# **[APPLIED ENGINEERING AND TECHNOLOGY DIRECTORATE]{.underline}**

![](media/image1.png){width="1.4791666666666667in"
height="1.2395833333333333in"}

## Flight Dynamics Analysis Branch

## End of Fiscal Year 2001 Report

A.  *Barnes, T. Stengle, S. Truong*

![](media/image2.png){width="1.1875in" height="0.75in"}

National Aeronautics and

Space Administration

Goddard Space Flight Center

Flight Dynamics Analysis Branch/Code 572

Greenbelt, Maryland 20771

December 2001

##  

## Flight Dynamics Analysis Branch

## End of Fiscal Year 2001 Report

National Aeronautics and

Space Administration

Goddard Space Flight Center

Flight Dynamics Analysis Branch/Code 572

Greenbelt, Maryland 20771

## Acknowledgements

The contents of this report are based on inputs generously supplied by
members of the Guidance, Navigation, and Control Center (GNCC) Flight
Dynamics Analysis Branch (FDAB) at NASA/Goddard Space Flight Center
(GSFC).

The primary editors of this report are:

Alex Barnes, Flight Dynamics Analysis Branch, Code 572

Tom Stengle, Head, Flight Dynamics Analysis Branch, Code 572

Son Truong, Flight Dynamics Analysis Branch, Code 572

Additional information concerning FDAB activities may be obtained
through the following Branch management:

John P. Lynch, Associate Head, Flight Dynamics Analysis Branch, Code 572

Karen Richon, Associate Head, Flight Dynamics Analysis Branch, Code 572

Published copies of this report are available from:

Catherine A. Waltersdorff

Flight Dynamics Analysis Branch

Code 572

NASA/Goddard Space Flight Center

Greenbelt, MD 20771

A Hypertext Markup Language (HTML) version of this document will be
available on the Internet World Wide Web (WWW) in the Uniform Resource
Locator (URL):

http://fdab1.gsfc.nasa.gov/

**Abstract**

This report summarizes the major activities and accomplishments carried
out by the Flight Dynamics Analysis Branch (FDAB), Code 572, in support
of flight projects and technology development initiatives in Fiscal Year
(FY) 2001. The report is intended to serve as a summary of the type of
support carried out by the FDAB, as well as a concise reference of key
accomplishments and mission experience derived from the various mission
support roles. The primary focus of the FDAB is to provide expertise in
the disciplines of flight dynamics, spacecraft trajectory, attitude
analysis, and attitude determination and control. The FDAB currently
provides support for missions and technology development projects
involving NASA, government, university, and private industry.

**Table of Contents**

1.  Introduction

2.0 Flight Project Support

> 2.1 Development Missions

##### 2.1.1 Triana 

> 2.1.2 Space Technology (ST5)

## 2.1.3 Aqua Earth Observing System (EOS) 

> 2.1.4 Aura EOS
>
> 2.2 Operational Missions
>
> 2.2.1 Microwave Anisotropy Probe (MAP)

# 2.2.2 Earth Observing (EO) -1 

> 2.2.3 Geostationary Operational Environmental Satellite (GOES)
>
> 2.2.4 Landsat-4 Decommissioning
>
> 2.2.5 Tropical Rainfall Measuring Mission (TRMM)
>
> 2.2.6 InFocus Balloon Program
>
> 2.2.7 General Space Operations Management Office (SOMO) Support

3.0 Study Mission Support

## 3.1 Integrated Mission Design Center (IMDC) 

> 3.2 KRONOS
>
> 3.3 Constellation X
>
> 3.4 Global Precipitation Mission (GPM)
>
> 3.5 Next Generation Space Telescope (NGST)
>
> 3.6 Living With A Star: Solar Dynamics Observatory (SDO)
>
> 3.7 Magnetospheric Multiscale Mission (MMS)
>
> 3.8 Laser Interferometer Space Antenna (LISA)
>
> 3.9 Leonardo

4.0 Technology Development Activities

> 4.1 Advanced Mission Design
>
> 4.1.1 Generator Software Tool
>
> 4.1.2 SBIR Phase II Research: Optimal Orbit Transfer Analysis for
> Advanced Space Systems
>
> 4.1.3 Optimal Mission Design

# 4.2 Autonomous Navigation Technologies

## 4.2.1 Autonomous Onboard Navigation systems 4.2.2 Adaptive Kalman Filter for Autonomous Navigation

## 4.2.3 Magnetometer Based Navigation

> 4.3 Formation Flying Technologies

#### 4.3.1 EO-1 Formation Flying Experiment 

# 4.3.2 SBIR Phase II Research: Autonomous Unified Orbit and Attitude Control for Formation Flying Using GPS and LQG/LTR Controller 

> 4.3.3 Decentralized Estimation and Control of Distributed Spacecraft
>
> 4.4 Attitude Determination and Modeling Techniques
>
> 4.4.1 Attitude and Orbit Model Support
>
> 4.4.2 Advanced Attitude Determination and Sensor Calibration
>
> 4.4.3 Multi-Mission Attitude Determination System
>
> 4.5 Flight Dynamics Automation Studies

5.0 Branch Infrastructure

> 5.1 Flight Dynamics Tool Program
>
> 5.2 Flight Dynamics Lab

## 5.3 FDAB Web Page 

## 5.4 2001 Flight Mechanics Conference

6.0 Interagency Activities

## 

## 6.1 GSFC Standards Program

> 6.2 X-43A Anomaly Investigation

7.0 Outreach Activities

> 7.1 SAMPEX University Operations
>
> 7.2 Undergraduate Student Research Program (USRP)
>
> 7.3 PREST Program
>
> 7.4 Graduate Student Research Program (GSRP)
>
> 7.5 TEAMS Competition
>
> 7.6 Public Education/Community Outreach

Appendix A -- Awards

Appendix B - University Grants

Appendix C -- Conferences and Papers

Appendix D- Acronyms and Abbreviations

**[1.0 Introduction]{.underline}**

The Guidance, Navigation and Control Center (GNCC) at the Goddard Space
Flight Center (GSFC) provides the skills, vision and leadership in
guidance, navigation and control (GN&C) systems, engineering, operations
and mission analysis to enable revolutionary Earth and Space Science
discovery. The scope of technical disciplines encompassed by the GNCC is
broad and includes all aspects of flight dynamics, propulsion, flight
mechanics, guidance, navigation and control engineering for space
systems, experiments, and sub-orbital missions. The range of products
and services is also broad and requires expertise in skill areas such as
advanced component design, control system architecture, propulsion
design, trajectory analysis, autonomy and mission design.

Within the GNCC, the Flight Dynamics Analysis Branch (FDAB), Code 572,
is responsible for providing Guidance, Navigation and Control analytic
expertise for trajectory and attitude systems. This includes dynamics
and control analyses and simulations of space vehicles. The Branch
creates and maintains state-of-the-art analysis tools for mission
design, trajectory optimization, orbit analysis, navigation, attitude
determination, and controls analysis. The Branch also provides the
expertise to support a wide range of flight dynamics services such as
mission design, on-orbit sensor calibration, and launch/early orbit
operations. The FDAB also maintains an active technology development
program, with special emphasis on developing new techniques and
algorithms for autonomous orbit/attitude systems and advanced approaches
for trajectory design. Specific areas of expertise resident in the FDAB
are:

-   Attitude and trajectory analysis and control design

-   Control/Structure interaction analysis

-   Mission (attitude & trajectory) planning

-   Estimation techniques

-   Vehicle autonomy

-   Constellation analysis

-   Flight Dynamics model development

This document summarizes the major activities and accomplishments
performed by the FDAB in support of flight projects and technology
development initiatives in Fiscal Year (FY) 2001. The document is
intended to serve as both an introduction to the type of support carried
out by the FDAB, as well as a concise reference summarizing key analysis
results and mission experience derived from the various mission support
roles assumed over the past year. The FDAB staff that are involved in
the various analysis activities within the branch prepared this
document. Where applicable, these staff members are identified and can
be contacted for additional information on their respective projects.

Among the major highlights by engineers in the FDAB during FY2001 are:

-   ***Successful launch of EO-1 and demonstration of fully autonomous
    formation flying (with Landsat-7)***. Principal Investigators for
    this flight demonstration included David Folta and David Quinn of
    the FDAB

-   ***Successful flight dynamics support to the MAP mission.*** Branch
    engineers were responsible for attitude control system development,
    testing, early mission checkout and attitude sensor calibration. The
    FDAB was also responsible for trajectory design (to take the
    spacecraft to an L2 libration orbit) and trajectory operations.

-   ***Successful launch support for the GOES-M mission.*** Flight
    dynamics engineers planned and executed trajectory maneuvers to put
    this spacecraft on station.

-   ***Successful decommissioning of Landsat-4.*** FDAB engineers
    planned and executed orbit maneuvers to take Landsat-4 out of its
    operational orbit and assure reentry within 25 years.

-   ***Provisional patent application submitted for the MATLAB-ADS
    system.*** This generalized system can provide attitude
    determination and sensor calibration for both 3 axis stabilized and
    spinning spacecraft.

**[2.0 Flight Project Support]{.underline}**

This section summarizes FDAB support to Goddard flight projects during
FY01. For purposes of this report, these projects are classified as:

-   *Development Missions*: Approved missions under development.

-   *Operational Missions*: Missions that were in-flight in FY01. This
    includes missions that were in the final stages of development and
    were successfully launched in FY01 (e.g. MAP).

Support to future mission concept studies and proposal support for
missions seeking project approval are covered in section 3.

**[2.1 Development Missions]{.underline}**

######## **[2.1.1 Triana]{.underline}**

<http://triana.gsfc.nasa.gov/home/>

####### [Triana Trajectory Design]{.underline}

Flight Dynamics Analysis Branch (FDAB) personnel provided cost-effective
flight dynamics support for an April 2002 launch. This support included

-   automating the nominal trajectory generation

-   developing and refining correction maneuver strategies,

-   analyzing orbit maneuver propellant use,

-   analyzing initial acquisition and orbit determination errors, and

-   developing initial acquisition strategies and tracking schedules.

After Triana was removed from the STS manifest (i.e. launch postponed)
in March 2001, FDAB performed preliminary analysis of different Triana
launch options (e.g. STS at different inclinations, Delta ELV) and
presented a trajectory design review in September, 2001 at the request
of the Triana Project office.

[Triana Attitude Control System (ACS)]{.underline}

The Triana Attitude Control Systems Analysis team successfully completed
its final design review. The review covered the structure\'s flexible
modes and the modes' impact on controller performance and stability. All
requirements were met or exceeded, and the Triana Project Office was
satisfied with the controller's performance.

Extensive work was required for the implementation of the controller's
final design. Since examination of the finer aspects of controller
performance requires a real-time test environment, the team spent
considerable time planning and completing the flight software testing.
The team's final software build should be completed by October 2001.
Meanwhile, the integration and testing work required ongoing support,
especially in the phasing test area where two hardware wiring problems
were identified and successfully corrected. Also key to this year\'s
activities was the design, implementation, and test of the on-board
Failure Detection and Correction software and the ancillary contingency
planning and simulation. In preparation for the Triana spacecraft's
storage, the team has completed all action items and analysis reports.
The team also documented suggested design refinements.

[Triana Gyroscopic Upper Stage (GUS)]{.underline}

In addition to the trajectory design and ACS effort applied to the
Triana Observatory, the Flight Dynamics Analysis Branch (FDAB) is also
responsible for the end-to-end flight performance of the GSFC designed
and built Gyroscopic Upper Stage (GUS). The GUS is a multi-function
platform specifically designed for Triana. Two Marman clamp interfaces
attach the GUS to both Triana and the IRIS, an Alena-Spazio (*Italy*)
built spin table (which was originally used on the LAGEOS mission) that
is mounted across the Space Shuttle cargo bay. The major components that
comprise the GUS include a Thiokol built Star-48B solid rocket motor
(SRM), the Triana Event Sequencing System (TESS), the Nutation Control
System (NCS), and a dedicated power and telemetry system.

The mission time line begins with GUS spin-up in the Orbiter to roughly
60 RPM followed by a command to the pyrotechnic bolt-cutters that open
the first Marman clamp and allow the spring-driven actuators to eject
Triana. After Triana has cleared the IRIS and surrounding structures,
the NCS is enabled. Correct inertial pointing is passively maintained by
the gyro dynamics throughout one-half (47 minutes) of a 283 km altitude
orbit, while the Shuttle maneuvers to a distance and attitude protected
from SRM ignition. During the 84 second burn, the SRM provides a 3,175
meters per second perigee boost. The NCS continues to function through
powered flight in a second operational mode. Four minutes after burnout,
the GUS is separated from the Triana Observatory by triggering the
second Marman clamp and releasing additional actuators. The NCS is
inactive and its fuel reserves are passively depleted. The GUS is then
earmarked for orbital disposal.

