously constructed satellite systems these parameters have been
shown to correlate to the RDT&E and production of the theoretical first unit (TFU) cost. (SMAD
pp. 723) When producing multiple spacecraft a learning curve factor/parameter would be applied
to the TFU in order to account for cost decreases due to efficiencies associated with increased
66
unit production. (SMAD pp. 718) Given that Triana is a one spacecraft constellation, cost
estimation beyond the TFU was not required.
The primary advantage of parametric cost estimation is that it is relatively quick
to execute and provides a traceable and logical path for "outsiders" to follow. The ability to quickly
and accurately communicate cost relationships facilitates trade studies, design comparisons and
identifies areas that may require more detailed cost analysis. Therefore, the primary benefit of
parametric modeling is that system designers and users can receive direct answers to the
question of "Why does the system cost this much?"
Despite this significant advantage the primary downfall of parametric cost
estimation is that the range of historical data used to create the CERs bound its usefulness.
Given that CERs are derived from case studies of previously completed systems, they inherently
can not account for systems that fall outside the historical bounds. Parameters exist that can be
used to predict, based on historical evidence, the cost impact to programs that attempt to employ
emerging state of the art technologies, but these values are highly situation dependent and can
lead to reduced cost accuracy. Furthermore, since it is difficult to model the exact mix of
components and technologies cost prediction accuracy is further degraded. Despite these draw
backs, cost estimation must be performed and a method for doing so must be employed. The
Team determined that use of the parametric cost method would provide a system focused cost
value that would be easily traceable by outsiders.
b)
The Triana Cost Estimation
The overall system cost estimate used a cost breakdown structure as described
below.
67
Production
RDT&E
Space
Mission
Architecture
Space
Segment
Launch
Segment
Ground
Segment
Operations &
Support
Figure V-1 Cost breakdown structure (after SMAD pp. 718).
Costs for each of the four segments, Space, Launch, Ground, and Operations &
Support were calculated for both RDT&E and TFU production. The final overall system cost is the
sum of RDT&E and TFU production costs. It is also important to note that all cost values were
calculated in constant year 1992 dollars.
c)
Space Segment Costs
Space segment costs are sub-divided into three primary categories, hardware,
software and program level costs. The hardware category includes cost components which make
up the spacecraft itself. The CERs for these cost items are predominantly weight based. The
software category provides an estimate of RDT&E costs for coding spacecraft applications. The
program level cost are "associated with labor-intensive activities where a level of manpower is
operating over some period of performance" and includes functions such as management,
systems engineering, product assurance, and system tests. (SMAD pp. 725) The results of the
space segment cost estimation are described in the following tables.
Table V-2 Hardware RDT&E.
Cost Component
Parameter
X
(unit)
CER
(FY92$K)
68
Triana
X
Value
Triana
CER
Value
(FY92$K)
Payload
Visible
Comm
Antenna
Comm electronics
Subtotal
Spacecraft Bus
Structure/Thermal
TT&C
Attitude Control
Attitude Reaction Control
Power
Subtotal
Apogee Kick Motor
3-axis stabilized
Total (FY92$K)
Aperture dia. (m)
0+110791*X^0.56
.22
$47,452.73
Wt (kg)
Wt (kg)
20+1015*X^0.59
0+917*X^1.0
7
10
Dry Wt (kg)
Wt (kg)
Wt (kg)
Dry Wt (kg)
Dry Wt (kg)
EPS Wt X BOL Pwr (kgW)
16253+110*X^1.0
2640+416*X^.66
1955+199*X^1.0
