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American Institute of Aeronautics and Astronautics
1
LEO-1: Development Of AUniversity Microsatellite For
Flight Testing New Technologies
Trevor C. Sorensen, Carol V. Hude, Marcelo H. Kobyashi, Eric J. Pilger, Amit K. Sanyal, Lance K. Yoneshige
Hawaii Space Flight Laboratory, University of Hawaii, Honolulu, HI, 96822
The Hawaii Space Flight Laboratory (HSFL) was established at the University of Hawaii
at Manoa to educate students and help prepare them to enter the technical workforce, and to
help establish a viable space industry that will benefit the State of Hawaii. In 2011 the first
mission, STU-1, will be launched. It includes a spacecraft being designed and built by the
HSFL called LEO-1. The primary objectives for LEO-1 mission are: (1) to demonstrate the
ability of the HSFL to design, build, and operate a small satellite in the 50-kg class as a
platform to test new technologies; and (2) support the C-band radar transponder
experiment (CRATEX) for the USAF; the Coherent Electromagnetic Radio Tomography
(CERTO) experiment for the Naval Research Laboratory; support the LEO-1 Solar Array
Experiment; and to demonstrate the suitability of the LEO-1 spacecraft to test technologies
suitable for missile detection and tracking. The radar calibration experiment uses C-band
transponders and precise orbit determination to test new technologies designed to aid the
Department of Defense to calibrate their radars around the world. The CERTO Doppler
beacon provides accurate orbit determination using various ground stations around the
world, but also provides data that will be used to help characterize the ionosphere for space
weather monitoring. A secondary payload consists of two digital imagers to provide color
images of the Earth. The LEO-1 spacecraft will be placed into a 600-km circular 9 p.m.
ascending Sun Synchronous Orbit to optimize its support of the radar calibration
experiment. The 60-kg LEO-1 spacecraft is 3-axis stabilized using three magnetic torque
rods and a reaction wheel for attitude control; and two sun sensors, a 3-axis magnetometer,
and an inertial measurement unit for attitude determination. Communication is provided by
a UHF- transceiver linked to a ground station located in the Leeward Community College in
Honolulu and other partner ground stations. Control of the mission will be done in the
HSFL Mission Operations Center located on the University of Hawaii campus at Manoa.
Integration and testing of the spacecraft will be done in the clean rooms at the HSFL
facilities on the UH campus. The HSFL is using a core team of experienced professionals
supplemented with graduate and undergraduate students to design, build, and test the LEO-
1 spacecraft. Part of the design philosophy includes the development of two nearly identical
spacecraft – the Engineering Model which will be used for integrated testing and then
transitioned to be part of the tesbed/simulator for operations, and the Flight Model. Another
means employed to keep the spacecraft cost to a minimum is the use of commercial-off-the-
shelf (COTS) as much as possible.
I. Introduction
he Hawaii Space Flight Laboratory (HSFL) was established at the University of Hawaii at Manoa for two
primary purposes: (1) to educate students and help prepare them to enter the technical workforce, and (2) to help
establish a viable space industry that will benefit the State of Hawaii.
The Low Earth Orbit Nanosat Integrated Defense Autonomous Systems (LEONIDAS) Program is a
Congressionally sponsored program named after the Spartan King who led a small Greek force that held a much
larger Persian army at bay at the Battle of Thermopylae in 480 BC. This program is of importance to both the United
States and to the State of Hawaii. Its purpose is to develop and demonstrate small-satellite orbital launch capability
from the Pacific Missile Range Facility (PMRF). The objectives of the program include establishing a technical
work force in the State of Hawaii with a development program that trains students to enter the technical work force.
The original vision for the program was to develop a chain of small satellites to pass information to a single ground
station and assist with disaster-relief efforts. It later evolved into a constellation of very small (nano-) autonomous
T
AIAA SPACE 2009 Conference & Exposition
14 - 17 September 2009, Pasadena, California
AIAA 2009-6812
Copyright © 2009 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
American Institute of Aeronautics and Astronautics
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Students&
Support
C-Band
Transponders
HSFL (
HSFL (
Manoa
Manoa
)
)
HMOC
Mission Ops
Project
Management
S/C Design, Build,
I&T S/C+PAD I&T
Ground
Station
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Program Mgt
Launch
Support
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C-Band
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C-Band Xpndrs (VAFB)
Imagers
Images
CERTO Beacon (NRL)
B/U MOC
B/U MOC
UHF Doppler
TLM
Imagers (HSFL)
ORS
Schedule
Tasking
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Solar
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Students&
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HSFL (
HSFL (
Manoa
Manoa
)
)
HMOC
Mission Ops
Project
Management
S/C Design, Build,
I&T S/C+PAD I&T
Ground
Station
Integrated
PAD
Program Mgt
Launch
Support
GS
LEO-1
Super Strypi
Launcher Dev
Oversight
Launch
Support
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Imagers
Images
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B/U MOC
UHF Doppler
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Figure 1. STU-1 Mission Architecture
satellites to aid the Department of Defense (DoD). By the time funding was approved in 2007, the program was
designated to cover two launches and two spacecraft. Oversight of this program was given to the U.S. Army’s Space
and Missile Defense Command (SMDC) with headquarters in Huntsville, AL. In February 2008 the oversight of the
program was transferred to the Operationally Responsive Space (ORS) Office with headquarters at the Kirtland Air
Force Base in Albuquerque, NM.
The first LEONIDAS mission was to consist of a small (30 – 50 kg) satellite to demonstrate a stable platform for
remote sensing (especially missile detection and tracking). The LEONIDAS Missions #1is designated as Science
and Technology for the University (STU) #1. This is the official mission launch designation provided to PMRF. The
first launch of the LEONIDAS Program and HSFL is STU-1, which includes the LEO-1 spacecraft; other secondary
satellites (CubeSats); the Super Strypi launch vehicle, being developed by Sandia National Laboratories for HSFL;
and the payload adapter and deployer (PAD), designed and built by the NASA Ames Research Center.
With the transfer of oversight of the LEONIDAS Program from SMDC to ORS, the mission for the first satellite,
LEO-1, also changed. Originally it was designed to show the suitability of the microsat bus to be a platform for
remote sensing of missile launches and tracking. It then was changed to be a microsat platform to test various
technologies of interest to the Department of Defense including the ORS office. During the six months following the
switchover, new experimental payloads were found and feasibility studies were performed. Finally the payload
manifest was finalized with the inclusion of the C-band radar transponder experiment (CRATEX) for the USAF; the
Coherent Electromagnetic Radio Tomography (CERTO) experiment for the Naval Research Laboratory (NRL); and
the LEO-1 Solar Array Experiment (LSAE). The current design of LEO-1 has successfully passed its Preliminary
Design Review (PDR) in May 2009, and is on course for a launch in 2011.
The LEO-1 is a three-axis stabilized octagonal spacecraft of approximately 60 kg mass and will be launched into
a 600-km 9 p.m. ascending node sun synchronous orbit (SSO). Communication is provided by a UHF- transceiver
linked to a ground station located in the Leeward Community College in Honolulu and other partner ground stations
around the world. Control of the mission will be done in the HSFL Mission Operations Center located on the
University of Hawaii campus at Manoa.
II. Mission Description
This section describes the mission architecture, project management, the launch trajectory, and orbits analysis.
A. Mission Architecture
The LEO-1 spacecraft is part of the designated STU-1 launch from PMRF. Figure 1 shows the top level mission
architecture for STU-1.
HSFL, located at the UH
Manoa campus, is at the
center of the system and will
be responsible for design,
fabrication, project
management; integration &
testing of both the LEO-1
spacecraft and the PAD. All
the payloads being flown on
STU-1 will be integrated on
the PAD at HSFL before
shipment to PMRF for
launch vehicle integration.
HSFL will also be
responsible for the ground
segment including the
ground stations, and mission
operations centered in the
HSFL Mission Operations
Center (HMOC).