The FDAB has provided analysis of all the major time line events
described above. These analyses include clearance analysis of the
ejection from the STS, flight dynamics analysis from STS ejection to
Observatory separation, design and analysis of the NCS, and
post-separation Observatory/GUS recontact avoidance.

The ejection analysis uses a multi-body, non-linear simulation to
determine the envelope of motion possible for the spacecraft as it exits
the IRIS cradle. The model incorporates twelve degrees-of-freedom, six
for a rigid model of the spacecraft structure and six for a damped
pendulum lumped-mass model of the on-board Hydrazine fuel. The family of
trajectories is continuously compared against a dynamic envelope of the
surrounding structure. STS safety criteria require that no recontact can
occur for any worst-case combination of up to two failures in either the
payload and/or Shuttle systems, for all dispersions of system
parameters. Analysis showed that these criteria are met. Milestones
completed during FY01 include completion of the Phase II Payload Safety
Review at Johnson Space Center and application of Monte Carlo techniques
to determine the statistics of the initial spacecraft state for orbit
propagation after ejection.

Since the Triana spacecraft spins about its minor principle axis, an
active NCS is required during the coasting phase of the mission. The NCS
maintains the proper orientation along the pointing (angular momentum)
vector in the face of environmental disturbances and internal energy
dissipation (e.g. fuel slosh). Since there is no feedback of the
inertial reference, the NCS must maintain attitude without disturbing
pointing while also minimizing propellant consumption (nitrogen
cold-gas). Additionally, during the powered flight phase of the mission,
the NCS is available to counteract the end-of-burn coning instability
known to exist with the Star-48 class of motors. In this second
operational mode, the NCS control strategy seeks to minimize response
latency and maximize control authority. The entire NCS design and
fabrication was a GNCC effort. Considerable analytical support was
provided by the FDAB in algorithm design, simulation, and testing. This
included the completion of a high fidelity dynamic simulator capable of
both off-line analysis and real-time hardware-in-the-loop testing with
the flight electronics modules. One of the more challenging aspects of
the simulator design was maintaining the 1 kilohertz clock cycle rate
used by the NCS micro-controller. By the end of FY01, the simulator will
have been used to support flight software testing, integrated system
testing of the NCS to the GUS, and stand-alone performance testing of
the flight NCS modules on a 3-axes rate table. Additional effort in the
NCS development includes completion of the Peer Review, preparation of
an Independent Verification and Validation (IV&V) data package, thruster
performance testing at Glenn Research Center, and the Phase II STS
Safety Review at Johnson Space Center.

Lastly, both analytical and numerical-based analyses have been used to
access the potential for recontact between the Triana Observatory and
GUS after the planned separation. Of primary concern regarding this
event are the clearance reducing effects for a flat (major axis) spin
separation, residual thrusting of the spent upper stage, and an NCS
thruster failure.

\[Technical contacts: Wendy Morgenstern, Greg Marr, Steve Queen\]

##### 2.1.2 Space Technology (ST5) *(launch 5/04*) 

[http://nmp.jpl.nasa.gov/st5/](http://nmp.jpl.nasa.gov/st5/%0D%0B)

[\
](http://nmp.jpl.nasa.gov/st5/%0D%0B)Space Technology 5 (ST5) is a
mission in the New Millennium Program and NASA's first experiment in the
design of miniaturized satellite constellations. The mission will last 3
months. During this time the constellation of three spin-stabilized
spacecraft will validate new technology for spaceflight. These
technologies include a miniature cold gas thruster, x-band transponder,
flexible interconnects, variable-emissivity coatings, a constellation
communications and navigation transceiver, ultra lower-power logic,
autonomous constellation management ground software, as well as, various
technology improvements embedded in the spacecraft itself. In addition
to validating these new technologies and instruments the mission goal is
to reduce the weight, size and cost of space missions, while preserving
or improving technical capabilities.

The planned orbit will be a Geostationary Transfer Orbit (GTO) with a
perigee as low as 270 km in altitude. Analysis has shown that lower
perigees run the risk of being lowered by lunar perturbations below 200
km in altitude, which must be avoided due to undesirable thermal
conditions and high aerodynamic disturbances caused by the increased
atmospheric pressure at lower altitudes. Ongoing flight dynamics
analysis is focused on maintaining the constellation. The spacecraft are
required to be separated by no more than 1000 km over a two hour window
centered at the apogee of the orbit.
![](media/image3.png){width="3.2083333333333335in"
height="1.6409722222222223in"}

The ST5 attitude control system (ACS) must provide an autonomous sun
acquisition mode as well as the capability to reorient the spin axis by
ground command.  The onboard ACS hardware consists of a sun sensor
mounted perpendicular to the spin axis, a three-axis magnetometer, and a
single cold gas thruster.  The cold gas thruster is mounted parallel but
offset from the spin axis in order to provide control torque as well as
translational acceleration.  The challenge was to provide an ACS that
uses simple algorithms to minimize onboard processing and work with the
limited sensor and actuator compliment.  A method that uses Rhumb line
precession (common on many spin stabilized spacecraft) to reorient the
spin axis will be implemented and tested in hi-fidelity simulations.
Nutation will be passively damped using a fluid filled ring damper that
is being designed and tested at NASA-GSFC.

Due to tight constraints of the relative orbits of the three spacecraft,
careful consideration must be used when planning attitude maneuvers,
which change the orbit while precessing the spacecraft spin axis due to
the translational and rotational components of the thrust vector. Monte
Carlo analysis is being used to determine optimal maneuver planning. The
results of this analysis will be used in a high fidelity simulation
including high order earth harmonics, solar and lunar perturbations, and
accurate atmospheric density models in order to determine the lifetime
of the constellation verify that the orbits meet the relative distance
requirement.

\[Technical contacts: [J. Morrissey]{.underline}, [M.
Woodard]{.underline}, [M. Concha]{.underline}\]

######## [**2.1.3 Aqua Earth Observing System (EOS)** (launch 3/02)]{.underline}

<http://aqua.gsfc.nasa.gov/>

The Flight Dynamics System (FDS) Team provided extensive pre-launch
support for the Aqua mission in the areas of attitude determination,
mission planning, and flight operations.

The FDS Attitude Team developed specifications for and tested software
to support real-time and off-line attitude determination, attitude
sensor calibration, and other product planning functions. The new
attitude utilities were tested with simulated data from the Aqua Flight
Operations Team (FOT). The Attitude Team created stand-alone utilities
for project simulation support to meet new Aqua mission requirements.

The FDS Maneuver Team performed various analyses and provided data to
support Aqua mission planning. In addition, the Team provided multiple
analyses and presentations to demonstrate the feasibility and benefits
of phasing Aqua with the EOS AM Constellation. Audiences included the
Aqua Project, Earth Science Mission Operation (ESMO) office, and
representatives of EOS PM Constellation constituents. The Maneuver Team
developed an initial Aqua ascent scenario and, in response to evolving
Aqua mission phasing requirements, re-planned the ascent several times.
Other analyses completed by the Maneuver Team include Tracking and Data
Relay Satellite (TDRS) antenna contact predicts from fairing jettison to
separation, TDRS/Polar Ground Station (PGN) contacts during early
mission phase, and mission planning product generation. The Maneuver
Team provided specification updates and tested the functions for mission
products planning. The Maneuver Team provided predict information used
in development of the Integrated Mission Timeline, provided inputs for
the Delta Detailed Test Objectives (DTO), and generated products for
internal GSFC and external Project simulations. In addition, the Team
produced simulator initialization parameters required for simulation and
testing support.

The FDS Team supported the Aqua FOT with frequent communications,
development of documents, and simulations with the Flight Dynamics
Facility (FDF). To support testing and to answer questions regarding FDS
mission support strategies, the FDS Team communicated frequently with
the FOT, Mission Readiness Test Team (MRTT), TRW (spacecraft
manufacturer), Raytheon (ground system contractor), and Aqua Project
personnel. Communication was both in person and via teleconference. An
Operations Procedures Handbook was started and other documents were
developed or updated. These other documents include the FDS Timeline,
the FDS/Aqua Project Interface Control Documents (ICDs), the FDS/FDF
Memorandum of Understanding (MOU), and Operations Agreements between FDS
and Earth Mission Operations System (EMOS), and between FDS and FOT. The
FDS Team assisted with FDS hardware installation and checkout in the EOS
Mission Operations Center (MOC) and defined and tested interfaces of FDS
with external entities. The FDS Team coordinated test support with FDF
personnel and performed simulations for FDS/FDF product exchanges.
Multiple project and internal simulations were also supported.

\[Technical contact: [D. Tracewell]{.underline}\]

## [2.1.4 Aura EOS *(launch 11/2003)*]{.underline}

<http://eos-aura.gsfc.nasa.gov/>

Aura's major science objective is the study of the chemical interactions
and climate change in the Earth\'s atmosphere, focusing on the upper
troposphere and lower stratosphere. The Aura spacecraft is 3-axis
stabilized and will operate in a near-circular, sun-synchronous polar
orbit at an altitude of about 705 km and an ascending nodal crossing at
approximately 1:45 PM mean local solar time. Pre-launch flight dynamics
services include mission design, sensor analysis, and operations
planning.

During FY2001, FDAB completed a preliminary ascent maneuver plan,
station-keeping maneuver analysis, and constellation flying analysis
with Aura, Aqua, and other spacecraft. FDAB also presented flight
dynamics material at the Aura Ground System Review, provided updates to
the Mission Specific Requirements Document (MSRD), refined
specifications for products, began development of preliminary Interface
Control Documents (ICDs), and participated in meetings of the Afternoon
Constellation Working Group. Support was also provided to the Aura
Project Scientist at the Aura Science Team meeting in Pasadena, and to
the Microwave Limb Sounder (MLS) and Tropospheric Emission Spectrometer
(TES) instrument teams. Attitude analysis was performed to determine the
best methods for computing attitude data at a frequency of eight Hz.

FDAB worked with the Aura Project Scientist to devise a scheme for
flying Aura relative to Aqua in order to obtain coincident imaging
between the two spacecraft. The desire is to coordinate observations
between the Aqua spacecraft instruments looking in the nadir direction,
and the Aura MLS looking forward at the Earth's limb. The idea is for
Aura to view a point on the Earth's limb a short time after Aqua has
flown directly over that point. Aqua will be flying on the World
Reference System (WRS) ground track. To accomplish the desired viewing
geometry, Aura will fly on an adjacent WRS path with a given offset such
that the Aqua ground track will always intersect the Aura MLS field of
view at the Earth's limb, as shown in Figure 2-2. Aura will follow Aqua
with an along-track separation between 15 and 22 minutes.

![](media/image4.wmf)**Figure 2-2. Relative Positions of Aqua and Aura
Showing MLS Viewpoint**

\[Technical contact: [L. Newman]{.underline}\]

**[2.2 Operational Missions]{.underline}**

# **[2.2.1 Microwave Anisotropy Probe (MAP)]{.underline}**

[http://map.gsfc.nasa.gov](http://map.gsfc.nasa.gov/)

####### [Background]{.underline}

FY 2001 has been an extremely productive and busy year for the MAP team
as we prepared for launch in the summer of 2001. MAP, shown in Figure
2-3, was successfully launched on June 30, 2001. At this time, MAP is
working extremely well, and its thermal stability is as good as the
science team had hoped.

MAP is a MIDEX-class mission produced by GSFC in partnership with
Princeton University. The goal of MAP is to produce an accurate full-sky
map of the cosmic microwave background temperature fluctuations
(anisotropy). This map will shed light on several key questions related
to the Big Bang theory and expand on the information provided by the
NASA Cosmic Background Explorer (COBE) mission. The MAP Mission lifetime
is two years with a goal of four years.

## ![](media/image5.png){width="6.0236220472440944in" height="4.099212598425197in"}Figure 2-3: The MAP Observatory

##### 

##### 

##### 

##### 

##### 

##### Pre-Launch Spacecraft Testing

Mission simulations were run in October 2000, and spacecraft
electro-magnetic interference (EMI) tests were run in early November.
All the components and the spacecraft as a whole passed the EMI tests.
One of the first deployment tests was done in late fall, and the backup
two-wheel controller design kickoff meeting was held in early November.
Also in November, the spacecraft was prepared for vibration and shock
testing, and scripts and procedures for performing maneuvers were
generated.