0+3330*X^0.46
935+153*X^1.0
5303+.108*X^0.97
87.12
38.2
4
20.92
28
5384
$3219.44
$9170.00
$59,842.17
$ 25,836.20
$7,245.27
$ 2,751.00
$13,486.60
$5219.00
$5754.60
$60,292.67
Total Imp (Ns)
0+0.0156*X^1.0
48663
$759.14
$120,893.98
Table V-3 Hardware Theoretical First Unit Cost.
Cost Component
Parameter
X
(unit)
CER
(FY92$K)
Payload
Visible
Aperture dia. (m)
Comm
Antenna
Wt (kg)
Comm electronics
Wt (kg)
Subtotal
Spacecraft Bus
Dry Wt (kg)
Structure/Thermal
Wt (kg)
TT&C
Wt (kg)
Attitude Control
Dry Wt (kg)
Attitude Reaction Control Dry Wt (kg)
Power
EPS Wt X BOL Pwr (kgW)
Subtotal
Apogee Kick Motor
3-axis stabilized
Total Imp (Ns)
Launch Operations and Orbital Support
with AKM
Satellite Wet Wt (kg)
Total (FY92$K)
69
Triana
X
Value
Triana
CER
Value
(FY92$K)
0+44263*X^0.56
0.22
$18,958.20
20+230*X^0.59
0+179*X^1
7
10
$1,630.00
$1,790.00
$ 22,378.20
$5,768.49
$3,285.20
$ 688.33
$4,072.3
$1,754.06
$ 2210.40
$17,465.06
0+185*X^0.77
0+86*X
93+164*X^0.93
0+1244*X^0.39
-364+186*X^0.73
0+183*X^0.29
87.12
38.2
4
20.92
28
5384
0+0.0052*X^1
48663
64+1.44*X^1
150
$253.01
$280.00
$40,376.27
Table V-4 Space Segment Software Development Costs (RDT&E Costs only, using ADA).
CER
LOC
Flight Software
375*LOC
Ground Software 190*LOC
Total
25k
75k
Cost
(FY92$K)
$4,750
$28,125
$32,875
This software estimate appears grossly inaccurate and therefore a selected value of
$10M will be used for software development costs.
Table V-5 Program Level Costs.
Program
Level
Component
Program Management
Systems Engineering
Product Assurance
System Test & Evaluation
Total
RDT&E
Factor
%
20
40
20
20
RDT&E
TFU
Cost
Factor
(FY92$K)
%
$24,178.80 30
$48,357.59 20
$24,178.80 30
$24,178.80 20
$120,893.98
TFU
Cost
(FY92$K)
$12,112.88
$8,075.25
$12,112.88
$8,075.25
$40,376.27
Triana
Total
(FY92$K)
$36,291.68
$56,432.84
$36,291.68
$32,254.05
$161,270.25
Table V-6 Space Segment Costs.
Component
Hardware RDT&E
Hardware TFU
Software
Program Level
Total
d)
Cost (FY92$K)
$120,893.98
$40,376.27
$10,000.00
$161,270.25
$332,540.50
Launch Segment Costs
Launch costs were determined using values obtained from literature and Internet
surveys and are $24M as described in the Launch Vehicle Section of the Mission Design chapter
of this report.
70
e)
Ground Segment Costs
Ground segment costs vary considerably and are a function of such things as the
use of legacy or in-use systems and mission unique infrastructure needs. For the Triana concept
the Team chose to estimate ground segment development costs based on a "distribution of costs
between software, equipment, facilities" and program level costs. The development costs were
derived from a CER that relates software development costs to ground segment development
costs. (SMAD pp. 729)
Table V-7 Ground Segment Development Cost Model.
Ground
Station
Element
Facilities (FAC)
Equipment (EQ)
Software (SW)
Logistics
Management
Systems Engineering
Product Assurance
Integration and Test
Total
f)
Development
Cost as %
of Software Cost
18
81
100
15
18
30
15
24
Triana Cost
(FY92$K)
$
$
$
$
$
$
$
$
$
1800.00
8100.00
10000.00
1500.00
1800.00
3000.00
1500.00
2400.00
30,100.00
Development Cost
Cost Distribution
%
6
27
33
5
6
10
5
8
Operations and Support Costs
Operations and support costs were estimated using the approach set forth in
SMAD (pp. 730). This approach presents cost as a function of contractor and government
personnel labor rates (overhead included) as "well as the maintenance costs of the equipment,
software, and facilities." (SMAD pp. 729).
Table V-8 Operations and Support Cost.
CER
Maintenance
Contractor Labor
Government Labor
Total
0.1*(SW+EQ+FAC)/Year
$140K/Staff Year
$95K/Staff Year
71
# of Years Triana 5 year Cost
(FY92$K)
5
$9950.00
5
$700.00
5
$475.00
$11125.00
72
VI.
A.
APPENDIX
SPACE SHUTTLE PRICING ALGORITHM
 SpaceSystemWeight 
CostByWeight  $210 M  