American Institute of Aeronautics and Astronautics
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B. Project Management
1. Organization and Staffing
The HSFL employs a small and dedicated team to lead the effort of developing the LEO-1 spacecraft. To fulfill
the university’s mission to educate students, HSFL involves several students for these projects. However, students
are almost by definition inexperienced and relying solely on them would result in an unacceptable level of risk for a
multi-million dollar highly technical project such as LEO-1. The HSFL management has thus hired a core team of
experienced experts in the various fields required to complete the development effort. This cadre of experts serves as
the mentors for the students and help mitigate the risks inherent to the development of a new launch vehicle and
spacecraft.
The HSFL is an element of the University of Hawaii and thus has a primary purpose to educate students. This
purpose is reflected in the organizations employed for the LEO-1 project. By employing a number of students, both
undergraduate and graduate, the HSFL will accomplish the education of students to prepare them to enter the
workforce as already experienced and productive engineers and scientists, and to help keep the development costs of
the projects low. The project management and lead engineers are all experienced full-time HSFL staff or part-time
faculty. The experience of the key personnel allows them to perform multiple tasks within the organization with the
help of undergraduate and graduate students. In some cases the lead engineers will do the work themselves, usually
where a high level of technical expertise or experience is required, or they will supervise the performance of tasks
by students. Even when the lead engineers do the task themselves, one or more students will be involved if at all
possible to observe and learn. The staffing level for most of the project, including staff engineers, faculty, and
students, stays in the range of 8 to 9 Full Time Equivalents (FTEs).
The LEO-1 project organization is led by Dr. Trevor Sorensen, project manager, reporting to the HSFL Director,
Dr. Luke Flynn. The organization of the LEO-1 project is shown in Figure 2.
2. Development Process
The basic design process being followed by the HSFL Team is fairly standard for similar space projects. It is
based loosely on the design process outlined in Reference 1. The development process has been divided into five
phases as defined in Table 1. The HSFL phase nomenclature is based on the NASA system (A-E) while the
Figure 2. LEO-1 Organization
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corresponding DoD nomenclature is in parentheses (I-V). Several milestones have been planned into the
development process including the following:
1. Conceptual Design Review (CoDR)
2. System Requirements Review (SRR)
3. Preliminary Design Review (PDR)
4. Critical Design Review (CDR)
5. Test Readiness Review (TRR)
6. Mission Readiness Review (MRR)
Table 1. LEO-1 Development Process Phases
Phase Name Description
A (0) Concept Definition Defines requirements and baseline conceptual design – ends in SRR
B (I) Definition & Acquisition
Planning
Improves on baseline design with further analyses and trades –
ends in PDR with
preliminary design
C (II) Detailed Design Completion of system design – ends in CDR
D (III) Development Includes procurement, fabrication, integration & testing – ends in launch
E (III) Operations
All operations after launch including Launch & Early Orbit (L&EO), Engineering
Evaluations & Checkout (EE&C), nominal operations, and terminal operations –
ends with reentry or loss of satellite
The developmental process for Phases A & B being followed by HSFL in the development of LEO-1 is shown in
Figure 3. The status of the steps in this process at the time of the PDR is indicated by the color: green = completed,
yellow = in progress, red = not yet started. The Work Breakdown Structure (WBS) for the project contains over 500
tasks.
The LEO-1 project schedule is shown in Figure 4. A temporary funding shortfall with resultant reduction in force
caused the CDR date to be pushed back about six months until May 2010. The spacecraft is currently scheduled to
be completed by June 2011.
Figure 3. Phases A & B Development Process
American Institute of Aeronautics and Astronautics
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2. Systems Engineering
The role of systems engineering in the project is to define and develop an operable system capable of meeting
mission requirements within the imposed constraints, including performance, cost, risk, and schedule. The HSFL
approach follows common systems engineering best practices appropriate to the project and include the following
phases: functional analysis; system synthesis; system evaluation and decision; and system definition. The Systems
Engineering Management Plan document captures in greater detail the approach to the systems engineering effort.
The project systems engineering effort also includes configuration management; oversight of project
documentation; and places a major role in the project risk management strategy.
C. Trajectories and Orbit Analysis
1. Launch Profile
The LEO-1 spacecraft will be ready for launch by June 2011. It will be launched by an enhanced Super Strypi
three-stage, spin-stabilized, solid-propellant rocket from the Pacific Missile Range Facility (PMRF) located on
Kauai in a southerly retrograde direction. The notional launch profile is shown in Figure 5.
The spacecraft will be deployed from the Payload Adapter and Deployer (PAD), being developed by NASA
Ames Research Center. The PAD will be capable of carrying one primary microsatellite and up to 24 or 32 cubesats
or equivalents. The nominal insertion orbit is into a 9 p.m. ascending circular sun synchronous orbit (SSO) with an
altitude of 600 km. and inclination of ~ 98º. After the third stage combustion has terminated, a yo-yo mechanism
will despin the vehicle from 1 rps to no more than 4º per second.
Figure 4. LEO-1 Development Schedule
Figure 5. Notional Enhanced Super Strypi Launch Profile
2
Up to 8 P-PodsUp to 8 P-Pods
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2. Orbit Analysis
The LEO-1 spacecraft has no propulsion subsystem and will be unable to alter its orbit. It is expected to reenter
the Earth’s atmosphere within 25 years after the end of the design mission (three years). The orbital analysis was
performed by the AFRL based upon orbit insertion data provided by Sandia National Laboratories. The assumed
baseline orbit is shown in Table 2. The design length of the mission is three years. Dispersion analyses were also
done, but are not included in this paper.
Three ground stations were assumed for the initial study: Honolulu, Santa Clara CA, and Guildford England.
Figure 6 shows the nominal orbit ground traces of the first three complete orbits of the LEO-1. The importance of a
ground station in Europe is evident from this plot. The altitude prediction is shown in Figure 7, the inclination
prediction in Figure 8, and the ascending node solar time (local time) over the three-year lifetime is shown in Figure
9. This latter plot is important for a desired Sun Synchronous Orbit (SSO) to show how well the satellite keeps to the
desired local time. This particular analysis shows the satellite passing a little sooner than the desired 2100 hours
local time, but can be tweaked by the launch time.
Launch Date
1 December 2010 20:11:05 UTC
Altitude
600 km
Orbit Type
Circular, Sun Synchronous
Inclination
97.79 degrees
Circular Velocity
7.558 km/s
Orbit Angular Velocity
3.723 deg/min
Period
96.69 minutes
Revolution per Day
14.85
Force Model Parameters
21 x 21 EGM96 Earth gravitation
Solar and lunar gravitation
C
D
= 2.2
A/m = 0.0067 m
2
/kg
Jacchia 1970 Lifetime atmosphere
Figure 6. Ground Traces of First Three Nominal Orbits
Table 2. Baseline Parameters for Orbital Analysis
American Institute of Aeronautics and Astronautics
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Figure 7. Orbital Altitude Prediction
Figure 8. Orbital Inclination Prediction
American Institute of Aeronautics and Astronautics
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III. Spacecraft Description
This section describes the spacecraft bus, its configuration, and its subsystems.
D. Configuration and Mass Budget
The initial mass budget of the LEO-1 spacecraft is 62 kg. The mass budget is shown in Figure 10. The external
and internal configurations are shown in Figures 11 and 12.