Throughout December 2000, the ACS Team worked hard to prepare for three
weeks of thermal vacuum testing followed by thermal balance testing
(TV/TB). Procedures were written, and a schedule, temperature profile,
and test script were written and refined many times over the course of
several months. Functional tests and comprehensive performance tests
(CPT) would be performed during TV/TB testing, so all our procedures had
to be written and debugged before the testing could begin. Most of
January was spent preparing for TV/TB. The testing was 24 hours a day
for three weeks, so everyone had to be familiar with every aspect of the
hardware, software, scripts, and procedures.

TV testing started January 24 and finished February 20. The thermal
profile tested on MAP was the most complicated ever done at Goddard.
Everyone involved worked hard, paid attention to details, and didn't try
to cut corners. The only issues were the failure of a survival heater on
Autonomous Star Tracker 1 (or AST A), and a question about reaction
wheel assembly (RWA) survival versus operational temperatures and
conditions. These were both successfully resolved, and the test finished
with no outstanding hardware issues. At the end of the testing, everyone
on the team had a much better feel for how the system as a whole worked.

####### [Trajectory Design]{.underline}

The MAP trajectory design team presented its work at several major
reviews, including peer reviews in January and March, and a red team
review in May. The team presented its work on the trajectory design,
maneuver operations, and navigation support. The review panels consisted
of several internal and external flight dynamics experts. By the red
team review, all action items from previous reviews had been fully
addressed and closed.

The MAP trajectory team also completed a significant amount of analysis
work. In particular, the trajectory team completed and delivered to
Boeing the nominal trajectories for the prospective launch months of
April, May, June, July, and August; as well as the launch window
analysis for the months of July and August. The team completed and
documented a number of contingency analyses as well as other trajectory
analysis requested by the Trajectory Peer review panel and by the MAP
project. A few of the contingency analyses completed and documented are
Using Mid-Course Correction (MCC) delta-V to remove L~2~ lunar shadows,
lunar shadow avoidance at L~2~, missed first perigee (P1) maneuver,
strategy to move perigee maneuvers into a station contact, and splitting
perigee maneuvers. In addition, the team completed an orbit
determination covariance analysis for the phasing loops and L~2~ phases
that was used to derive the orbit determination requirements and
tracking requirements for maneuver planning and calibration.

Moreover, the trajectory team built an analytic model for the phasing
loops. The model determines delta-V distribution across perigee
maneuvers in order to achieve the proper timing and energy. In addition,
the trajectory team performed a parametric study and a Monte Carlo
analysis to investigate the launch vehicle dispersions so as to
guarantee that trajectories that satisfy all mission requirements were
available for all dispersions and within the propellant budget. The team
developed a Matlab^®^ script to automatically evaluate the maneuver
execution errors.

##### Launch Preparations

Due to the lack of a separate flight ops team, the entire project team
held a Flight Ops Retreat in early April 2001. This got the whole team
focused on flight operations. Calibration burns and contingencies were
discussed throughout the spring. The ACS team also supported the
Pre-ship and Red Team reviews held in April 2001. Several Two-Wheel
Control Mode design reviews and code walk-throughs were held in the
spring for the ACS, project, and Flight Software teams.

The spacecraft was shipped to KSC in mid-April 2001. The ACS team spent
the last week of April at KSC doing post-ship testing, including a full
multi-day CPT and some pre-launch testing, both of which occurred around
the clock. The observatory made the trip with no problems, and all
procedures were completed successfully. Stray light problems with the
star trackers and the Coarse Sun Sensors (CSSs) were discovered after
the solar arrays were attached to the observatory. This was eventually
fixed by adding layers of black Kapton tape to the back of the solar
arrays and webbing, and around the CSS heads.

May 2001 brought the Operations Readiness Review and the Mission
Readiness Review. The GNCC also held its own internal Mission Readiness
Review. The Flight Readiness Review was held in June 2001. Also in June,
the MAP HiFi simulation was ported to the MAP SMOC. The benefit of
having this software readily available in the SMOC was shown many times
over when it was used for quick data analysis and verification. The
final functional tests were done before and after the spacecraft was
moved to the launch pad on June 20, 2001.

##### Maneuver Operations Planning

The Maneuver Ops Team was formed to address the problems of planning,
executing, and verifying all maneuvers. It consisted of representatives
from trajectory design, ACS, propulsion, flight software, spacecraft
controllers, and navigation disciplines.

The Maneuver Team generated information flow diagrams and developed
processes to plan, execute, and verify maneuvers. These processes were
demonstrated during several maneuver simulations. Based on the
simulations, the processes were refined and updated. The Maneuver team
was also instrumental in setting up for these simulations, evaluating
anomalies, and verifying the performance of the maneuvers. The team
identified interfaces between subsystems, software required to plan
maneuvers, and data and file formats. An Interface Control Document
(ICD) was written to document the exchange of products within the
different sub-teams. In addition, various procedures were written and
implemented as a result of the many simulations held. To properly model
the effects of thruster firings on the fuel mass and thereby increase
the accuracy of the simulations, the team added a propulsion blow down
model to the Hybrid Dynamic Simulator (HDS) and to the ACS HiFi
simulator. Moreover, propellant budgets were produced many times in
support of many reviews during the year.

##### Post Launch

MAP was launched at 19:46:46 Z from the Cape Canaveral Spaceport aboard
a Delta II 7425 expendable launch vehicle. The Delta vehicle placed the
MAP spacecraft into a highly elliptical parking orbit with a 28.7°
inclination and a 185 km perigee altitude. Transfer from the parking
orbit to the mission orbit started in July and will end in December (see
Table 2-1).

The transfer consisted of phasing loops followed by a lunar gravity
assist (see Figure 2-4). The FDAB Trajectory Team successfully planned
and calibrated all phasing loop apsis maneuvers leading to a nominal
lunar swing-by. The MAP mission orbit will be a Lissajous orbit about
the L~2~ Sun-Earth Lagrange point, approximately 1.5 million km from
Earth in the anti-Sun direction. This location and orbit were selected
to minimize environmental disturbances and maximize observing
efficiency. At L~2~, the spacecraft will maintain a Lissajous orbit such
that the MAP-Earth vector remains between 0.5° and 10.5° off the
Sun-Earth vector to satisfy communications requirements while avoiding
eclipses. Four station-keeping maneuvers per year are planned to
maintain the Lissajous orbit. MAP will be the first spacecraft to orbit
about the L~2~ Lagrange point for a period of two years.

![](media/image6.png){width="5.7555555555555555in"
height="2.8055555555555554in"}Figure 2-4: Representation of the MAP
3-Loop Trajectory

## Table 2-1: MAP Maneuvers

+-----------+------------------+-----+----+---------+--------+------+
| Maneuver  | ### Description  | Ma  |    |         | D      | ##   |
| Location  |                  | neu |    |         | elta-V | # St |
|           |                  | ver |    |         |        | atus |
|           |                  | St  |    |         | (m/s)  |      |
|           |                  | art |    |         |        |      |
|           |                  |     |    |         |        |      |
|           |                  | #   |    |         |        |      |
|           |                  | ##  |    |         |        |      |
|           |                  | UTC |    |         |        |      |
|           |                  |  Da |    |         |        |      |
|           |                  | te  |    |         |        |      |
|           |                  | & T |    |         |        |      |
|           |                  | ime |    |         |        |      |
+-----------+------------------+-----+----+---------+--------+------+
| A1        | Calibration of   | J   | 04 | 1       | 2.01   | *    |
|           | Thrusters        | uly |    | 3:22:38 |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
| P1        | Apogee Raising   |     | 08 | 0       | 20.19  | *    |
|           |                  |     |    | 4:43:40 |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
| A2        | Engineering Burn |     | 12 | 1       | 0.25   | *    |
|           |                  |     |    | 6:11:54 |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
| P2        | Apogee Raising   | []{ | 17 | 0       | 2.51   | *    |
|           |                  | .ma |    | 3:38:25 |        | *** |
|           |                  | rk} |    |         |        |      |
+-----------+------------------+-----+----+---------+--------+------+
| A3        | Engineering burn |     | 21 | 1       | 0.30   | *    |
|           |                  |     |    | 8:54:43 |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
| P3        | Apogee Raising   |     | 26 | 10:29   | 7.35   | *    |
|           |                  |     |    |         |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
|           | P3 Correction    |     | 27 | 04:30   | 0.31   | *    |
|           |                  |     |    |         |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
| Moon      | Gravity Assist   |     | 30 | 16:37   | #      | *    |
|           |                  |     |    |         | ## N/A | *** |
+-----------+------------------+-----+----+---------+--------+------+
|           | MCCM 1           | Aug | 06 | 16:37   | 0.09   | *    |
|           |                  |     |    |         |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
|           | MCCM 2           | S   | 14 | 16:37   | 0.04   | *    |
|           |                  | ep. |    |         |        | *** |
+-----------+------------------+-----+----+---------+--------+------+
| L2        | LOI              | D   | 14 | 16:37   | #      | TBD  |
|           |                  | ec. |    |         | ## TBD |      |
+-----------+------------------+-----+----+---------+--------+------+

> A1, A2, A3 = First, Second, and Third Apogee
>
> **** = Successfully Completed

#### P1, P2, P3 = First, Second, and Third Perigee

> MCCM = Mid-Course Correction Maneuver

###### LOI = Lissajous Orbit Insertion

The MAP ACS team had an incredibly busy month of July. The initial
separation and sun acquisition went well. Over the course of a grueling
first week or so, team members worked around the clock to check out the
spacecraft's Safe hold modes, inertial modes, and all the thrusters.
Calibration pulses were commanded, and sensor calibration slews were
performed. Some real-time issues were resolved quickly and efficiently,
with the primary goal being spacecraft and orbit safety. A rough
alignment and gyro calibration was performed in the first week, with the
only glitch being an incorrect velocity aberration correction to the AST
data onboard. All the sensors and actuators worked as advertised, and
all the flight software worked as it was designed. Additional work done
using the high fidelity (HiFi) simulator in the Science Mission
Operations Center (SMOC) was a calibration of the wheel tachometers to
reduce the error in the system momentum calculation while the
observatory is spinning, and a recalculation of the CSS eye outputs to
better model the actual output of the heads.

The ACS team had to remain flexible to deal with things like moon and
earth interference in the star trackers (this prevented a full Observing
Mode), modified thruster calibration maneuvers, and an "anomalous force"
on the spacecraft near perigee that increased the system momentum just
before a maneuver. Analysis eventually showed that the anomalous force
was due to the baking off of outgassed volatiles, and it was properly
modeled and predicted for subsequent perigee and perilune passes. Also
during July, the Real Time Attitude Determination System was used to
generate true attitude knowledge and gyro bias information. The ground
analysis shows that we are easily meeting our 1.3 arc minute attitude
determination performance. The sensor calibration went well, with
residual errors on the order of tens of arc seconds. All of the
engineers worked hard to capture all the flight data, and feed back
attitude and maneuver information to the rest of the team as quickly as
possible, aided by the availability of the MAP HiFi in the SMOC. Often,
the trajectory team performed a quick first-order verification of a
maneuver within a half hour of the actual burn.

##### Conclusions

The remaining work includes finishing the coding and testing of the
two-wheel controller (Observing II) so that the mission can be
accomplished even if one wheel fails. As part of that process, Scott
Starin presented his Professional Intern Program (PIP) project on the
analysis and simulation of the two-wheel controller. The Maneuver Team
continues to support all maneuvers, including planning, generating
command sequences, simulation, and validation.

Thank you to all the people on all the teams who made this mission such
a success. Thanks should also be given to the people in the GNCC who
supported our critical operations by maintaining the computer "Nichols"
and working around our hectic schedule, often on short notice. Again, a
hearty thanks and congratulations to all who have made MAP a success.

\[Technical Contacts: [O. Cuevas]{.underline}; [S.
Andrews]{.underline}\]

##### 

##### 2.2.2 Earth Observing --1 (EO-1)

<http://eo1.gsfc.nasa.gov/miscPages/home.html>

In November 2000, the Goddard Space Flight Center launched the Earth
Observing -1 (EO-1) spacecraft. The EO-1 mission was dedicated to
testing a wide range of spacecraft subsystem and operational
technologies, including the next generation of earth sensing instruments
that may fly on future Landsat spacecraft. Among the technologies that
had a direct bearing on future Flight Dynamics support was a
demonstration of formation flying. EO-1 and the Landsat-7 spacecraft
operate in a close formation that keeps the two spacecraft about one
minute apart along track. The goal of operating in this formation was to
have each spacecraft take a series of co-fly images of the same scene
approximately one minute apart. The images could then be compared to
measure improvements in the EO-1 imaging instruments. See section 4.3.1
for more details on the formation flying experiment.