 0.75  ShuttleCapability 
Space System Weight = 791.1 kg
Shuttle Capability = 23090 kg

7911
. kg

CostByWeight  $210 M  
  $9.59 M  $10 M
 0.75  23090kg 
73
V BUDGET
B.
Heliocentric Hohmann Transfer
sun = 1.327x1011 km3/s2
r1 = 1 AU = 1.495978x108 km
r2 = 0.9926 AU = 1.485x108 km
Vse
Vse
Vsl1
Vsa
Vsp
Vs1
Vs2
V se 
V sl1
= Earth velocity wrt Sun
= L1 velocity wrt Sun
= Apogee velocity of transfer
orbit wrt Sun
= Perigee velocity of transfer
orbit wrt Sun
= Hohmann transfer insertion
velocity
= L1 insertion velocity
sun
r1

1327
.
 1011 km
Vs1
Vsa
L1
Vsl1
SUN
EARTH
Vsp
Vs2
3
s2
 29.783 km s
1495978
.
 10 8 km


2


   r2  
.
 10 8 km  29.564 km s
  1485
 365.256day  86400 s
day 

Vsa 
Vsp 


3
2  1327
.
 1011 km 2  1485
.
 10 8 km
2  sun  r2
s

 29.728 km s
8
8
8
r1   r1  r2 
1495978
.
 10 km  1495978
.
 10 km  1485
.
 10 km


3
2  1327
.
 1011 km 2  1495978
.
 10 8 km
2  sun  r1
s

 29.949 km s
r2   r1  r2 
1485
.
 10 8 km  1495978
.
 10 8 km  1485
.
 10 8 km
Vs1  Vse  Vsa  29.783 km s  29.728 km s  0.055 km s
Vs 2  Vsp  Vsl1  29.949 km s  29.564 km s  0.385 km s
Hyperbolic Earth Transfer
earth = 3.986x105 km3/s2
SOI = 9.24x105 km
74
V
Vei
Vep
V
Ve1
Ve1
= Initial s/c velocity wrt Earth
= Hyperbola perigee velocity
wrt Earth
= s/c escape velocity wrt
Earth
= Hyperbolic escape insertion
velocity
Vep
Vei
EARTH
Vis-Viva Equation:
2  earth
2
 2 1
V 2        V ep 
 V 2
 r a
rep
1) Space Shuttle
rep = 6678 km (300 km altitude)
Vei = 7.725 km/s
Vep 2 

3
2  3986
.
 105 km
s2
  0.055 km 
2
s
6678km
Vep  10.926 km s
Ve1  Vep  Vei  10.926 km s  7.725 km s  3.201 km s
2) Athena II
rep = 7578 km (1200 km altitude)
Vei = 7.25 km/s
Vep 2 