RAAN Solar Time
19.9
20.0
20.1
20.2
2
0.3
20.4
2
0.5
20.6
2
0.7
20.8
2
0.9
12/1/2010
2/1/2011
4/1/2011
6/1/2011
8/1/2011
10/1/2011
12/1/2011
2/1/2012
4/1/2012
6/1/2012
8/1/2012
10/1/2012
12/1/2012
2/1/2013
4/1/2013
6/1/2013
8/1/2013
10/1/2013
12/1/2013
Date
Solar Time (Hours)
LST
Figure 9. Ascending Node Solar Time
Mass (kg)
Structures 16.3
EPS 8.9
CDH 1.9
Telecom 3.0
ADCS 4.4
Thermal 2.2
Systems 4.4
Payload 11.2
Margin 9.8
TOTAL 62.0
Mass (kg)
Structures 16.3
EPS 8.9
CDH 1.9
Telecom 3.0
ADCS 4.4
Thermal 2.2
Systems 4.4
Payload 11.2
Margin 9.8
TOTAL 62.0
Figure 10. LEO
-
1 Mass Budget
American Institute of Aeronautics and Astronautics
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Zenith
Nadir
Radiator
C-Band
A
ntennas
(2x)
HSFL
Camera
15 in. Lightband
CERTO
Antennas
(5x)
HSFL
Camera
Feedthroughs
Y
Y
Z
Z
Z
Z
X
X
B
aselined
Novatel GPS
Z
Z
Y
Y
X
X
Mass = ~ 57 kg
Figure 11. LEO-1 External View
1111
SST-177
C-band
Transponder
Avionics
Zenith View
Reaction
Wheel
Magtorquers
(3x)
IMU
(2x)
CERTO
HSFL
Camera
(1 of 2)
Batteries
MD2000C-1
C-band
Transponder
Y
Y
X
X
Z
Z
Y
Y
X
X
Z
Z
Nadir View
HSFL
Camera
(1 of 2)
Magnetorquer
Control Unit
Reaction Wheel
Control Unit
GPS
CERTO
Splitter
RDAQs
TTC
CDH
EPS
Figure 12. LEO-1 Internal Configuration
American Institute of Aeronautics and Astronautics
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B. Bus Subsystems
1. System
The spacecraft functional block diagram (FBD) is shown in Figure 13. The basic philosophy used in the design is
a distributed avionics architecture, where each subsystem is controlled by a remote data acquisition (RDAQ) unit
that interfaces with the central processing unit (CPU), which provides the top level executive control. Each of the
spacecraft bus subsystems will be described in this section. The subsystems are: Command and Data Handling
(C&DH), Electrical Power (EPS), Telecommunications (Telecom), Structures, Attitude Determination and Control
(ADCS), Thermal Control (TCS), Payloads, and Flight Software (FSW).
2. Command & Data Handling (C&DH) Subsystem
The spacecraft (S/C) is controllable by real-time uplinked commands, time-delayed command scripts, event-
driven command scripts (e.g. establishing lock with a ground station during a pass), and autonomous onboard flight
software. It has sufficient onboard storage to store at least 24 hours worth of state-of-health (SOH) and payload data.
The S/C is capable of storing delayed commands to support the spacecraft and mission for at least 48 hours.
CDH Subsystem contains one central processor unit (CPU), and multiple microcontroller based Remote Data
Acquisition modules (RDAQs). The Main CPU is a PowerPC 405 processor performing attitude control, telemetry,
command processing, and CMOS image capture and processing. RDAQ processors are 16-bit Harvard RISC or 32-
bit ARM processors performing sensor sampling and Analog-to-Digital, Digital-to-Analog conversion.
Communication between the flight computer and RDAQ processors is via a RS-485 bus at 115kbps.
Communication between RDAQ processors is via I2C bus at 400 kbps. The MIP 405 CPU has been flight-tested in
space.
3
The Attitude Determination and Control Subsystem (ADCS) can be controlled by the ADCS RDAQ in a limited
mode. Telecom transceivers can be handled by COM RDAQ in limited mode. Basic telemetry data sent via separate
beacon radio RDAQ inter-module communication on separate I2C data bus.
The functional allocations for the C&DH are as follows:
• MIP405 Flight Computer
o Command processing and scheduling
o Wideband telemetry generation
o ADCS estimation and control
o HSFL imager interfacing and image data compression
Figure 13. LEO-1 Functional Block
American Institute of Aeronautics and Astronautics
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•
RDAQ Controllers
o
Housekeeping telemetry generation
o EPS subsystem control and measurement
o Launch sensors interfacing (temperature, pressure, acceleration, spin-rate, vibration)
o ADCS sensors interfacing (IMUs, sun sensors, momentum wheel tachometer)
o ADCS actuators interfacing (Magtorquer rods, momentum wheel)
o Payload enable / disable
o Real Time Camera Control
The C&DH Block Diagram is shown in Figure 14.
3. Electrical Power Subsystem (EPS)
The EPS provides enough power to cover all the spacecraft’s needs during all phases of the mission after launch.
It uses photovoltaic elements (solar cells) assembled into solar arrays to generate power and secondary batteries to
store excess energy until it is required to cover a shortage of power from the solar arrays. The EPS fulfills the power
requirements of the fully operating S/C at all times, including during eclipse periods, for the duration of the mission
(two years) of the S/C. The spacecraft generates 28 VDC unregulated current which provides 5 VDC and 3.3 VDC
on a main power bus to all subsystems as needed. The EPS Functional Block Diagram is shown in Figure 15.
The eight body-mounted solar array panels each consist of two solar array modules (strings) with 19 solar cells
per module (Figure 16). To support LEO-1 Solar Array Experiment (LSAE), the arrays are made up of several types
of Triple Junction GaAs solar cells – BTJM (28% efficiency), ZTJ (29.5%), XTJ (29.9%), and ITJ (26.8%). Each
solar array module has Peak Power Tracking (PPT) controlled by the EPS RDAQ, which also measures the voltage
and power of the modules. The solar arrays generate a peak panel power of ~42 W, with average power over an orbit
(with losses) of ~25 W. The average power margin over an orbit is ~5.5 W, while in an emergency low power mode
the power margin is ~15.9 W.
The LEO-1 batteries are A123 Systems High Power Lithium Ion ANR26650 (LiFePO4) mounted in two arrays
providing 242.9 Wh (9.2 Ah @ 26.4 V). These batteries have been flight tested and tested in a HSFL vacuum
chamber and are not subject to explosive failures like so other type of Li Ion batteries.
The nominal weekly operations schedule is shown in Figure 17. This determined the loading for the EPS
performance shown in Figures 18-21. The nominal S/C attitude is Local Vertical Local Horizontal (LVLH) hold,
which keeps the nadir end (+Z) of the S/C always pointed to geocenter.
CRATEXCRATEX
Figure 14. C&DH Block Diagram
American Institute of Aeronautics and Astronautics
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Figure 1
5
.
EPS
Block Diagram
CRATEX
LSAE Step Mode
LSAE Spin Mode
CRATEX
LSAE Step Mode
LSAE Spin Mode
Figure 17. LEO-1 Nominal Weekly Activities
Figure 16. Solar Array Panel Consisting of Two 19-cell Modules
American Institute of Aeronautics and Astronautics
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x
y
-z
1
86
7
5
4
3
2
Sun: 9 AM,
descending
x
y
-z
1
86
7
5
4
3
2
Sun: 9 AM,
descending
Figure 18. Weekly Payload Power Consumption
Figure 19. Solar Flux Per Solar Panel (LVLH Mode)
American Institute of Aeronautics and Astronautics
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4. Telecommunications Subsystem (Telecom)
The spacecraft is capable of two-way communications with the ground regardless of attitude using two redundant
UHF full-duplex RF Datatech ART series transceivers to downlink SOH and payload data (excluding the C-band
transponders and the CERTO beacon) and to receive uplinked commands and flat files. These transceivers have
selectable output power from 50 mW to 5 W, data rate is 150-9600 bps, using 4-level FSK. One receiver is always
on (i.e., it is not possible to turn off both S/C receivers at the same time). This is to prevent the spacecraft from
accidentally being severed from ground control. The Telecom subsystem is designed to not interfere with payload
communications. The Telecom subsystem also includes a simplex VHF beacon that will continuously transmit the
spacecraft identification and some essential SOH data. The beacon is a RF Datatech ZRT series radio modem, with
selectable power output of 100 mW to 5 W and is configured for 2-level FSK 1200 bps. There are three independent
monopole antennas for these transceivers. From the link budget analyses, the UHF downlink margin to the LCC
ground station is 18.23 dB, uplink margin of 14.35 dB, and the VHF Beacon downlink margin is 32.05 dB. The
Telecom Functional Block Diagram is shown in Figure 22.