The FDAB led the Flight Dynamics team that supported the EO-1 Project
through four years of pre-launch preparations. Extensive mission
analysis was performed to tailor the EO-1 launch for a window only
seconds in duration. This analysis was necessitated by the stringent
formation flying requirements, launch vehicle dispersions, and a very
limited propellant budget. Following launch on a Delta II 7320-10
vehicle, the EO-1 Flight Dynamics team supported nine orbit maneuvers
over a four week period to attain the mission orbit. In addition to
orbit maneuver support, the FDAB provided support for sensor calibration
and alignment, real time attitude computation, orbit determination, and
generation of orbit and attitude products for scheduling and image
planning. About two months after launch, the FDAB turned over the Flight
Dynamics portion of the EO-1 ground system to accomplished Flight
Dynamics analysts who were members of the EO-1 Flight Operations Team.

What had been developed as a one-year mission to take 200 co-fly images
between EO-1 and Landsat-7 was successfully completed in about six
months. Through careful analysis of the launch window and the ascent
maneuver sequence, the limited propellant budget was conserved;
therefore, the mission can be extended for several years.

\[Technical contacts: [R. DeFazio]{.underline}, [R.
Luquette]{.underline}, [C. Mendelson]{.underline}\]

**[2.2.3 Geostationary Operational Environmental Satellite
(GOES)]{.underline}**

<http://www.gsfc.nasa.gov/topstory/20010620goesm.html>

GOES-M was the fifth and last of a series of geosynchronous
meteorological satellites built and launched under NASA/GSFC supervision
for NOAA. On July 23, 2001 GOES-M was launched aboard an Atlas IIA
launch vehicle into a geosynchronous transfer orbit of 42271 km. x 275
km. with an inclination of 20.54 deg. The launch was nominal and the
Flight Dynamics team, led by FDAB personnel, prepared to execute a
seven-orbit maneuver sequence to place GOES-M into its checkout orbit.
All went as planned until about 25 % of the way into the first Apogee
Motor Firing (AMF). At that point the temperature on the Main Satellite
Thruster (MST) exceeded the maximum allowable value. The maneuver was
aborted at 13 minutes into a 54.5 minute burn. After studying the
situation, a second AMF was planned several days later for a maximum of
12 minutes. This second maneuver aborted just four to six seconds short
of its completion. From that point onward, all GOES-M orbit maneuvers
with the MST were limited to less than 11 minutes. No further aborts
were encountered, but the limitation on burn size required a total of
nine AMFs instead of the nominal three AMFs of the pre-launch plan. Five
additional burns were required to adjust the final orbit with the entire
station acquisition sequence taking 29 days instead of the nominally
planned 18 days. Except for the problems with the MST, the spacecraft
proved to be healthy.

The Flight Dynamics team provided fine support to the GSFC GOES Project
during five launches dating from 1994 to the present. This was a
dedicated team that worked tirelessly through both good times and bad.
This may be the swan song for GOES Flight Dynamics support at GSFC,
since the next GOES series will be launched and placed on station by the
spacecraft contractor. This saddens many of us who have watched GSFC
Flight Dynamics provide quality support to the GOES Project for more
than a quarter century. All good things must eventually end and thus it
is with GOES Flight Dynamics at GSFC; however, the people who have
supported GOES missions will make their presence felt on many other GSFC
missions.

\[Technical contact: [R. DeFazio]{.underline}\]

######## **[2.2.4 Landsat-4 Decommissioning]{.underline}**

The Landsat 4 (L4) spacecraft was funded and launched by the U.S.
government and operated by Space Imaging (formerly known as the Earth
Observation Satellite Company). In late February, Space Imaging notified
the government that they no longer intended to operate the L4
spacecraft; therefore, they transferred L4 to the USGS for
decommissioning. In support of the decommissioning, the FDAB was asked
to design and execute the maneuvers necessary to comply with NASA
guidelines for disposal.

Several days of maneuvers lowered L4's altitude from 705 km to 580 km.
At its original altitude, L4 would have entered the Earth's atmosphere
after 42 to 75 years. At the lower altitude, L4 will enter the
atmosphere after 8 to 25 years which satisfies the NMI guidelines for
decommissioned spacecraft to enter within 25 years.

In addition to the orbit lowering, on-board energy sources, such as
propellants and batteries, were depleted. This satisfied other NMI
guidelines based on preventing an accidental explosion, which would
create a debris field in space.

## [**http://www.earth.nasa.gov/history/landsat/landsat4.html**](http://www.earth.nasa.gov/history/landsat/landsat4.html) {#httpwww.earth.nasa.govhistorylandsatlandsat4.html .sub-heading}

<http://www.gsfc.nasa.gov/gsfc/gnews/071301/071301.htm#landsat>

<http://www.usgs.gov/public/press/public_affairs/press_releases/pr1455m.html>

\[Technical contact: [D. Quinn]{.underline}\]

**[2.2.5 Tropical Rainfall Measuring Mission (TRMM)]{.underline}**

<http://www.gsfc.nasa.gov/news-release/releases/2001/01-84.htm>

In July 2000, Flight Dynamics Analysis Branch (FDAB) personnel began
preliminary analysis related to TRMM deorbit planning. This analysis
indicated early on that the 68 kg of propellant budgeted for the deorbit
operation would be inadequate to conform to the requirements and
guidelines imposed by NASA Management Instruction (NMI) 1700.8 and NASA
Safety Standard (NSS) 1740.14 for spacecraft end-of-life disposal.
Detailed analysis performed during the first half of 2001 determined
that a controlled deorbit from TRMM's 350 km altitude could be performed
in a manner which satisfied the requirements and guidelines with as
little as 158 kg of propellant, 90 kg more than was budgeted. The
additional propellant would be taken from the orbit maintenance
propellant budget, effectively shortening the mission life by 15 to 17
months. FDAB personnel presented these results at a TRMM deorbit status
review in April 2001.

A recommendation was made to assess the operational and science impact
of raising TRMM's mean orbit altitude to 400 km. Raising the orbit
altitude would decrease the aerodynamic drag thus reducing the amount of
propellant required for orbit maintenance. Analysis indicated that this
had the potential of extending the mission life by 4 to 5 years. By May
2001, raising TRMM began to look like a viable option; the deorbit
planning was put on hold; and work was focused on the analysis and
planning required for raising the orbit. Simulations were run to ensure
that the TRMM attitude control system, designed for 350 km, would
continue to meet the pointing requirements at 400 km. The orbit raising
took place during August 2001 and the spacecraft is now operating at a
mean altitude of 402.5 km. A design flaw was discovered in the Earth
sensors at the higher altitude. Because of this flaw, TRMM is using the
back-up attitude determination algorithms, which use the three-axis
magnetometer and digital sun sensors.

The focus is now turning back to the deorbit plan. Over the next several
months, the plan will be modified to account for the new orbit altitude
as well as operational changes that stem from the orbit raising.
Presently, the plan is to let drag decay the orbit to 350 km before
starting the burn sequence to deorbit TRMM. The final version of the
plan is currently expected to be completed by December 2001.

\[Technical contacts: [F. Vaughn]{.underline}, [J.
Morrissey]{.underline}\]

**[2.2.6 InFocus Balloon Program]{.underline}**

<http://infocus.gsfc.nasa.gov/>

# Over the past year, the Flight Dynamics Analysis Branch continued its engineering support to develop and operate pointing control systems for use on stratospheric balloon payloads. All efforts were directed to bringing to the balloon community significant new pointing capabilities for stratospheric ballooning.

# A primary goal of the science ballooning community is to achieve arc second pointing on X-ray target sources. A major impediment to precision pointing on balloon payloads is caused by disturbances originating from the cable and parachute segment connecting the gondola to the balloon. This segment, called the load train, is 250 ft long. To characterize load train disturbances, GNCC attitude sensors were included on a test flight of the InFocus telescope in August 2000. InFocus is a long focal length (9 meter) imaging X-ray telescope designed as a payload for stratospheric balloons. The measurements of the load train and gondola motions clearly showed that disturbance oscillations occur at several frequencies, and that these disturbances last for hours after the balloon reaches its float altitude.

A second InFocus flight was launched on July 5, 2001 from Palestine,
Texas. An existing azimuth/elevation gondola pointing control system was
adapted to point the telescope with a desired accuracy of several arc
minutes. The GNCC was responsible for establishing the slewing and
pointing control parameters and also the azimuth reaction wheel size
necessary to maneuver the large telescope. The only practical approach
for establishing these values was to conduct pointing tests in high bays
at GSFC and in Palestine, Texas. The entire gondola was suspended from
the ceiling and allowed to maneuver in azimuth and elevation in
conditions as close as practical to the actual flight configuration.
Several schemes, both active and passive, were tested in an effort to
increase damping. In addition to the tests, simulations of the elevation
and the azimuth control loops were generated to further predict what the
performance would be when the gondola was suspended from the much longer
load train.

![](media/image7.bmp)

From the second InFocus flight it was learned that the azimuth/elevation
systems are easily overpowered by high winds; thus, more robust
approaches are needed if arc second accuracy is ever to be obtained. One
approach now being considered, which was proposed by the science team,
is to support the InFocus telescope on the gondola base via a ball and
cup having three degrees of rotational freedom. Another lesson from the
InFocus flight is that suspended pointing tests, though necessary, can
be misleading in judging the performance of the pointing system in
flight. This reinforces the need for simulations, which require
aerodynamic models based on flight data. Several ideas to collect this
data on a regular basis are being pursued including compact measurement
modules that could monitor the local rotational motion of where ever
they are placed on the balloon, gondola, or load train.

\[Technical contact: [D. Olney]{.underline}\]

**[2.2.7 General Space Operations Management Office (SOMO)
Support]{.underline}**

The Space Operations Management Office (SOMO), located at Johnson Space
Center, was an important sponsor and funding source in FY01 for many of
the FDAB activities. This includes much of the branch's technology work
(covered in Section 4) as well as general mission design and concept
development work for future missions (some of the work covered in
Section 3 is sponsored by SOMO). The FDAB periodically assisted SOMO in
its management of mission services and operations activities, including
its management of the Consolidated Space Operations Contract (CSOC).
FDAB management meet regularly with CSOC management responsible for the
operation of the Flight Dynamics Facility. The purpose of these meetings
is not to give direction to routine operations, but to continue to
maintain awareness of facility upgrade plans and share knowledge of
future mission plans, technology development activities relevant to the
facility, and software system upgrades.

\[Technical Contact: [Tom
Stengle](mailto:thomas.stengle@gsfc.nasa.gov)\]

**[3.0 Study Mission Support]{.underline}**

One of the primary roles of the Flight Dynamics Analysis Branch (FDAB)
within the Guidance, Navigation and Control Center (GNCC) is to serve
the science community by providing analysis of advanced mission
concepts. This includes development of orbit/attitude designs based on
science constraints, evaluation of orbit/attitude errors and attitude
dynamics analysis. Members of the branch often represent "first access"
by Earth Science and Space Science customers to the services offered by
the GNCC.

In FY2001, the FDAB continued its participation in supporting a wide
variety of future mission concepts. This section describes some of the
analyses performed.

**[3.1 Integrated Mission Design Center (IMDC)]{.underline}**

The Integrated Mission design Center (IMDC) is a human and technology
resource dedicated to innovation in the development of advanced space
mission design concepts to increase scientific value for NASA and its
customers. The IMDC provides specific engineering analysis and services
for mission design and provides end-to-end mission design products. For
information about the IMDC, refer to
[http://imdc.gsfc.nasa.gov/]{.underline}.

Trajectory engineers from the FDAB supported the IMDC's customers by
providing mission planning, and trajectory analysis and design. Attitude
determination and control (ACS) personnel supported IMDC in the area of
ACS conceptual design and analysis. This included ACS requirement
definition, identification and computation of significant worse case
disturbance torques, sensor selection, actuator sizing, component
placement specification, control modes design, identification of ACS
imposed requirements on other subsystems, risk assessment, issues,
concerns and future work identification. In addition, special design
consideration and analysis were performed to solve or pinpoint each
mission's unique issue.