2  3986
.
 105 km
7578km
Vep  10.256 km s
3
s2
  0.055 km 
2
s
Ve1  Vep  Vei  10.256 km s  7.25 km s  3.006 km s
75
C.
SWINGBY MISSION FILE
START 0: Initial State
a = 6676.93 km
e = 0.000245
i = 28.5
 = 0
 = 337.8
TA = 142.2117
x = -3340 km
y = 5106 km
z = 2715 km
Vx = -6.69 km/s
Vy = -3.41 km/s
Vz = -1.81 km/s
EVENT 1: Maneuver (Earth)
Impulsive/Thrust Vector (VBN)
Tvelocity = 3.12 km/s
Tnormal = 0
Tbinormal = 0
EVENT 2: Propagate (LunarTran model)
Dist From Body = 1300000 km; Earth
EVENT 3: LHLV Burn (Sun)
Radial V = 0.236 km/s
Tangential V = 0
Normal V = 0.05 km/s
EVENT 4: Propagate (LunarTran model)
XZ Cross: Earth-Centered Mean of J2000.0
EVENT 5: LHLV Burn (Sun)
Radial V = -0.0055 km/s
Tangential V = 0.12 km/s
Normal V = -0.025 km/s
EVENT 6: Propagate (LunarTran model)
Duration 140 days
EVENT 7: LHLV Burn (Sun)
Radial V = -0.07 km/s
Tangential V = 0.005 km/s
Normal V = 0
V Summary:
Transfer Orbit Insertion = 3120 m/s
Halo Orbit Insertion
LHLV Burn 1 = 241.24 m/s
LHLV Burn 2 = 122.7 m/s
LHLV Burn 3 = 70.2 m/s
Total = 434.14 m/s
76
D.
PROPELLANT MASS ESTIMATION
Assumptions:
mf = 100 kg (“Dry”)
N2H4 Thrusters, Isp = 220 sec
g = 9.81 m/s2
Estimated Total V = 465 m/s (mid-course, halo orbit insertion, stationkeeping)
  V  
I g
m p  m f  e  sp    1





 


465 m s


 

 220 s  9.81m 2  

s

m p  100kg   e
 1  24.04kg  50% m arg in  50kg






77
E.
KICK MOTOR COMPARISON
Table VI-1 Kick Motor Comparision
STAR 27E
STAR 30C
STAR 37XFP
Performance
Burn Time
Total Impulse
Effective Isp
Burn Time Tavg
Max Thrust
Spin Capability
33.46 s
8.662e5 Ns
287.4 s
2.7e4 N
2.864e4 N
100 rpm
51 s
1.652e6 Ns
284.6 s
3.196e4 N
3,703e4 N
40-100 rpm
66.5 s
2.537e6 Ns
289.9 s
3.727e4 N
4.205e4 N
65-90 rpm
Weight
Total Loaded
Propellant
Diameter
Length
330.76 kg
304.63 kg
0.6934 m
1.237 m
626.1 kg
586.9 kg
0.762 m
1.493 m
956.08 kg
883.68 kg
0.933 m
1.502 m
Assumptions:
ms/c = 100 kg (“Dry”) + 50 kg (propellant) = 150 kg
madapter = 15 kg
mo = ms/c + madapter + mkick motor
STAR 27E
mo = 150 kg + 15 kg + 330.76 kg = 495.76 kg
mf = 495.76 kg - 304.63 kg = 191.13 kg
 m
V  g  I sp  ln
 mf



 495.76kg 
  9.81 m 2   287.4s  ln
  2687.27 m s
s
19113
.
kg



STAR 30C
mo = 150 kg + 15 kg + 626.1 kg = 791.1 kg
mf = 791.1 kg - 586.9 kg = 204.2 kg
m
V  g   I sp  ln 
 mf



 7911
. kg 
  9.81 m 2  284.6s  ln
  3781 m s
s
204
.
2
kg



STAR 37XFP
mo = 150 kg + 15 kg + 956.08 kg = 1121.08 kg
mf = 1121.08 kg - 883.68 kg = 237.4 kg
m
V  g   I sp  ln 
 mf



 112108
. kg 
  9.81 m 2  289.9s  ln
  4414 m s
s
 237.4kg 

78
F.
STRENGTH AND RIGIDITY ESTIMATIONS
r2
r1
STRENGTH: Compression - Kick Motor Adapter
Assumptions:
Max Load = 8 g’s
Mass = 150 kg
F.S. = 1.25
cu = 110 MPa