Figure 20. Solar Array Power Over An Orbit
Figure 21. Battery Charge Status Over One Week
American Institute of Aeronautics and Astronautics
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Figure 23. Avionics Shelf
Sun Sensors
HSFL
Cameras
Figure 24. FOV Analysis
5. Structures
The structure ensures all subsystem components are housed within the spacecraft's mass and volumetric footprint,
and adequately protected from physical forces and environment during all phases of the mission, from integration
and testing, through ground transportation, launch, and space flight. The
structure also accommodates subsystems mounting and interconnect.
The design of the satellite structure incorporates a number of primary
structural members supported by a system of trusses (see Figure 12). The
general shape of the structure is an octagon to maximize the power
obtained from the solar cells. The solar cell substrate doubles as the side
panels for the spacecraft and mount onto the vertical and horizontal
members. The upper solar panels mount onto the top surfaces of the
frame. The basic material used in the structure is Aluminum 7075 T6. A
composite structure was considered, but aluminum was selected because
of its reduced cost and risk/complexity. The structure is designed for
accessibility, which was part of our criteria for a standard bus baseline.
Helping this are removable solar panels (also fulfills an objective of the
LSAE), and external electrical/power feedthroughs. The LEO-1 design
was also kept mechanism-free for reduced associated risk and
complexity.
After the System Requirements Review (SRR), a central avionics
shelf was added to the S/C as shown in Figure 23. This provided
modularity, aided assembly and integration, shortened cable runs,
provided additional radiation shielding, and provided stiffening to the
S/C structure.
One of the requirements that had to be met in the design of the S/C
structure was to not have any object protrude into the field of view
(FOV) of the HSFL cameras or the sun sensors. This was accomplished
as is shown in Figure 24.
The force constraints used during the design of the structure was 16
g axial acceleration, 7 g lateral acceleration, 2.5 rps maximum spin rate,
and the launch environment and vibration tests specified in MIL-STD-
1540E. This resulted in a limit load of 1.25 times maximum load of 16
g axial and 7 g lateral, 1.25 factor of safety with respect to yield, and 1.40 factor of safety with respect to ultimate.
Extensive finite element analyses (FEA) were performed on the LEO-1 structure. The analyses performed included:
Figure 22. Telecom Subsystem Functional Block Diagram
American Institute of Aeronautics and Astronautics
16
Table 3. ADCS Operational Modes
MODE DESCRIPTION
Deploy
ADCS inactive after deployment separation
Tumble
Attitude control disabled to allow tumbling
(e.g. Lost mode)
Detumble
To some fixed inertial attitude
Nominal
LVLH hold attitude mode with nadir facing
Earth and zenith facing Space (along
outward radial) and +X-axis pointing along
velocity vector – this mode is used for
CRATEX contacts
Zenith Imaging
LVLH hold attitude mode with zenith facing
Earth and +X-axis pointing along velocity
vector
Special
Non-nominal mode for special tasks (e.g.,
Earth pointing, inertial pointing)
Sun Lock
To attitude that maximizes solar array
charging
Safe
To attitude that maximizes power
generation and communication coverage
Analysis Spin
1 rpm spin-up and down, for all solar arrays
to reach thermal uniformity (± 3° C)
static (Von Mises Stress), frequency (up to 15
th
mode), and thermo-mechanical. An example is shown in Figure 25.
The maximum displacement of about 4 mm is considered to be too much and a stiffening of the structure was done
post PDR, although the latest information from the launch vehicle designers indicate that the loads will be less than
half of what was used in the initial design.
6. Attitude Determination and Control Subsystem (ADCS)
The spacecraft is three-axis stabilized and capable of autonomous, closed-loop inertial pointing with an accuracy
of 5 deg. or better. Attitude measurement
accuracy is adequate to determine where the
spacecraft is pointing to 1-2 deg. This accuracy is
achievable in real-time, in darkness or sunlight,
and during all phases of the flight after
deployment.
The ADCS provides the spacecraft with the
capability of maintaining the +z face of the
spacecraft pointing in a nadir (geocentric)
direction with an accuracy of ±5°. This is the
attitude required for the C-band transponder,
CERTO, and imager payloads. The operational
modes of the ADCS are given in Table 3. Most of
the mission is spent in the nominal (LVLH Hold)
mode with the nadir (+Z-axis) pointing towards
geocenter. However, the spacecraft has the
flexibility to point to any inertial attitude and
even to spin when required. It is also desirable,
but not mandatory, for spacecraft to be capable of
pointing to and tracking a stationary Earth target.
The ADCS is also responsible for maintaining
zones of exclusion to prevent the S/C from pointing imagers towards the sun where it could result in damage to
Static Analysis : 20 g axial, 10 g lateral (+X), 3 rps
Displacement
(- 4.0 mm max)
Figure 25. FEA Showing Axial Displacements
American Institute of Aeronautics and Astronautics
17
Y
Y
Z
Z
X
X
x
y
z
Top
Right
z
z
x
y
y
x
Front
Magnetorquer
IMU
Reaction wheel
Sunsensor
Magnetorquer
IMU
Reaction wheel
Magnetorquer
IMU
Reaction wheel
Sunsensor
Figure 27. Location of ADCS Components
them. The ADCS software, including algorithms and settings, are accessible to the ground and can be updated or
replaced from the ground during the flight if needed.
Attitude data obtained from the spacecraft’s ADCS sensors are autonomously processed aboard the spacecraft to
determine spacecraft orientation. The attitude determination is accomplished using three space-qualified Aero Astro
MSS-01 Sun Sensors, each with 60
°
full-angle circular FOV providing an accuracy of ~1
°
, and two Microstrain
3DM-GX2 inertial measurement units (IMUs), each of which contains a 3-axis magnetometer, three accelerometers,
and a rate gyroscope.
The ADCS Functional Block Diagram is shown in Figure 26. The commanded control torque vector and the
input vector to the actuators (magnetorquers) are related by: B(R)
a
=
c
. R is the attitude,
a
is the actuator input
vector,
c
is the commanded torque vector (from the control law). B(R)=[B
1
(R)×B
g
B
2
(R)×B
g
B
3
(R)×B
g
] with three
magnetorquers, where B
i
(R), i=1,2,3 are the magnetorquer magnetic fields that depend on the attitude, and B
g
is the
local geomagnetic field. Whenever one of the magnetorquers is aligned closely with the local geomagnetic field,
B(R) cannot be inverted to get the
a
from the
c
. The controller has to switch to a slower-acting underactuated
control law in this situation. Therefore a small reaction wheel was added for full three-axis attitude actuation
capability.
The ADCS provides attitude control by the
use of control actuators, which for LEO-1 are
three magnetorquers and a small reaction wheel.
The magnetorquers are space-qualified
Vectronic VMT-35 which each have a magnetic
moment of 18 Am2 @ 100 mA and are
controlled by a Vectronic Torque Control Unit.
The reaction wheel (RW) is a space-qualified
Sinclair RW-0.03-4-ASYNC-2-1-0 which has a
nominal torque greater than 2 mNm and
nominal momentum of 30 mNm-sec @ 5600
rpm. The layout of the ADCS actuators and
components is shown in Figure 27.
The ADCS control scheme is nonlinear and
continuous, based on feedback of trajectory
tracking errors. Control law
c
designed for
global tracking of desired attitude and angular
velocity trajectories.
4
Figure 26. ADCS Functional Block Diagram
American Institute of Aeronautics and Astronautics
18
Figure 28. TCS Functional Flow Block Diagram
The actuator model developed generates inputs to actuators (RW and magnetorquers) and the control torque
a
applied by the actuators. This control torque allocation scheme, for a RW along the pitch axis, is given in Reference
5. This scheme is singular when the local geomagnetic field is perpendicular to the pitch (RW) axis.
The ADCS estimation and filtering scheme is nonlinear and discrete, based on deterministic ellipsoidal bounds
on measurement errors and dynamic flow uncertainty. Measurements are: direction vectors (for attitude) and angular
velocity vector, all in spacecraft body frame. Estimation scheme for an orbiting satellite without control torques
given in Reference 6. Propagation of attitude and angular velocity between measurements carried out using a Lie
group variational integrator. Estimation scheme for an orbiting satellite, with the inclusion of a feedback control
torque, is being developed and prepared for publication.