A wide range of mission types was supported, including near earth which
focus on the earth's surface and atmosphere and cis-lunar that are
interested in the earth's space environment. Specific studies supported
were: Ocean Observer, NPOESS, Geospace Electrodynamic Connections,
Carbon Cycle Initiative (Lider, BIRCH), GradSat, Gas & Aerosol
Monitoring Sensorcraft Mission (GAMS), Carbon, Black Carbon, Radiation
and Aerosols (COBRA), Survey of Infrared Cosmic Evolution (CIRCE), X-ray
Mirror Array Sky Survey (XMASS), Magnetospheric Tail Constellation (MTC
or name MAGIC), Constellation for Aerosols and Cloud Heights (COACH),
ASA Space Science, Super Nova, Stella Image, Global Precipitation
Mission (GPM), ST-7 and ACCESS.

(Technical contact: Josephine San, Charles Petruzzo, Aprille Ericsson,
Marco Concha, Eric Stoneking)

**[3.2 KRONOS]{.underline}**

The FDAB continued work on an extensive mission feasibility study for
High Earth Orbit (HEO) missions such as KRONOS. These unique orbits
require a lunar swingby to increase the orbit perigee radius, lift the
orbit out of the ecliptic plane, and rotate the line of apsides such
that apogee is in the northern hemisphere. The final HEO orbit obtained
via this lunar swingby has perigee near 10 Earth Radii (R~E~) and apogee
near lunar distance ( 60 R~E~). This study, whose results were
presented at the 2001 Flight Mechanics Symposium, contains a detailed
analysis of HEO orbit characteristics, launch window opportunities, and
fuel budget estimates. It is anticipated that the paper presented,
"*High Earth Orbit Design for Lunar-Assisted Medium Class Explorer
Missions*," will be used as part of future mission proposals for KRONOS
and other HEO missions. The KRONOS trajectory is shown in Figures 3-1,
3-2, and 3-3.

![](media/image1.wmf)

\[Technical contact: [Steven
Cooley](mailto:steven.cooley@gsfc.nasa.gov)\]

##### 3.3 Constellation X

![](media/image2.wmf){width="3.0in" height="2.379166666666667in"}
provided support to the Constellation X study team in the area of
trajectory design. The baseline plan is to launch the spacecraft aboard
an Atlas V launch vehicle. The FDAB performed a study to consider the
pros and cons of two mission orbit types that would provide
low-radiation environments. The baseline trajectory is an orbit about
the Sun-Earth L2 libration point -- this was contrasted against a High
Earth Orbit (HEO) with apogee at lunar distance to determine differences
in shadows and delta-v costs.

Also, in the event that the Atlas V launch vehicle is not available in
the 2010 mission timeframe, the FDAB completed a study to determine the
feasibility of launching the spacecraft on Delta II vehicles, then using
a ![](media/image3.wmf){width="3.2161417322834644in"
height="2.425196850393701in"}combination of hydrazine and low-thrust
propulsion to achieve the desired mission orbit. Various low-thrust
options were investigated for the Constellation-X mission using a Solar
Electric Propulsion (SEP) system and/or a hydrazine or Bi-Propellant
system. The SEP was sized to meet orbit requirements and attain the L2
mission orbit within one year. Initial orbit options covered a range of
LEO and GTO orbits. Thrust levels were analyzed which provided the
required trip time. To obtain a propulsion system thrust level needed to
be on the order of 300mN. The propulsion and launch vehicle support
groups then used these trajectory designs to define their requirements
and cost.

A sample of the transfer trajectory for this type of propulsion system
is shown in Figure 3-5.

\[Technical contact: [Lauri Newman](mailto:lauri.newman@gsfc.nasa.gov),
Dave Folta\]

**[3.4 Global Precipitation Mission (GPM)]{.underline}**

![](media/image4.wmf){width="3.0in" height="2.1756944444444444in"}The
main objective of the Global Precipitation Measurement (GPM) mission is
to provide sufficient sampling to reduce uncertainty in short-term
rainfall accumulations in a coverage region between 70º N and 70º S
latitude, with a goal being a 3-hour revisit time. The strategy for
achieving this objective is to form a constellation of
radiometer-carrying satellites using both GPM-specific satellites
(Drones) and other satellites with suitable radiometers (Co-op). The
Co-op satellites are in defined, fixed orbits, while the Drone
satellites are placed in orbits to best complete the desired coverage.

A significant portion of

FDAB effort this year has been related to the development and exercise
of a spacecraft constellation optimization tool. Given that some of the
s/c are dedicated to other missions and already have their orbits
defined, the question becomes: "Where do we place the GPM Drones to best
(or optimally) achieve maximum coverage?" Then, of course, we have to
ask ourselves how we're going to deploy and maintain the constellation.
It's a problem that becomes increasingly complex .

The optimization tool developed to attack this problem is centered
around a Genetic Algorithm (GA) approach. This is a technique that is
based on natural selection and Darwinian genetics, in that it is
population-based and progresses toward survival of the "fittest"
solution. One form of output from this tool is a graphical display of
coverage regions superimposed on a world map. Since the mission
objective is to map the Earth's surface every 3 hours, one approach
we've looked at is dividing each day into 8 3-hour "bins." If a point on
the Earth grid has been visited by one or more of the constellation
sensor footprints during one of these periods, it is said to occupy that
bin. An ideal solution would be to have each point in the study region
occupy each bin each day of the mission life.

![](media/image5.jpeg){width="3.38125in" height="2.45625in"}

In addition to the constellation optimization effort, FDAB personnel
have performed analyses on: ∆V for insertion and maintenance; end of
mission life re-entry/disposal; ground station coverage; launch
scenarios; sensor calibration opportunities; ground-track repeat cycles;
frozen orbit options; and a GPM Core spacecraft-to-ISS range study.

\[Technical contacts: [Chad
Mendelsohn](mailto:chad.mendelsohn@gsfc.nasa.gov), [David
Folta](mailto:david.folta@gsfc.nasa.gov)\]

###### 

###### 

###### [3.5 Next Generation Space Telescope (NGST)]{.underline}

The Next Generation Space Telescope (NGST) is slated to replace the
Hubble Telescope at the end of this decade. The scientific goals for
NGST are discovering and understanding the formation of the first stars
and galaxies, the evolution of galaxies and the production of elements
by stars, and the process of star and planet formation.

![](media/image6.jpeg){width="3.408333333333333in"
height="3.747916666666667in"}Among the various mission/design
architectures being considered for NGST, a proposed architecture, which
is based on a non-contacting payload and support modules, was
investigated. Linear and nonlinear simulation models were validated and
used to verify nominal fine pointing performance (LOS stability),
stability margins, and slewing times. Both, time-domain and
frequency-domain analysis were used in this verification. A Sensitivity
analysis with respect to the variations in the plant and control
parameters was performed. A host of parameters were included in the
sensitivity analysis, such as mass and inertia properties of the support
and payload modules, flexible modal frequencies and damping, geometric
properties, sampling time and delays, and many more. An overall
assessment of the feasibility of the proposed architecture was made
based on the results of the aforementioned analyses.

The NGST mission is baselined as a mission located near the Sun-Earth L2
Libration point. While the final orbit configuration is still being
determined, several analyses have been performed and project reviews
attended to discuss and provide feedback to potential spacecraft
contractors. Analysis by the FDAB was concerned with the effects of the
large solar shade and location of thruster placement in relationship to
orbit control. Using recent developments in dynamical system theory and
its implementation into FDAB software, the analysis focused on biased
libration orbits and single axis control. This trajectory design
analysis addressed improved methods for attaining constrained orbit
parameters and their control at the exterior collinear libration point,
L~2~. The use of a dynamical systems approach, state-space equations for
initial libration orbit control, and optimization to achieve constrained
orbit parameters were emphasized. The NGST trajectory design encompasses
a direct transfer and orbit maintenance under a constant acceleration. A
dynamical systems approach can be used to provide a biased orbit and
stationkeeping maintenance method that incorporates the constraint of a
single axis correction scheme. This analysis was performed in
partnership with Purdue University and showed that several strategies
are possible including one with a central manifold design that
incorporated pre-defined maneuvers in a given axis, and another that
incorporates the acceleration of the solar shade as a solar sail. Two
options shown in Figure 3-9 show that NGST control requirements can be
met. These are taken from a paper titled "Trajectory Design Strategies
for the NGST L~2~ Libration Point Mission" presented at the AAS/AIAA
conference in February, 2001. The left figure shows central manifold
maintenance. For this case, the accelerations can optionally be included
in the dynamical systems approach that compute a baseline libration
orbit or included in the targeting procedure afterwards. The orbital C3
energy is maintained below zero and return to the Earth region. The
figure on the right presents a biased orbit with deterministic control.

![](media/image7.wmf){width="3.3986111111111112in"
height="2.5493055555555557in"}

![](media/image8.wmf){width="2.745138888888889in"
height="2.395138888888889in"}

\[Technical Contact: Peiman Maghami, David Folta \]

####### **[3.6 Living With A Star: Solar Dynamics Observatory (SDO)]{.underline}**

The Flight Dynamics Analysis Branch (FDAB) has been involved in the
formulation of missions for the Living with a Star (LWS) Program. The
FDAB was involved with the generation of orbits and products related to
a distributed system of spacecraft used to understand the Earth's
environment and the interaction with the sun. The LWS's Solar Dynamics
Observatory (SDO) mission concept was investigated by the FDAB.

![](media/image9.wmf){width="6.0in" height="3.0909722222222222in"}Solar
Dynamics Observatory is a solar observer performing continuous and high
cadence observations of the full solar disk and coronal imaging in
multiple wavelengths to improve understanding and forecasting of the
Sun's impact. FDAB personnel investigated various orbit design
considerations from LEO and GEO to Highly Elliptical to Libration
orbits. These orbit designs considered the general aspects of these
designs and included shadow analysis, station coverage, Delta-V and fuel
budgets, sun related orbit parameters, and onboard antenna coverage
patterns. From this analysis, a geosynchronous inclined orbit was
chosen. Figure 3-10 depicts one final orbit consideration for the
geosynchronous inclined orbit option.

\[ Technical Contact: David Folta\]

**[3.7 Magnetospheric Multiscale Mission (MMS)]{.underline}**

MMS is part of the Sun-Earth Connection program, one of the four
principal science themes of NASA's Office of Space Sciences. The major
focus of the Sun-Earth Connection program is investigating the physical
processes that link the Sun and the Earth. MMS is a four-spacecraft
solar-terrestrial probe designed to study magnetic reconnection, charged
particle acceleration, and turbulence in the key boundary regions of the
Earth's magnetosphere. A draft version of an Announcement of Opportunity
for the instrument complement and principal investigator teams is
expected to be issued in October of 2001.

The mission is in its study phase. As far as was possible without the
Principal Investigator teams having been selected, a statement of
trajectory requirements was developed by the orbit analyst in
consultation with MMS study management. The analysis effort is not
complete, but much about the orbit dynamics of the mission has been
characterized. The mission is divided into distinct phases, each of
which uses a trajectory significantly different from the others.
Analysis related to the transitions between and within the phases is
highly complex and some techniques have yet to be developed. We have
demonstrated, in a gross sense, that trajectories can be designed to
meet most of the known science and engineering requirements and have
estimated the amount of propellant required.

A major consideration in our ability to do the necessary analysis is the
availability of software appropriate for the effort. Little of it is in
the "off the shelf" category because this is not a routine mission for
which analysis techniques are readily available. Though lacking some
major features, prototype software finds and analyzes trajectories
similar to what MMS would use so that engineering and science
characteristics can be considered. A graphics program, in prototype
form, illustrates the behavior of the tetrahedron as it changes size and
shape throughout each orbit.