 150kg   11772 N
s2

110MPa
 design  cu 
 88MPa
F .S .
1.25
Load
Load
11772 N

 Area 

 1.3377  10  4 m 2
6 N
Area

88  10
m2
Area
1.3377  10  4 m 2
t

 3.554  10  5 m  0.03554mm
2  r1  r2  2  0.305m  0.294m 
Load  8.0  9.81 m
STRENGTH: Tensile - Kick Motor Adapter
Assumptions:
Max Load = -2 g’s
Mass = 150 kg
F.S. = 1.25
tu = 40 MPa
79
Dimensions:
R1 =30.5 cm
R2 = 29.4 cm
Area = 2t(r1 + r2)


 150kg   2943N
s2

40 MPa
 design  tu 
 32 MPa
F. S.
125
.
Load
Load
2943N

 Area 

Area

32  10 6 N
Load  2.0  9.81 m
t
Area
2  r1  r2 

 9196
.
 10  5 m 2
m2
9196
.
 10  5 m 2
 2.443  10  5 m  0.02443mm
2  0.305m  0.294m
RIGIDITY: Fundamental Frequency - Kick Motor Adapter
r1
Dimensions:
R1 = 30.5 cm
R2 = 29.4 cm
L = 14 cm
I = t(r13 - r23)
r2
14 cm
Axial:
Assumptions:
Fnatural = 30 Hz
Mass = 150 kg
E = 12 GPa
2
Fnat  0.250 
A E
 Area 
m L
 Fnat 

  m L
 0.25
E
2
 30 Hz 
. m

  150kg   014
 0.25 
Area 
 2.52  10  5 m 2
9 N
12  10
m2
Area
2.52  10  5 m 2
t

 6.696  10  6 m  0.006696mm


2


0
.
305
m

0
.
294
m
2   r1  r2 
Lateral:
Assumptions:
Fnatural = 13 Hz
Mass = 150 kg
E = 12 GPa
80
2
Fnat  0.560 
EI
I
m  L3
 Fnat 
3