The nominal (LVLH Hold) attitude mode requires a pitch rate of 0.063 º/s. The ADCS pitch rate capability is
0.286º/s. ADCS control software simulations suggest it would require about 190 seconds (less than 1/30 of an orbit
at 600 km altitude) to rotate the satellite by 30º (to 0.4º accuracy), beginning with an initial attitude error of about
15º, an initial angular velocity error of about 1.15º/s, and a maximum torque <1.8 milli-Nm. De-tumbling maneuver
followed by stabilization of nadir-pointing attitude carried out in ADCS simulation shows satisfactory de-tumbling
from an initial attitude error of 30º and initial angular velocity of 0.098 rad/s within 1/20 of an orbit at 600 km
altitude.
7. Thermal Control Subsystem (TCS)
The TCS is designed to ensure all S/C subsystem
components are thermally controlled for operation,
both in sunlight and eclipse periods. Due to the fairly
benign thermal environment of low Earth orbit,
passive thermal control was determined to be
sufficient. The internal power dissipation required to
be handled by the TCS is from a minimum of 14 W to
a maximum of 90 W. The TCS hardware consists of
multi-layer insulation (MLI) blankets on the interior
S/C panels, appropriate surface finishing (e.g., paints),
five thermostats with redundancy, heaters for critical
components, temperature sensors, and a radiator. The
TCS Functional Flow Block Diagram is shown in
Figure 28.
Detailed simulations were performed of the spacecraft including all sub-systems and payload using Thermal
Desktop using finite differences. Simulations were done for a nominal hot case and a cold case. Trades studies for
the thermal management changed the layout of internal components.
One of the LSAE tests require the spacecraft obtain temperature uniformity of the solar array panels by spinning.
Simulations were also done of the spinning spacecraft for this test of the solar cells. The results are shown in Figure
29 and show that thermal uniformity of the solar panels can be achieved after only about 30 minutes with a spin rate
of 1 rpm. After achieving thermal uniformity, the spacecraft has to continue the experiment for one orbit.
~ 30 min
Figure 29. Solar Array Panels Temperature With S/C Spinning at 1 rpm
American Institute of Aeronautics and Astronautics
19
8. Flight Software(FSW)
The flight software autonomously monitors and maintains the spacecraft’s state of health, controls the operation
of the spacecraft, monitors the SOH of the payloads, performs the calculations for the ADCS, and can activate or
deactivate the payloads. The FSW will always be recoverable from loss of power or function without ground
support. It can also be reloaded or modified from the ground without endangering the spacecraft SOH or mission.
The FSW architecture is shown in Figure 30.
The FSW operates as a state machine, where the spacecraft operational modes are defined as states. The Image of
State is contained in shared memory. There is a subprocess for each important subtask, with the Executive Process
controlling system as a whole. Commands are implemented through successive changes in state commands stored as
a linear sequence in time on the command queue.
The FSW performs the navigation calculations for the spacecraft to generate an orbit and position that is used by
other subsystems. The current position of the spacecraft can be determined on the ground using the signal from the
CERTO Doppler Beacon and then uplinked to the spacecraft as either an ephemeris or state vector. A backup
method is to uplink a two-line element (TLE) obtained from the NORAD website. An autonomous method for
determining the orbital position is to process onboard the data obtained from the NovaTel GPS unit. Thus the S/C
will be able to determine its position even if it is out of touch with the ground for extended periods. The FSW
operating system is the QNX Neutrino Real Time OS, which was selected for its stability and support for the MIP
405 C&DH architecture.
IV. Payloads
This section describes the payloads being carried by the LEO-1 spacecraft.
A. LEO-1 Solar Array Experiment (LSAE)
The LEO-1 spacecraft is performing the LEO-1 Solar Array Experiment (LSAE) to test new solar array
technologies, including a modular solar array technology developed at DR Technologies, Inc.
7
The technology has
Operationally Responsive Space (ORS) applications and would benefit from a flight demonstration. One objective
of the LSAE is to flight test plug-and-play solar panel modules and the LSAE incorporates conventional as well as
advanced technology photovoltaics that could be used in future space missions.
IMU Mag
Torque
·
Sensors
GPS
Power
Sense
Control
Power
Sense
Control
Beacon Sec.
Rx/Tx
Primary
Rx/Tx
COMM RDAQ
Monitor
ADCS RDAQ
Monitor
GPS RDAQ
Monitor
EPS_B RDAQ
Monitor
E
PS_A RDAQ
Monitor
SENS RDAQ
Monitor
Sun Sensors
Main CPU
Core
Executive
S
hared Memory
RS-485
Ethernet
D
isk
Attitude
Determination
and Control
Imager
Control
RADCAL
Control
CERTO
Control
Telecomm
Command
Server
State of
Health
Command
Queue
Data Server
RS-
232
RS-
232
RS-
232
RS-
232
Localnet
IMU Mag
Torque
·
Sensors
GPS
Power
Sense
Control
Power
Sense
Control
Beacon Sec.
Rx/Tx
Primary
Rx/Tx
COMM RDAQ
Monitor
ADCS RDAQ
Monitor
GPS RDAQ
Monitor
EPS_B RDAQ
Monitor
E
PS_A RDAQ
Monitor
SENS RDAQ
Monitor
Sun Sensors
Main CPU
Core
Executive
S
hared Memory
RS-485
Ethernet
D
isk
Attitude
Determination
and Control
Imager
Control
RADCAL
Control
CERTO
Control
Telecomm
Command
Server
State of
Health
Command
Queue
Data Server
RS-
232
RS-
232
RS-
IMU Mag
Torque
·
Sensors
GPS
Power
Sense
Control
Power
Sense
Control
Beacon Sec.
Rx/Tx
Primary
Rx/Tx
COMM RDAQ
Monitor
ADCS RDAQ
Monitor
GPS RDAQ
Monitor
EPS_B RDAQ
Monitor
E
PS_A RDAQ
Monitor
SENS RDAQ
Monitor
Sun Sensors
Main CPU
Core
Executive
S
hared Memory
RS-485
Ethernet
D
isk
Attitude
Determination
and Control
Imager
Control
RADCAL
Control
CERTO
Control
Telecomm
Command
Server
State of
Health
Command
Queue
Data Server
RS-
232
RS-
232
RS-
232
RS-
232
Localnet
CRATEX
IMU Mag
Torque
·
Sensors
GPS
Power
Sense
Control
Power
Sense
Control
Beacon Sec.
Rx/Tx
Primary
Rx/Tx
COMM RDAQ
Monitor
ADCS RDAQ
Monitor
GPS RDAQ
Monitor
EPS_B RDAQ
Monitor
E
PS_A RDAQ
Monitor
SENS RDAQ
Monitor
Sun Sensors
Main CPU
Core
Executive
S
hared Memory
RS-485
Ethernet
D
isk
Attitude
Determination
and Control
Imager
Control
RADCAL
Control
CERTO
Control
Telecomm
Command
Server
State of
Health
Command
Queue
Data Server
RS-
232
RS-
232
RS-
232
RS-
232
Localnet
IMU Mag
Torque
·
Sensors
GPS
Power
Sense
Control
Power
Sense
Control
Beacon Sec.
Rx/Tx
Primary
Rx/Tx
COMM RDAQ
Monitor
ADCS RDAQ
Monitor
GPS RDAQ
Monitor
EPS_B RDAQ
Monitor
E
PS_A RDAQ
Monitor
SENS RDAQ
Monitor
Sun Sensors
Main CPU
Core
Executive
S
hared Memory
RS-485
Ethernet
D
isk
Attitude
Determination
and Control
Imager
Control
RADCAL
Control
CERTO
Control
Telecomm
Command
Server
State of
Health
Command
Queue
Data Server
RS-
232
RS-
232
RS-
IMU Mag
Torque
·
Sensors
GPS
Power
Sense
Control
Power
Sense
Control
Beacon Sec.