For more detailed information about the mission, visit
[http://mms.gsfc.nasa.gov](http://mms.gsfc.nasa.gov/).

\[Technical contact: Charles Petruzzo\]

**[3.8 Laser Interferometer Space Antenna (LISA)]{.underline}**

The primary objective of Laser Interferometer Space Antenna (LISA)
mission is to detect and measure gravitational waves from massive black
holes and galactic binaries in the frequency range of 10^--4^ and 0.1
Hz. The LISA mission comprises three identical spacecraft, 500,000 km
apart, which form an equilateral triangle (Figure 3-11). The center of
the spacecraft formation is in the ecliptic plane, 1 AU from the Sun and
20° behind the Earth. LISA can essentially be viewed as a Michelson
interferometer in space, with a third arm to provide wave polarization
information as well as redundancy. Each spacecraft contains two optical
assemblies, with each assembly pointing towards an identical assembly on
each of the other two spacecraft. A 1 W infrared laser beam (1 μm
wavelength) is transmitted to the remote spacecraft via a telescope. The
incoming beam is focused on a sensitive photodetector where it is
superimposed with a fraction of the original local light. Each optical
assembly includes an enclosure containing a free-flying proof mass,
which serves as an optical reference mirror for the light beams. A
passing gravitational wave changes the length of the optical path
between the proof masses in one arm relative to the other arm. The
spacecraft is used to provide a drag-free environment for each of the
proof masses within it, by shielding the masses from solar radiation
pressures. In order to be able to detect gravitational strain levels to
the order of 10^--23^, tight pointing and positioning requirements are
placed on the spacecraft and the proof masses (e.g., acceleration
requirement on each proof mass: 3x10^-15^ m/s^2^/Hz^-1/2^). To achieve
these requirements, the LISA spacecraft are baselined to use electric
propulsion thrusters and quadrant photodiodes for position and attitude
control of each spacecraft, and capacitive sensing and actuation for
relative positioning of each proof mass to the spacecraft.

The FDAB personnel supported the LISA mission in a number of areas: (a)
Orbital design, analysis, and optimization; (b) Dynamics and control
modeling and analysis; (c) Design and analysis of Disturbance Reduction
System (DRS) control; (d) Control system design and analysis of thrust
stand facility. Details of some of the orbit design analysis can be
found in section 4.1.3.

A number of simulation and analysis models of a single LISA spacecraft
were developed and used to assess the feasibility of various
technologies, such as Micro-newton thrusters, inertial sensors,
capacitive actuation, as well as the Drag-Free Control concept. These
models, which have varying degrees of complexities, have been utilized
for trade studies, control design and analysis, etc. The most complete
of these is the 18-DOF LISA model, which includes full nonlinear
translational and rotational dynamics of the spacecraft and each of the
proof masses. Gravitational forces from the Sun, the Earth, the moon,
and other significant planets are included. DRS control has been fully
incorporated, along with instrument models of varying complexity.
Approximations for self-gravity and nonlinear stiffness effects (from
capacitive sensing and actuation) are included as well.

DRS control is a critical part of the LISA mission. It includes the
overall control system architecture for the positioning and pointing of
the spacecraft as well as the proof masses relative to the spacecraft.
In the baseline configuration, the spacecraft is responsible for
maintaining a total drag-free environment (or as close as possible to
it) for each of the proof masses. At the same time, fine pointing of
each spacecraft with respect to the other two has to be maintained
continuously. Preliminary design work for DRS control to achieve the
desired pointing and positioning accuracy has been completed. This
design is based on a decentralized approach to DRS control, wherein the
spacecraft position control is designed to center about the proof
masses, and the proof mass control maintains relative position and
attitude with respect to the spacecraft. Two options were considered for
proof mass translational control in the measurement axis, one with no
control and the other with a very low-bandwidth controller.

As part of the technology validation effort for LISA and other missions,
a thrust stand facility is being developed at Goddard for
characterization of the dynamics and noise characteristics of
micro-newton thrusters. The stand is based on a torsional pendulum
concept, where a thruster is placed at an offset from the torsion fiber.
A thrust force produces a torque about the fiber, and causes it to
twist. In an open-loop mode, the twist angle measurement is used to
compute the thruster force output. In a so-called "null" mode,
capacitive sensing and actuation is used to regulate the twist angle,
and the net actuation force/torque is used as a measure of the thruster
force output. A digital controller was designed for actuating the
capacitors in the null mode as well as regulating the power supply. A
detailed simulation and analysis model for the thrust stand was
developed to analyze the controller performance.

![](media/image10.png){width="4.4in" height="3.4097222222222223in"}

\[Technical Contacts: Peiman Maghami, Steve Hughes\]

**[3.9 Leonardo]{.underline}**

The purpose of the Leonardo mission is to define the Bi-directional
Reflectance Distribution Function (BRDF) of sunlight off of Earth's
clouds. Leonardo uses a formation of several spacecraft to measure the
sun light reflectance off of a cloud simultaneously from different
perspectives.

![](media/image11.png){width="3.099212598425197in"
height="2.399212598425197in"}Early FDAB analysis of the Leonardo mission
considered a formation of three spacecraft in a sun synchronous orbit.
The analysis showed that this configuration required too much
propellant. Subsequent analysis has been on a formation of six
spacecraft in near equatorial low Earth orbits.

A sophisticated model has been developed to analyze the Leonardo
mission. The model includes genetic search and sequential quadratic
programming search algorithms, and high fidelity orbit propagation. All
of the orbit parameters for each spacecraft are control variables and
the scientist's BRDF algorithms form the objective function. Future work
will focus on determining the propellant required for formation
initialization and maintenance over the mission life.

\[Technical Contact: Steve Hughes\]

**[4.0 Technology Development Activities]{.underline}**

[4.1 Advanced Mission Design]{.underline}

[4.1.1 Generator Software Tool]{.underline}

This year the FDAB partnered with Purdue University to develop advanced
mission design capabilities for libration orbits using dynamical systems
theory, or more specifically, manifold applications. This work
culminated in the delivery of a software tool called Generator.

The research goals are to develop and infuse advanced mission design
technologies to enable rapid and robust mission design planning and
contingency analysis. The primary goal is to reduce operations cost
while providing a highly accurate system to increase the efficiency of
the orbit design process. The benefits of this effort provide a
tremendous time savings in trajectory design and robustness in
trajectory design, especially for libration orbit design. This work has
application to contingency analysis and enables innovative classes of
mission orbits to support science goals.

The potential customers of this research include GSFC's LWS, SEC, NGST,
TRIANA, and Orginis Programs just to name a few. Collaboration with JPL
and JSC, are ongoing and will be enhanced with this tool. Our approach
was to develop algorithms, prototype/finalize code, and write
mathematical specifications for integration into a GN&C MATLAB based
system. This provides a design of trajectories using dynamical system
theory / manifolds and advanced numerical techniques. The research
followed an iterative prototyping approach, with frequent user
demonstrations, feedback, and revisions based on user input. Techniques
(solar sail accelerations, etc.) and linear and non-linear feedback
control were also addressed. TRIANA and NGST were used as the test bed.
Below are menu and output from this tool (Figure 4-1).

![](media/image12.png){width="1.7326771653543307in"
height="2.811811023622047in"}![](media/image13.png){width="1.8381889763779529in"
height="1.9137795275590552in"}![](media/image14.png){width="3.7118110236220474in"
height="1.2881889763779528in"}![](media/image15.png){width="2.4645669291338583in"
height="1.8838582677165354in"}

Figure 4-1. Menu and Output from the Generator Software Tool

\[Technical Contact: [David Folta](mailto:david.folta@gsfc.nasa.gov)\]

[4.1.2 SBIR Phase II Research: Optimal Orbit Transfer Analysis for
Advanced Space Systems]{.underline}

As part of overall FDAB research support, we have been supporting SBIR
phase-II contracts for advanced methods in Optimal Orbit Transfer
Analysis for Advanced Space Systems. Techniques for effectively
analyzing orbit transfers of advanced space systems employing low-thrust
propulsion are developed and employed. The first two methods, known as
collocation and parallel shooting, are trajectory modeling methods
useful for solving boundary-value-problems (BVP). Both modeling
techniques may be used in either a direct optimization formulation where
optimal control problems are transformed into a mathematical programming
(MP) problem or a two-point-boundary-value-problem (TPBVP) that results
from applying the calculus-of-variations (COV) to optimal control
problems. The last two techniques, known as sparse nonlinear programming
(SNLP) and genetic algorithms (GA), are mathematical programming methods
that together promise to provide a valuable strategy for optimizing a
variety of complex orbit transfer problems. Orbit transfer dynamics are
modeled using equations-of-motion based upon modified equinoctial orbit
elements that include restricted, third-body gravity effects. Software
based on these techniques is developed for use in optimizing low-thrust
orbit transfers on a variety of challenging science mission scenarios
and the relationship between the MP-BVP and the COV-TPBVP formulations
is investigated and exploited for the missions investigated.

# A delivery of a beta version of an advanced low thrust trajectory design tool with optimization capability was delivered. The tool has capabilities for low thrust ( and near impulsive thrust) for orbit in highly elliptical orbits, transfers to the moon, and libration orbits.

#  

# \[Technical Contact: [David Folta](mailto:david.folta@gsfc.nasa.gov)\]

# 

[4.1.3 Optimal Mission Design]{.underline}

Recent trends in Distributed Spacecraft (DS) mission concepts are
challenging conventional approaches to mission design and optimization.
In response, we are developing new design approaches that permit optimal
mission design for a wide variety of mission objectives. The approach
uses Sequential Quadratic Programming to optimize orbits according to a
user defined performance measure. The performance measure can be a
function of relative spacecraft geometry, science performance, inertial
orbit location, and orbit stability to name a few. Furthermore, we can
impose constraints on the orbit to satisfy a wide range of mission
constraints including maximum eclipse time, periapsis altitude, and
relative vehicle dynamics.

![](media/image16.wmf){width="3.94375in" height="3.088888888888889in"}

**Figure 4-2. Performance and Constraints History**

The approach has been applied to two different scenarios that we present
briefly here. In the first example the science objective is to study the
plasma sheet, and the engineering requirement is to maximize the amount
of time that the spacecraft spend in the Plasma sheet. Due to power
limitations and orbit correction limitations, there are several
constraints that the optimal solution must satisfy. The maximum eclipse
time must be four hours or less. Also, the periapsis must be between 1.1
Re and 1.5 Re and apoapsis must be between 17.5 Re and 18.5 Re. In the
figure below we can see that by minimizing the negative of the days
spent in the Plasma Sheet and satisfying the design constraints each
spacecraft will spend about 90 days in the primary region of interest
over the two year mission life.

The second design example is the Laser Interferrometer Space Antenna
(LISA). LISA is a three spacecraft concept designed to measure gravity
waves. One possible measure of performance for LISA is to keep two of
the legs of the formation equal over the entire mission life. A baseline
orbit was designed by Folker *et.al.* and was used as an initial guess
in the optimization of the formation. Due to science constraints the
baseline orbit was chosen to be heliocentric with the three spacecraft
rotating about a virtual hub. Some further constraints are imposed on
the design due to mechanical limitations of the spacecraft. The rate of
change of distance between any two spacecraft must be less than 15 m/s.
The angle between the spacecraft must always be between 59 and 61
degrees. Finally, for science, the range between the spacecraft must be
on the order of 5 million kilometers. In the figure below, we see an
optimal solution that satisfies all of the above constraints. In this
case we chose to make the distance between spacecraft 2 and 3 and
between spacecraft 1 and 3 as close to equal as possible over the entire
mission life. For this particular solution the distances are equal to
within a few hundredths of a percent.

![](media/image17.wmf){width="4.875694444444444in"
height="3.6569444444444446in"}

Figure 4-3. Relative Geometry Evolution

References:

Folkner,W.M., Hechler, F., Sweetser, T.H., Vincent, M.A., and Bender,
P.L. "LISA Orbit Selection and Stability," *Classical Quantum Gravity*,
Vol. 14, 1997, pp. 1405-1410.

\[Technical Contact: [Steven
Hughes](mailto:shughes@pop500.gsfc.nasa.gov)\]

# **[4.2 Autonomous Navigation Technologies]{.underline}**

# 

# **[4.2.1 Autonomous Onboard Navigation Systems]{.underline}** 

![](media/image18.wmf){width="5.984645669291338in"
height="3.2763779527559054in"}![](media/image19.wmf){width="0.16889763779527558in"
height="0.2043307086614173in"}![](media/image19.wmf){width="0.16889763779527558in"
height="0.2043307086614173in"}![](media/image19.wmf){width="0.16889763779527558in"
height="0.2043307086614173in"}![](media/image19.wmf){width="0.16889763779527558in"
height="0.2043307086614173in"}The technologies developed in this work
area enable the following: advanced mission concepts such as formation
flying, solar sailing, and low-thrust orbit transfer; autonomy for all
aspects of navigation including maneuver planning and execution,
communication signal acquisition, realtime onboard attitude
determination and control; design flexibility by providing a single
navigation software system for multiple mission scenarios to enhance
autonomy; and highly accurate autonomous onboard inertial and relative
navigation for multiple satellites. The approach optimizes use of
available sensor data onboard the vehicles. It reduces mission
life-cycle cost for single and multi-spacecraft platforms, by minimizing
ground and tracking operations, and by reducing the development and test
cost of autonomous navigation while increasing the efficiency of the
navigation process.