  m L
 0.56 
E
2
 13Hz 
3
. m

  150kg   014
 0.56 
I
 1848
.
 10  8 m 4
9 N
12  10
m2
I
1848
.
 10  8 m 4
t

 1987
.
 10  6 m  0.001987mm
3
3
3
3
   r1  r 2     0.305m   0.294m


81
G.
STRUCTURE SUBSYSTEM WEIGHT ESTIMATE
Thrust Tube (Adapter & Load Bearing Structure)
Mass:
5.29 kg
Volume:
0.016285.71 m3
Bounding box:
X: -50.00 -- 50.00 cm
Y: -50.00 -- 50.00 cm
Z: -64.40 -- -1.63 cm
Centroid:
X: 0.00 cm
Y: 0.01 cm
Z: -31.29 cm
Moments of inertia:
X: 1.1268 kg m2
Y: 1.1268 kg m2
Z: 0.7984 kg m2
Principal moments and X-Y-Z directions about centroid:
I: 0.6091 kg m2 along [0.76 -0.65 0.00]
J: 0.6091 kg m2 along [0.65 0.76 0.00]
K: 0.7984 kg m2 along [0.00 0.00 1.00]
Outer Drum
Mass:
Volume:
Bounding box:
14.18 kg
0.04363823 m3
X: -51.30 -- 51.30 cm
Y: -51.30 -- 51.30 cm
Z: -50.33 -- 49.67 cm
Centroid:
X: 0.00 cm
Y: 0.00 cm
Z: -0.33 cm
Moments of inertia:
X: 3.1055 kg m2
Y: 3.1055 kg m2
Z: 3.8471 kg m2
Principal moments and X-Y-Z directions about centroid:
I: 3.1054 kg m2 along [1.00 0.00 0.00]
J: 3.1054 kg m2 along [0.00 1.00 0.00]
K: 3.8471 kg m2 along [0.00 0.00 1.00]
Top Panel
Mass:
Volume:
Bounding box:
3.33 kg
0.01027538 m3
X: -49.97 -- 50.03 cm
Y: -50.02 -- 49.98 cm
Z: 48.42 -- 49.72 cm
Centroid:
X: 1.81 cm
Y: -0.02 cm
Z: 49.07 cm
Moments of inertia:
X: 1.0349 kg m2
Y: 1.0129 kg m2
Z: 0.4395 kg m2
Principal moments and X-Y-Z directions about centroid:
I: 0.2308 kg m2 along [1.00 0.00 0.00]
J: 0.2076 kg m2 along [0.00 1.00 0.00]
K: 0.4384 kg m2 along [0.00 0.00 1.00]
82
Main Deck Shelf
Mass:
Volume:
Bounding box:
3.318 kg
0.01021018 m3
X: -50.00 -- 50.00 cm
Y: -50.00 -- 50.00 cm
Z: -1.47 -- -0.17 cm
Centroid:
X: 0.00 cm
Y: 0.00 cm
Z: -0.82 cm
Moments of inertia:
X: 0.20766 kg m2
Y: 0.20766 kg m2
Z: 0.41478 kg m2
Principal moments and X-Y-Z directions about centroid:
I: 0.20744 kg m2 along [1.00 0.00 0.00]
J: 0.20744 kg m2 along [0.00 1.00 0.00]
K: 0.41478 kg m2 along [0.00 0.00 1.00]
Equipment Bay Dividers
Mass:
4.0775 kg
Volume:
0.01254628 m3
Bounding box:
X: -35.82 -- 35.81 cm
Y: -35.82 -- 35.81 cm
Z: -0.17 -- 48.40 cm
Centroid:
X: 0.00 cm
Y: 0.00 cm
Z: 24.12 cm
Moments of inertia:
X: 0.48839 kg m2
Y: 0.48839 kg m2
Z: 0.34207 kg m2
Principal moments and X-Y-Z directions about centroid:
I: 0.25119 kg m2 along [1.00 0.00 0.00]
J: 0.25119 kg m2 along [0.00 1.00 0.00]
K: 0.34207 kg m2 along [0.00 0.00 1.00]
Bottom Support Spars
Mass:
0.4048 kg
Volume:
0.00124558 m3
Bounding box:
X: -50.05 -- 49.95 cm
Y: -50.02 -- 49.98 cm
Z: -50.28 -- -48.98 cm
Centroid:
X: -0.05 cm
Y: -0.02 cm
Z: -49.63 cm
Moments of inertia:
X: 0.14524 kg m2
Y: 0.14524 kg m2
Z: 0.09105 kg m2
Principal moments and X-Y-Z directions about centroid:
I: 0.04553 kg m2 along [1.00 0.00 0.00]
J: 0.04553 kg m2 along [0.00 1.00 0.00]
K: 0.09105 kg m2 along [0.00 0.00 1.00]
83
H.
PROPELLANT TANK SIZING
Requirements:
Mass of propellant = 50 kg
Hydrazine density = 1.004 gm/cm3
Blowdown ratio = 3:1
50kg
Volume of propellant;
Vp 
Volume of tank; Vt 
2
Vp  74700cm 3
3
1004
.
g / cm 3
 49800cm 3
4
  r 3  r  2613
. cm
3
1
4
Vt  r 3  r  16.45cm
Radius of four tanks;
4
3
Radius of single tank;
Vt 
84
VII.
REFERENCES
Wiley J Larson and James R Wertz , Space Mission Analysis and Design, Second
Edition, Microcosm, Inc.
Roger L. Freeman, Telecommunication Transmission Handbook, Third Edition, John
Wiley & Sons, Inc., 1991.
John E. Prussing and Bruce A. Conway, Orbital Mechanics, Oxford University Press,
1993
LMLV Mission Planner's Guide, Lockheed Martin
PAM-D User's Requirements Document, McDonnell Douglas
Commercial Delta II Payload Planner's Guide, McDonnell Douglas
Space Rocket Motor Catalog, Morton Thiokol, Inc.
Commercial Taurus Launch System, Orbital Sciences
Hydrazine Handbook, Rocket Research Company
Robert Thurman and Patrick A. Worfolk ,"The geometry of halo orbits in the circular
restricted three-body problem" University of Minnesota, October 25, 1996.
85