Rx/Tx
Primary
Rx/Tx
COMM RDAQ
Monitor
ADCS RDAQ
Monitor
GPS RDAQ
Monitor
EPS_B RDAQ
Monitor
E
PS_A RDAQ
Monitor
SENS RDAQ
Monitor
Sun Sensors
Main CPU
Core
Executive
S
hared Memory
RS-485
Ethernet
D
isk
Attitude
Determination
and Control
Imager
Control
RADCAL
Control
CERTO
Control
Telecomm
Command
Server
State of
Health
Command
Queue
Data Server
RS-
232
RS-
232
RS-
232
RS-
232
Localnet
CRATEX
Figure 30. Spacecraft Flight Software Architecture
American Institute of Aeronautics and Astronautics
20
The LSAE uses high efficiency (~29%) solar cells that are laid out in modules (strings) of 19 cells, with two
modules per side panel. There are eight such panels on the spacecraft as shown in Figure 31. Panels 1-5 contain
mostly standard technology solar arrays and provide power for the periods when not in a LSAE test mode. The
remaining three panels (6-8) contain the experimental modules. Some of the cells have experimental coatings on
their glass covers.
During the weekdays, the spacecraft will be oriented such that five of the arrays provide power throughout the
sunlit portion of the orbit (indicated as Weekday Arrays in the figure). Each weekend (assuming no other spacecraft
operations are planned), the spacecraft will be rotated about the z-axis (major axis of the cylindrical spacecraft) by
180 degrees in order to accumulate exposure time on the three arrays normally shadowed. Approximately once per
month, the weekend configuration will include a slow spin about the z-axis (~1 rpm). These LSAE tests are
described in more detail in the Operations section.
As part of its normal SOH data collection, the spacecraft collects and stores the following data for each solar
array module (string): string current, string voltage, panel temperature, sun sensor readings, and time stamp. SOH
data are downlinked to ground stations when possible.
B. C-Band Radar Transponder Experiment (CRATEX)
The CRATEX experiment payload is designed to flight test a replacement to the RADCAL payload currently
operating in a degraded fashion on the RADCAL satellite, which was launched in 1993 from Vandenberg Air Force
Base (VAFB) with an expected lifetime of one year. LEO-1 with the CRATEX payload will test some new
technology that can be used to support ground based Range Instrumentation’s performance monitoring functions.
Ground based Range Instrumentation users are minimally comprised of Tri-services and NASA Test Ranges.
Although the optimal satellite orbit is 850 Km altitude, with an inclination of 65° or higher, the LEO-1 orbital
altitude of 600 km in a Sun-Synchronous (inclination ~98°) is considered by VAFB to be satisfactory for testing
purposes. A C-Band transponder is used to calibrate the ground radars. An orbit determination system is required
that achieves a minimum of 5 meter accuracy. The orbit may be determined on-board, on the ground or through a
combination of space and ground calculations, but must be available to users within 48 hours. The primary method
used on LEO-1 for orbit determination is the CERTO Doppler system and the backup method is a NovaTel GPS
receiver.
Primary
(
“weekday”)
panels
Panel 1
Panel 2
Panel 3
Panel 4
Panel 5
Panel 6
Panel 7
Panel 8
Standard
Module
Experimental
Module
Primary
(
“weekday”)
panels
Panel 1
Panel 2
Panel 3
Panel 4
Panel 5
Panel 6
Panel 7
Panel 8
Standard
Module
Experimental
Module
Figure 31. Solar Array Module Arrangement to Support LSAE
American Institute of Aeronautics and Astronautics
21
Figure 32. SST 177C C-Band Transponder
Figure 33. MD2000C C-Band Transponder
Figure 34. CERTO-R2 Beacon and External Components
VAFB is supplying two C-band transponders – the
space-qualified SST-177C (currently being flown on
DMSP 15 satellite) and the transponder to be flight-
tested in space, the MD2000C (Figures 32 and 33
respectively). Two AntDevCorp Quadrifilar Helix C-
Band antennas are being used. Both transponders
receive ground C-Band radar signal of 5765 MHz and
respond coherently with 5765 MHz on the SST 177C or
at 5690 MHz from the MD2000C. This allows
monitoring of the tracking performance of C-Band
radars, while testing the performance of the MD2000C
in comparison to the older technology of the SST 177C.
Radar data (TRAE) is compared with a known satellite
position (precise ephemeris) obtained from processed
Doppler data or GPS data. A coherent transponder is
necessary for radars equipped with range rate
subsystem. The satellite provides a dynamic target for
testing the total radar system. C-Band transponders are
normally off and must be commanded on just prior to
requested use and off again immediately after each
requested use. Track period is software limited to a
maximum of 20 minutes. Transponders are routinely
activated on an average of 6 times a day.
Ephemeris must be determined to better than five
meters accuracy. The CERTO Doppler Beacon is the
primary means of determining precise ephemeris on
LEO-1. It achieves less than 5 meters accuracy; the
average 2004 Orbit Overlap Total Magnitude RMS was
2.2 meters. The Doppler transmitter remains on at all
times. Satellite Doppler data are automatically collected
and stored at numerous geodetically distributed sites.
Once a day, the data are transferred to Western Range
VAFB for initial editing and processing. The edited
data are then forwarded to the National Geospatial-
intelligence Agency (NGA) for final processing and posting of resultant ephemeris on RPMWEB site for user
community access.
C. Coherent Electromagnetic Radio
Tomography (CERTO) Experiment
The primary purposes of the CERTO
Experiment provided by the Naval Research
Laboratory (NRL) are to detect when and
where radiowave propagation through the
ionosphere is adversely affected by scintillation
and refraction and to provide a global map of
ionospheric densities and irregularities to
improve current models of the ionosphere. The
CERTO Experiment makes use of the Doppler
beacon required by the CRATEX payload for
precise orbit determination. The beacon signal
provides a carrier wave transmission to the
receiving ground stations, which use the shift
in frequencies to help determine the orbit.
However, the radio signal gains noise passing
through the ionosphere, which is the
American Institute of Aeronautics and Astronautics
22
Figure 35. Prosilica GE2040C
phenomenon that is being investigated by the CERTO Experiment. The noise is separated from the carrier signal at
the ground station and the resulting carrier signal is used to determine the orbit for the CRATEX experiment and the
noise is used by the CERTO Experiment to determine the ionospheric characteristics.
The CERTO-R2 beacon (Figure 34) is dual frequency – 150 MHz (VHF) and 400 MHz (UHF) and each
transmission uses its own antennas (four for the 150 MHz and one for the 400 MHz). The phase and frequency
fluctuations of the VHF and UHF signals help in the monitoring of scintillations in the ionosphere and resulting
regional maps of radio signal disruptions. By tracking the CERTO signals as LEO-1 passes over the ground station
allows data from oblique and vertical paths through the ionosphere to be obtained. Additional data are obtained from
GPS satellite occultations. The resulting data are used for reconstructions using computerized ionospheric
tomography.
NRL wants the CERTO beacon to be operating continuously during the mission if at all possible, but the beacon
can be turned to lower power standby mode or off if spacecraft emergency or power state requires it.
D. HSFL Imagers
The HSFL Imagers provide imagery that fulfills a secondary mission objective and will be used by HSFL for
public outreach and to inspire interest in science education. Color images of the Hawaiian Islands with little or no
cloud cover should provide good public relations value to
HSFL and its sponsors. A final product of this experiment is
a color image showing the Hawaiian Islands with little or no
cloud cover. This can be done with either a single image or
a mosaic of images. Two cameras are included in LEO-1.
One is a nadir-pointing (in normal LVLH Hold mode) and
the other is zenith pointing. This latter imager fulfills a
mission requirement to document the separation of LEO-1
from the PAD during orbital deployment, but it would
subsequently be a backup camera for the nadir imager.