The technical approach focuses on enabling formation flying and
distributed spacecraft by improving automation, autonomy, mission design
flexibility, and accuracy of flight dynamics functions. Evaluation of
future mission needs and research of existing and state of the art
methods is performed. Improved or new algorithms are then developed,
prototyped and tested. Feasible developments will advance to flight
experiments and eventually transfer to mission operations. Software
development emphasizes reusability,

Figure 4-4. Navigation Scenarios

maintainability and easy reconfiguration for various mission needs.
Realistic simulated and actual fight

data are used for analysis and testing whenever possible. Partnerships
with industry, other government agencies and grants with universities
are utilized when appropriate. Below, we describe accomplishments of the
FDAB navigation team in the last year in the areas of algorithm and
software development, sensor integration and test, flight experiments,
and future missions analysis.

## 

## *Algorithm and Software Development*

The primary software product of the FDAB navigation team is the
GPS-Enhanced Onboard Navigation System (GEONS), a multi-purpose
navigation software package that maximizes software reusability and
maintainability, and can be easily reconfigured to a user's needs. GEONS
is the "container" in which research and development activities are
captured. GEONS is based on GEODE, a runner-up in NASA's 2000 Software
of the Year competition, which has been successfully transferred to
industry, academia, and other U.S. government agencies. A version of
GEODE is now successfully operating onboard the EO-1 spacecraft, and it
is being integrated into several GPS receivers by teams inside and
outside the government. GEONS builds on GEODE by integrating one-way
Doppler measurements using GN stations, and celestial navigation
capabilities based on data from ACS sensors. A future TDRSS capability
will be built on the TONS software, which successfully provides onboard
autonomous navigation onboard the Terra spacecraft since early 2000.
GEONS will also incorporate the capability for onboard maneuvering,
including decentralized cooperative maneuvering by formations of
spacecraft, and support for GPS/INS integration, including U.S.
Government patented algorithms for GPS attitude determination. The
following paragraphs describe our major software releases in the past
year.

### **GEODE 5.4.2**

The FDAB navigation team delivered GEODE Release 5.4.2, and the
associated System Description and the User's Guide and Mathematical
Specifications. This release, which is available to all licensed GEODE
users, has undergone extensive acceptance testing and is intended to be
the final release of the GEODE flight software. This version contains
enhancements supporting relative navigation capabilities: simultaneous
estimation of multiple satellite states; estimation of GPS Space Vehicle
measurement biases; processing of measurements from geostationary (GEO)
satellites associated with the GPS Wide Area Augmentation System (WAAS);
processing of singly differenced GPS and WAAS GEO measurements; and
processing of measurements from cross-link receivers. It also fixes a
bug associated with filter restart conditions.

### **GEONS 1.1**

The FDAB navigation team delivered GEONS Release 1.1 and the associated
Mathematical Specifications and System Description and User's Guide.
This release, which will eventually be made available to licensed GEONS
users, incorporates enhancements for Ground-Station-to-satellite Doppler
measurements, Doppler compensation prediction, a backup ephemeris
computation algorithm, and inclusion of externally measured
accelerations in the spacecraft acceleration modeling. GEONS Release 1.1
provides all capabilities previously available in the Ground Onboard
Navigation System (GONS) and GEODE Release 5.4.2. We also delivered a
report that defines the mathematical algorithms that can be used to
provide autonomous navigation using standard spacecraft attitude sensors
and communication components. These algorithms have been implemented in
the CelNav program and will be implemented in the next version of GEONS.

### **DATSIM** 

The FDAB navigation team delivered user instructions and mathematical
specifications for the measurement data simulation (DATSIM) program.
DATSIM is used to simulate the one-way forward pseudorange, Doppler, and
antenna signal-to-noise ratios for the Global Positioning System (GPS),
Wide Area Augmentation System (WAAS), cross-link, and ground station
tracking systems. The simulated tracking data can then be processed
using the Embedded Onboard Navigation System (EONS) and/or GEODE/GEONS
software.

## 

## *Sensor Integration & Test*

### **Low Power Transceiver (LPT)**

The LPT is a navigation and communications sensor being developed as a
joint effort between ITT and NASA that is incorporating GEODE as its
navigation processing software. The FDAB navigation team has provided
extensive support to ITT personnel developing the LPT. We processed
receiver test data and provided ITT staff with filter tuning parameters
consistent with the quality of the pseudorange measurements and the
characteristics of the receiver clock. We also provided extensive
support in the investigation of timing issues. We also provided
extensive consultative support.

### **PiVoT GPS Receiver**

PiVoT is an open-architecture GPS receiver that GSFC is developing
in-house (Figure 4-5). We performed a detailed analysis of the PiVoT
software that interfaces with the GEODE navigation flight software,
identifying several inconsistencies related to the time systems and the
contents of interface structures expected by GEODE. We performed a
detailed analysis of PiVoT system testing results, identifying several
possible causes of the processing errors that we observed. In addition,
we recommended tests that could be performed to isolate the cause of the
errors.

![](media/image20.jpeg){width="2.3in" height="1.3888888888888888in"}

## 

## 

## 

## *Flight Experiments*

![](media/image21.wmf){width="4.0in"
height="3.4659722222222222in"}*Autonomous Navigation of EO-1 Using GPS*,
which provides a detailed assessment of the accuracy of the navigation
solutions computed by the receiver on EO-1 satellite. Figure 4-6 shows
two spacecraft in formation flying configuration.

## 

## *Future Missions Analysis*

### **Celestial Navigation**

The FDAB navigation team delivered the technical report *Autonomous
Navigation of Libration-Point Orbiters Using Celestial Objects and
Doppler Measurements*, which provides the results of an analysis of the
accuracy achievable using Doppler measurements for autonomous navigation
of a satellite orbiting the L1 Sun-Earth libration point. This analysis
indicates that after 22 days of processing, the estimated position and
velocity errors reach steady-state levels of 6.5 kilometers and 2
millimeters per second root-mean-square, which is comparable to the
accuracy of the operational reference solution obtained using DSN
round-trip range and range-rate measurements.

The FDAB navigation team also delivered a report that quantifies the
navigation performance that can be achieved using standard spacecraft
attitude sensors and communication components to provide autonomous
navigation for high-Earth orbit missions. Based on the processing of
real *Polar* spacecraft measurements, this analysis demonstrates that an
autonomous navigation accuracy of about 10 kilometers root mean square
can be achieved for a 1.8-by-9-Earth-radii spacecraft using realistic
Sun and Earth sensor measurements. Using high-quality forward-link
Doppler measurements, an autonomous navigation accuracy of 1.0 kilometer
root mean square is achievable.

### **Relative Navigation**

The FDAB navigation team presented "Evaluation of Relative Navigation
Algorithms for Formation Flying Satellites", at the GSFC Flight
Mechanics Symposium held June 19-21, 2001 at GSFC. This evaluation
indicates that very accurate relative navigation positions can be
achieved for formations in medium-attitude and high-altitude eccentric
orbits using only Global Positioning System (GPS) measurements. The
addition of round-trip intersatellite range measurements was shown to
significantly improve relative navigation accuracy for formations with
sparse tracking of the GPS signals. Figure 4-7 shows the orbital
geometry with respect to GPS broadcast signal.

### ![](media/image22.wmf)

### 

### 

### 

### 

### 

### 

### **Figure 4-7. Satellite Orbital Geometry With Respect to GPS Signal**

### **Constellation-X**

The FDAB navigation team performed an analysis of the autonomous
navigation accuracy that could be achieved for the *Constellation-X*
mission. We derived these accuracy estimates based on earlier detailed
simulations of navigation accuracy for the SOHO mission as a function of
measurement type, measurement accuracy and tracking frequency. We
presented these results to the *Constellation-X* project scientists.

### **Magic** 

The FDAB navigation team performed an analysis of the autonomous
navigation accuracy that could be achieved for the *Magic* mission. This
analysis included generating truth motion files, processing the GPS
measurements collected by the PiVoT receiver, and simulating and
processing realistic GPS measurements for two spacecraft in the *Magic*
constellation. We provided a detailed summary of the absolute and
relative navigation accuracies that can be achieved using an operational
scenario that meets the constraints of the *Magic* mission. As a result
of this analysis, *Magic* has selected GEONS/PiVoT as their primary
navigation system option, which GNCC will provide as GFE to *Magic* if
they are selected in the next phase of Midex missions.

\[Technical contacts: [Russell
Carpenter](mailto:jrcarpen@pop500.gsfc.nasa.gov), [David
Folta](mailto:david.folta@gsfc.nasa.gov), [Cheryl
Gramling](mailto:cheryl.gramling@gsfc.nasa.gov)\]

# **[4.2.2 Adaptive Kalman Filter for Autonomous Navigation]{.underline}** 

The Kalman filter produces recursively optimal estimators of the dynamic
state with well-defined statistical properties. It has been extended and
modified to support high-accuracy spacecraft navigation on earth orbits.
For autonomous navigation, the identification and optimization problem
is introduced since the navigation system needs to perform actions that
change the structural parameters of the plant the controller is
interacting with. In this case, an adaptive or self-tuning filter is
needed. In other words, there is a requirement to identify the relevant
characteristics of the system in order to control it optimally. Numerous
adaptive Kalman filters have been developed. Major adaptive Kalman
filters, such as Jazwinski's and Magill's, have been usually referred to
when dealing with autonomous navigation issues. There are some drawbacks
in these techniques when applying to geocentric orbits, where gravity
modeling errors play an important role in the orbit estimation problem.
For Kalman filters and adaptive Kalman filters, the white process noise
hypothesis is essential. Gravity modeling errors, however, have been
demonstrated as auto-correlated with respect to time to the extent that
any white noise approximation will yield a nonoptimal procedure.

A research was conducted to develop an adaptive technique for autonomous
navigation systems on Earth orbits. It proposes a sophisticated
application of neuro-fuzzy techniques to perform the self-tuning
capability. It also demonstrates the feasibility and efficiency of a
self-tuning component built from this concept to augment to a Kalman
filter, which performs the state estimation. The core requirement is a
method of state estimation that handles uncertainties robustly, is
capable of identifying estimation problems, flexible enough to make
decisions and adjustments to recover from these problems, and compact
enough to run on flight software.

![](media/image1.wmf)

**Figure 4-8. Architecture of a Self-Tuning Kalman Filter**

**for Autonomous Navigation (Using GPS)**

The scope of this research has both theoretical and experimental
dimensions. In the direction of theory, performance limits of Kalman
filter and related major adaptive techniques, and new technologies
popular in the areas of system identification and automatic controls are
studied, with special emphasis on mathematical issues leading to the
optimization of spacecraft navigation autonomy. In the experimental
direction, a prototype self-tuning system is designed, developed, and
tested. Filtered data from real and simulated GPS measurements are
carefully prepared to train and check the accuracy of the system. The
experimental implementation establishes the reliability and accuracy of
the mathematical foundations of neuro-fuzzy techniques underlying the
self-tuning process. Results from the testing of the prototype show that
this self-tuning technique can achieve the accuracy of less than 5 cm in
total position.

Figure 4-8 illustrates the architecture of a self-tuning Kalman filter
for Autonomous Navigation using GPS. The tuning subsystem prototype is
simply a three inputs/three outputs neuro-fuzzy system augmented by a
preprocessor that gathers filter outputs (i.e. state error covariance)
in time series, determines if the filter re-tuning is needed, and uses
least-squares process to linearly fit them. The preprocessor also builds
a vector that represents the behavior of the covariance and that is
input to the neuro-fuzzy system. Parameters are tuned using the hybrid
option that is a mixture of least squares and back propagation
techniques. The prototype neuro-fuzzy system consists of three Adaptive
Neuro Fuzzy Inference Systems (ANFIS). Each ANFIS is built from a
three-input Sugeno Fuzzy Model with 27 rules, as shown in Figure 4-9.
There are 536 samples generated to train and test this prototype. These
samples are selected outputs from 536 runs using GEODE Version 5 for
single satellite and clean simulated GPS data as input.