The Nadir Imager is a Prosilica GE2040C color 2048 x
2048 (4 MP) CCD image sensors with a narrow FOV
(NFOV) lens placed on nadir-pointing side of spacecraft
(see Figure 35). Its frame rate is up to 15 FPS. The lens
selected for this camera is a Zeiss 85mm f/1.4 ZF Planar T
manual focus lens with a FOV of 28.5°. Its focus range is 1
m to infinity and it has aperture range of f/1.4 – f/16.
The Zenith-pointing imager is a Prosilica GE1050C 1024
x 1024 (1 MP) color camera with WFOV lens to capture
spacecraft deployment from PAD. Its frame rate is 60 FPS.
The lens selected for this camera is an Edmund Optics
Varifocal 4.0-12 mm manual focus lens with a FOV of
93.6° to 31.2°. Its focus range is 0.3 m to infinity and it has
aperture range of f/1.2 – f/16.
American Institute of Aeronautics and Astronautics
23
V. Operations
This section describes the operations concept for the LEO-1 mission.
A
. Operations Concept and Phases
T
he LEO-1 on-orbit mission is divided into three phases.
1. Engineering Evaluation & Checkout to Initial Operational Capability
The Engineering Evaluation & Checkout (EE&C) is the checkout period of the LEO-1 after orbit insertion and is
expected to last about a month (or as long as necessary to complete the checkout). This phase concludes with its
commissioning when it has achieved initial operational capability (IOC). During the EE&C the following tasks are
performed:
•
Testing of spacecraft bus subsystems and modes.
•
Testing of ground segment and data flow.
•
Testing of CRATEX & CERTO payloads:
o Both C-band transponders and CERTO beacon tested
o Orbit determination method tested for 5-m accuracy
o Test interfaces with USAF for tasking and delivery of orbit data
•
Testing of LSAE payloads:
o Experimental solar panels and data acquisition tested
o LSAE operational modes tested
o Heat transfer model verified for various LSAE modes
• Testing of HSFL Imagers:
o Take images using real-time and delayed commands using all cameras with storage and downlink of
images.
2. Nominal Operations
Nominal (CRATEX) Mode
The nominal mode for the LEO-1 spacecraft (i.e., the mode in which is spends most of its time) is to support the
CRATEX and CERTO payloads. This mode has the following characteristics:
• LEO-1 in LVLH hold (+X or –X forward) attitude (see Figure 36).
• Each ground radar requires calibration at least once per week (using Monday to Friday schedule).
• A weekly schedule request is provided by VAFB (30 Space Wing) >3 days in advance. HSFL integrates
requests with the overall LEO-1 schedule. The final CRATEX contact schedule is provided to 30 SW by the
HMOC at least two days in advance for nominal commanding.
• A C-band transponder is turned on and off by time-delayed script before and after each calibration contact
(script nominally uploaded >1 day in advance). Only one transponder and one antenna are active at a time.
Transponders and antennas 1 and 2 are used in alternating months. Transponder 1 (SST 177C) is the
reference transponder and Transponder 2 (MD2000C) is the experimental transponder.
• There are a maximum of five calibration passes per day (each of 20 minutes maximum duration limited by
watchdog timer).
• The target ground radar activates the C-band transponder and executes a calibration contact. The CERTO
Doppler signal is also received by some ground stations.
• HSFL sends GPS position data to VAFB if required (backup mode).
• Ionospheric characterization data extracted from Doppler data are obtained by target ground stations from
the CERTO beacon (which is on continuously).
LSAE Modes
During the nominal CRATEX operations period (Monday-Friday), no LSAE activities are scheduled except
routine monitoring of solar array performance (solar array SOH data collected and stored at 4 minute intervals).
During weekends with no CRATEX contacts, the spacecraft is rotated spacecraft 180° about z-axis (yaw) to put –x
axis forward while maintaining LVLH hold attitude This exposes the experimental panels 6, 7, and 8 (refer to Figure
31) to sunlight and puts the standard solar arrays 2, 3, and 4 into shadow during the orbit. Solar array SOH data still
collected and stored at 4-minute intervals. This mode, called the LSAE Mode, is shown in Figure 37.
American Institute of Aeronautics and Astronautics
24
Once a weekend, if possible, the spacecraft collects data for a current-voltage (I-V) curve for each SA module.
This is called the LSAE Step Mode and is shown in Figure 38. To execute this mode, the following actions are done:
a) Spacecraft goes to inertial attitude with sun normal to side panel
b) Spacecraft points each of the eight sides normal to the sun (beta angle of zero) for long enough to
allow for settling of data.
Once a month the spacecraft goes into a slow (1 rpm) spin to achieve thermal stability of the solar arrays (all
solar arrays within 3 degrees Celsius) to determine their relative performance with a constant temperature. This
mode is called the LSAE Spin Mode and is shown in Figure 39. To execute this mode, the following actions are
done:
LV
L
H
LH
L
V
L
V
LH
Sun Vector
7
7
7
6
8
68
6
8
LSAE MODE
A
ttitude: LVLH Hold, -x
L
SAE: 5,6,7,8,1 sunlit
2,3,4 dark
SA SOH Interval: 4 minutes
Imager: Nadir
(Weekends)
L
V = Local Vertical
L
H = Local Horizontal
+z
+z
+z
-x
-x
-
x
Figure 36. Nominal (CRATEX) Mode Figure 37. LSAE Mode
Sun Vector
LSAE STEP MODE
(Once per weekend)
STEP 2: Rotate to Each Panel and Dwell
1
1
1
1
1
1
1
1
2
2
2
2
2
2
2
2
3
3
3
3
3
3
3
3
4
4
4
4
4
4
4
4
5
5
5
5
5
5
5
5
6
6
6
6
6
6
6
6
7
7
7
7
7
7
7
7
8
8
8
8
8
8
8
8
Attitude: Sun Normal
LSAE: All sunlit in turn
SA SOH Interval: 1.5 sec
Imager: N/A
Sun Vector
NOTE: Rotation can be CW or CCW
A
A
Figure 38. LSAE Step Mode
a) Rotate to Sun Normal Mode
b) Rotate to Each Panel and Dwell
Figure 39. LSAE Spin Mode
LV
LH
LH
LV
LV
LH
Sun Vector
7
7
7
6
8
6
8
6
8
ZENITH IMAGING MODE
Attitude: LVLH Hold, +x/-x
LSAE: 1,5,6,7,8 sunlit
2,3,4 dark
SA SOH Interval: 4 minutes
Imager: Zenith
(As needed – weekends)
LV = Local Vertical
LH = Local Horizontal
-z
-z
-z
+x
+x
+x
Figure 40. Zenith Imaging Mode
LV
LH
LH
LV
L
V
LH
S
un Vector
3
3
3
2
4
24
2
4
NOMINAL (CRATEX) MODE
Attitude: LVLH Hold, +x
L
SAE: 1,2,3,4,5 sunlit periods
6,7,8 dark always
SA SOH Interval: 4 minutes
Imager: Nadir
(Weekdays + priority)
LV = Local Vertical
LH = Local Horizontal
+x
+x
+
x
+
z
+z
+z
Sun Vector
LV (nadir)
L
V
LH
Sun Vector
~8
°
STEP 1: Rotate to Sun Normal Mode
LSAE STEP MODE
(Once per weekend)
Nominal
(LVLH Hold)
Sun Normal
(Inertial Hold)
Attitude: Sun Normal
LSAE: All sunlit in turn
SA SOH Interval: 1.5 sec
Imager: N/A
Sun Vector
LV (nadir)
LV
LH
Sun Vector
~
8°
STEP 1: Rotate to Sun Normal Mode
Attitude: Sun Normal
LSAE: All sunlit in turn
SA SOH Interval: 1.5 sec
Imager: N/A
LSAE SPIN MODE
(Once per month)
STEP 2: Spin at 1 rpm until 1
orbit past SA thermal stability
Sun Vector
Sun Vector
z-axis
N
OTE: Rotation can be CW or CCW
American Institute of Aeronautics and Astronautics
25
a) Spacecraft goes to inertial attitude with sun normal to side panel
b) Spacecraft goes to an attitude with the sun vector normal to the spacecraft z-axis
c) Spacecraft starts 1 rpm spin about the z-axis
d) Once thermal stability achieved (all solar arrays +/- 3
°
C) spacecraft completes one orbit with spin. Thermal
analysis has shown that this thermal stability can be achieved in less than one orbit. Data are collected at 1.5
second intervals.
Earth Imaging
In Nominal Mode the nadir camera takes images of the Earth. This camera has a comparatively narrow FOV. The
similar camera on the Zenith end of the spacecraft that is used to image the separation of the LEO-1 from the launch
vehicle, can also be used as a backup for the Nadir imager for obtaining images of the Earth, although with lower
resolution and a wider-angle lens. In order to accomplish this, the spacecraft has to be flipped upside-down, so that
the zenith end is pointing to the Earth. This mode, called the Zenith Imaging Mode, is shown in Figure 40.
Images can be taken by time-delayed script or real-time command on a non-interference basis. They will be
stored onboard and then downlinked when convenient.
Nominal Payload Operations Schedule
Having determined the various operational modes to fulfill the objectives of the payloads, a weekly schedule is
formed that ensures that there are no conflicts between the spacecraft bus and the various payloads. A nominal
weekly schedule showing the operations of the various payloads and modes is shown in Figure 41.
3. Mission Termination
New missions launched into LEO have to have a means to dispose of the satellite within 25 years of the mission
termination in order to remove it as potential space debris. This can be done either using a system to change the
orbital velocity of the satellite (e.g., using propulsive or drag devices), or by natural orbital decay. For LEO-1, which
has no propulsion subsystem, the latter method was selected. The current analysis shows that with an initial altitude
of 600 km, that the satellite will reenter within 25 years of the nominal mission end.
At the end of mission, the spacecraft payloads and transmitter will be turned off. There is no further action
required before reentry into the atmosphere.
B. Operational Modes
A number of operational modes have been identified for the LEO-1 spacecraft to accomplish its mission. These
operational modes are identified and described in Table 4. There is a close correlation between these modes and the
ADCS modes that were identified in Table 3.
C. Operations Staffing and Process
The basic staffing to support operations of the LEO-1 mission after launch is shown in Table 5. The basic
operations philosophy of the LEO-1 mission is to use the spacecraft engineers as operations personnel during the
EE&CO period of the first month until commissioning. During nominal operations, the engineers will be working
other projects and will be used to support operations as needed (e.g., anomaly resolution) or to just a very low level.
Most of the operations will be performed by the spacecraft controllers, which will be mainly students. This also
provides an educational benefit to the project, to familiarize students with the operation and workings of an
operational satellite. The spacecraft is being designed for mostly autonomous operation, including automatic
contacts with the ground where command files are uplinked, and data files are downlinked with human intervention.
Figure 41. Nominal Payloads Weekly Schedule
American Institute of Aeronautics and Astronautics
26
Once the mission has achieved a certain level of maturity in which there is a high level of confidence that the
automated systems will work, then the mission can enter the mature operations, which requires minimal staffing
by operations personnel. Only a single 8-hour shift for five days a week is anticipated for nominal and mature
operations, while the EE&CO period will require operations personnel for seven days a week. Although the
nominal and mature operations are only five days a week, they will usually not correspond to Monday to Friday,
since the LSAE modes, which will require supervision, are done on weekends.
EE&C Nominal Mature Shift
Position FTE FTE FTE h/d Comment
Operations Manager 1 0.5 0.1 8/5
Spacecraft Controller 2 0.9 0.1 8/5 8/7 for EE&CO
Ground Network Controller 1 0.1 0.1 8/5
Scheduler 1 0.5 0.5 8/5
Mission Planner 1 0.5 0.5 8/5
Data Manager 1 0.5 0.1 8/5
S/C Analyst 4 0.5 0.1 8/5 8/7 for EE&CO
Software Analyst 2 0.5 0.1 8/5 8/7 for EE&CO
Orbit Analyst 1 0.5 0.1 8/5
TOTAL
14
4.5
1.7
Table 4. Spacecraft Operational Modes
Table 5. LEO-1 Operations Staffing
American Institute of Aeronautics and Astronautics
27
The mission operations process developed for LEO-1shown in Figure 42 includes the following four sub-
processes:
a) Planning Process – this takes the Mission Operations Plan, and the Flight Rules & Mission Requirements
(all developed before launch), with the current status of the spacecraft, the ground network, and any payload
customer requests (such as CRATEX contact requests received from 30 Space Wing at VAFB) and with the
aid of the current orbit ephemeris and visibility windows, determines a master schedule, and generates
timelines for orbits as necessary. From the timeline a command script is automatically generated, and if
necessary a set of pointing quaternions. The command script and quaternions are then verified in the
Testbed/simulator, and if successful, passed along to the Execution Process for implementation.
a) Execution Process – this handles the real-time operations of the ground network (ground stations and
network) and the interactions with the LEO-1 spacecraft. Contact with the spacecraft are monitored, during
which the command script and flat files are uplinked, while the SOH (both real-time and archived) and
payload data are downlinked to the ground station and transferred to the HMOC. All contacts and events are
logged, both automatically and manually by the controllers.
b) Data Management Process – this takes the downlinked SOH and payload data from the ground station, then
archives, processes, and distributes the data.
c) Analysis Process – the processed data are analyzed and trended over time to determine the health of the
spacecraft and the success in accomplishing mission objectives. If anomalies are detected, then a spacecraft
or payload engineer is alerted to initiate an anomaly resolution process.
VI. Conclusion
HSFL is well on the path to design, build, and launch a microsatellite. In May 2009 the LEO-1 PDR was
successfully completed, with the government customer giving it an “A” grade. The paradigm being used of having a
core team of full-time professional engineers and supplementing this with a cadre of professors and students (both
graduate and undergraduate) to develop a technologically advanced satellite with a fairly complex mission at
moderate cost and risk seems to be working. HSFL plans this satellite to be the first on the path to many and even
more ambitious missions.
DATA MANAGEMENT
PROCESS
Contact Plan
Constraints
Configuration
Definitions
Mission
Planning &
Scheduling
Process
Mission
Planning &
Scheduling
Process
Contact Plans
Command Loads/Scripts
Ground
Network
Commands
(GN & S/C)
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Payload Data Pipeline
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(Level 0, Eng.
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ANALYSIS PROCESS
PLANNING
PROCESS
MP
Modify
& Verify
SC, GC
Verify or
Intervene
Spacecraft
Analyst
Verify
Spacecraft
Analyst
Verify
Orbit
Analyst
Verify
Orbit
Analyst
Verify
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Figure
42
.
Mission Operations Process
American Institute of Aeronautics and Astronautics
28
Acknowledgments
The authors wish to thank the following individuals for their contributions to the material and data presented in
this paper - from HSFL: Byron Wolfe, Jason Akagi, Miguel Nunes, Zachary Lee-Ho, Mark Wood, Harold Garbeil,
and Keith Horton; from VAFB: Martin Prochazka (WROCI/InDyne); from AFRL: Moriba Jah (AFMC
AFRL/RDSM); From NRL: Paul Bernhardt (Code 6754); and from DR Technologies, Inc: Ted Stern. We would
also especially like to thank Luke Flynn (Director, HSFL), Lavina Chatlani (HSFL), and Mark Franz (IPA USAF
AFSPC ORS) for their support of this project.
References
1
Wertz, J. R., and Larson, W. J., (editors), Space Mission Analysis and Design,, 3
rd
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Schindwolf, E. J., Swanson, B., E., Millard, W. A., “Launch of ‘Smallsats’ Using Lo-Cost Sounding Rocket Technologies,
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th
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http://www.mpl.ch/DOCs/MPLdoc_00000221.pdf
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Sanyal, A. K., and Chaturvedi, N. A., “Almost Global Robust Attitude Tracking Control,” AIAA-2008-6979, Proceedings of
the AIAA Guidance, Navigation, and Control Conference, Honolulu, HI, August, 2008.
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Sanyal, A. K., and Lee-Ho, Z., “Robust Attitude Tracking Control of a Small Satellite in Low Earth Orbit,” to be presented
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