> ![](media/image2.wmf){width="5.502777777777778in"
> height="3.372916666666667in"}

**Input Output**

**Figure 4-9. Diagram of a Three-Input ANFIS with 27 Rules**

This concept of a robust and self-tuning Kalman filter for autonomous
spacecraft navigation is also extended to broaden its mission scope to
include geosynchronous orbits and near-Earth high-eccentricity orbits.

\[Technical Contact: [Son Truong](mailto:son.truong@gsfc.nasa.gov)\]

# **[4.2.3 Magnetometer Based Navigation]{.underline}** 

##### *Overview*

Magnetometer based navigation (MAGNAV) provides low cost, autonomous
navigation for low Earth orbit (LEO) missions. The magnetometer has four
primary advantages. First, it is always part of the sensor complement
for LEO missions, primarily for momentum management. Second, it always
outputs data; that is, it is not subject to occultation or tracking
problems. Third, it is very reliable. Lastly, it provides information on
spacecraft attitude, rate, and orbit. The system developed in GNCC is
based on an Extended Kalman Filter algorithm, combined with a
pseudo-linear Kalman filter, producing the full set of navigation
parameters, namely attitude, orbit, and rate. Reducing the complexity of
onboard processing, eliminating costly sensors, and reducing ground
operating costs, while providing accuracy and reliability are additional
objectives of MAGNAV.

![](media/image3.wmf){width="0.8020833333333334in"
height="0.7368055555555556in"}

![](media/image4.wmf){width="0.5520833333333334in"
height="0.6131944444444445in"}

![](media/image5.wmf){width="0.8159722222222222in"
height="0.8159722222222222in"}

Figure 4-10. High Level Overview of Onboard Operation

Typically, the MAGNAV algorithm, in order to provide simultaneous
attitude, orbit, and rate estimates, also processes data from an
additional sensor, such as a gyro, sun sensor, or GPS (operating alone
the magnetometer can provide either attitude and rate or orbit
estimates). This improves the accuracy and speed of convergence, and
ensures robustness. A magnetometer-gyro configuration has been tested
with real data from four GSFC satellites. A magnetometer-sun sensor
configuration has been tested with TRACE data and is scheduled for an
inflight test onboard the WIRE spacecraft. The magnetometer-GPS
configuration (GPSMAG) underwent analytical testing in FY01, with the
goal of developing a 'black-box' spacecraft navigation system, as
depicted in Figure 4-10. Example results are given in Table 4-1 below.
It is expected that MAGNAV could be used in a backup mode; startup mode,
e.g. initialization; anomaly resolution; or as a prime navigation system
for a LEO mission with coarse requirements.

##### 

##### *In-Flight Experiment of MAGNAV on WIRE Spacecraft*

An in-flight experiment of the MAGNAV algorithm will take place onboard
the WIRE spacecraft. The flight code has been prepared and is undergoing
final ground testing. The code is expected to be uplinked and patched
into the WIRE onboard computer in early October of 2001, with two weeks
of testing to follow. MAGNAV will run as an independent task, in
parallel with the fine and coarse onboard attitude determination
systems. It is anticipated that this test will demonstrate the
capabilities of the MAGNAV algorithm to provide low cost, autonomous
estimates of orbit, attitude, and rate for low Earth orbit satellites.

## *IR&D Funded Research of GPS/Magnetometer Navigation*

Analytic testing of the GPSMAG, a version of MAGNAV that incorporates
GPS measurements, was conducted during FY01. The testing was performed
using Matlab. The spacecraft simulation was based on a UARS ephemeris
and included simulation of the GPS constellation. The algorithm was
successful in estimating the spacecraft orbit, attitude, rate, and GPS
clock errors using simulated measurements from two GPS satellites (both
phase and pseudo-range), along with magnetometer measurements. Starting
with initial errors of 500 km/axis in position, 0.5 km/sec/axis in
velocity, 103 degrees in attitude, and 5 deg/sec/axis in rate, the
average RSS errors after 12 hours were less than 0.3 deg in attitude,
0.003 deg/sec in rate, 30 meters in position, and 6.5 cm/sec in
velocity. Most of the convergence occurred within the first 50 minutes.
Additionally, the algorithm was able to follow a simulated 80-degree
rotation about a spacecraft body axis. The results of the testing were
presented in two conference papers.

\[Technical Contacts: [Julie
Thienel](mailto:jthienel@pop500.gsfc.nasa.gov), [Rick
Harman](mailto:rharman@pop500.gsfc.nasa.gov)\]

**[4.3 Formation Flying Technologies]{.underline}**

#### [4.3.1 EO-1 Formation Flying Experiment]{.underline} 

NASA's first-ever autonomous formation flying mission has been
successfully completed! With the launch of NASA's Earth Observer-1
satellite, called EO-1, NASA's Goddard Space Flight is demonstrating the
capability of satellites to react to each other and maintain a close
proximity without human intervention. This advancement allows satellites
to autonomously react to each other's orbit changes quickly and more
efficiently. It permits scientists to obtain unique measurements by
combining data from several satellites rather than flying all the
instruments on one costly satellite. It also enables the collection of
different types of scientific data unavailable from a single satellite,
such as stereo views or simultaneously collecting data of the same
ground scene at different angles.

![](media/image6.wmf){width="3.460236220472441in"
height="2.5098425196850394in"}Formation Flying is exactly that,
satellites flying in a predetermined formation, and maintained in that
formation by using onboard control. Therefore, when one satellite moves,
the others move to coordinate their measurements. EO-1 was launched this
past December as a technology mission designed to fly in formation with
another NASA satellite called Landsat-7, as shown in figure 4-11. Both
satellites carry instruments that enable scientists to study
high-resolution images and climatic trends in the Earth's environment.
The EO-1 satellite flies only 60 seconds (450 kilometers) behind
Landsat-7 and maintains the separation within 2 seconds. This separation
is necessary for EO-1 to observe the same ground location through the
same atmosphere region. It also demonstrates significantly improved
return of science data. The mission allows engineers to compare
technological advances made in ground observing instruments that are
smaller, cheaper, and more powerful. EO-1 also demonstrates technologies
for propulsion, onboard processing, and data storage.

Previously, satellites did not communicate directly with each other, did
not plan and execute orbital maneuvers onboard, nor were they equipped
to autonomously accommodate the actions of any other satellite in
support of a desired scientific experiment. Onboard EO-1 is an advanced
technological controller that is capable of autonomously planning,
executing, and calibrating satellite orbit maneuvers. On EO-1 it is used
for the computation of maneuvers to maintain the separation between the
two satellites. The idea and mathematical algorithm for this NASA first
was developed by Dave Folta, John Bristow, and Dave Quinn, Aerospace
Engineers of the Flight Dynamics Analysis. It is designed as a universal
3-Dimensional method for controlling the relative motion of multiple
satellites in any orbit. Their idea was then combined with a new flight
software that is the predecessor of a GSFC sponsored commercial software
call FreeFlyer produced by Lanham, MD based *a.i.-solutions* *inc*. This
flight software provides for the ingest of real-time navigation data
from the onboard Global Positioning System (GPS), the transfer of data
from the maneuver algorithm for maneuver commands, onboard predictions
of where the satellites will be in the future, the necessary attitude
pointing, and actual onboard commanding of the thruster firings.

Because maneuver calculations and decisions can be performed onboard the
satellite, the lengthy period of ground-based planning currently
required prior to maneuver execution will eventually be eliminated. The
system is also modular so that it can be easily extended to other
mission objectives such as simple orbit maintenance. Furthermore, the
flight controller is designed to be compatible with various onboard
navigation systems.

Formation flying technologies are primarily concerned with the
maintenance of the relative location between many satellites. Much
shorter and more precise baselines can be established between the
satellites. The satellites can then be combined as part of a "virtual
satellite" that should provide previously unobtainable science data
using mass produced, single-string, relatively cheap satellite. Multiple
scientific instruments often present competing and conflicting
requirements on a satellite design and its operation. So much science at
stake for a single satellite often requires a great deal of onboard
redundancy, which imposes its own overhead on the design process.
Separating scientific payloads onto several simpler single-string
satellites can accomplish the same complex missions without the added
design and operational overhead, while risking only one payload at a
time. The proposed approach for onboard formation control will enable a
large number of satellites to be managed with a minimum of ground
support. The result will be a group of satellites with the ability to
detect errors and cooperatively agree on the appropriate maneuver to
maintain the desired positions and orientations.

Since this technology is now fully developed and demonstrated,
synchronous science measurements occurring on multiple space vehicles
will become commonplace and the concept of Earth observing 'virtual
platforms' will become a reality. In the process, this technology
enables the development of autonomous rendezvous. Scientific payloads
could be launched from any launch vehicle, rendezvous with and join a
formation already in place, and then autonomously maintain this
condition or respond to specific requests for science data collection by
altering its own orbit. Thus, this technology addresses all of the NASA
directives to build revolutionary satellites that are better, faster,
and cheaper.

#### ![](media/image7.wmf){width="3.5083333333333333in" height="2.0805555555555557in"}

Figure 4.12. Orbit Mechanics of EO-1 Formation Flying

In Figure 4-12, EO-1 starts a formation at the red dot located behind
Landsat-7 by 450 kilometers and above by approximately 50 meters. Due to
the differences in the accelerations from atmosphere drag and spacecraft
design, the EO-1 satellite orbit decays faster than that of Landsat-7.
While above Landsat-7, EO-1 is drifting away from Landsat-7. After
several days of atmospheric drag, EO-1 will be below Landsat-7 and will
drift towards it. When EO-1 is outside the required separation distance
or if the Landsat-7 satellite has maneuvered away, EO-1 will
autonomously compute and perform a maneuver to reposition it to an
initial condition to repeat the relative motion and meet science data
collection requirements.

\[Technical Contact: [David Folta](mailto:david.folta@gsfc.nasa.gov)\]

# 

#### 

#### [4.3.2 SBIR Phase II Research: Autonomous Unified Orbit and Attitude Control for Formation Flying using GPS and LQG/LTR Controller]{.underline}

The objective of this work was to provide formation flying control
designs that can be used to support many of the GSFC distributed
spacecraft architectures of multiple imaging, interferometry, and robust
control. The design evolution compliments a complete design of an actual
Autonomous On-Board Orbit and Attitude Control System for Formation
Flying. The definition of Formation flying used here is where spacecraft
are reactive to each other and involve a cross-link communication path
between the spacecraft. The spacecraft will be with a line of sight
distribution and will revolutionize the way scientific measurements are
captured. Figure 4-13 shows spacecraft in formation flying.

The Satellites in these future formations need to be maintained or
controlled at designated positions in the formation group. The orbit
control system consists of an onboard closed-loop feedback controller
and an orbit real-time on-line estimator (i.e. Kalman Filter). The input
of the controller is the direct orbit measurement data, possibly GPS or
other ground generated states. The attitude needs to also be estimated
and controlled, concurrently with the orbit control. The attitude
determination system will use an onboard navigation system input with
other spacecraft hardware such as gyro inertial units. The attitude
controller will also use a non-linear attitude control law. The methods
used for control cover LQG/LTR and Lyapunov.

![](media/image8.wmf){width="3.272222222222222in"
height="2.8333333333333335in"}

Space Products and Applications (SPA) Inc. has completed a unified orbit
and attitude control system to meet distributed spacecraft requirements.
This work includes pulse modulation effects of a propulsion system used
in a finite maneuver model, GPS models, attitude and orbit estimation,
and the relative controllers for each. They incorporated their MATLAB
based design into a GSFC high fidelity simulator. Several papers have
been published on this topic, with the most recent paper entitled
"Design and Implementation of Synchronized Autonomous Orbit and Attitude
Control for Multiple Spacecraft Formation Flying Using GPS".

The completion of this SBIR-II allows GSFC to perform a unified control
approach to formation flying and these control algorithms are being
proposed for an EO-1 technical onboard formation flying demonstration.

\[Technical Contact: [David Folta](mailto:david.folta@gsfc.nasa.gov) \]

# 

# **[4.3.3 Decentralized Estimation and Control of Distributed Spacecraft]{.underline}** 

Decentralized control is an appealing approach to maintaining satellite
formations for several reasons. It is non-hierarchical, so that
coordination by a central supervisor is not required, but it retains the
optimality of centralized control. Each satellite need only process its
own local measurement data, in a form of parallel processing. Detected
failures degrade system performance gracefully. For a given level of
system reliability, a decentralized architecture may be cheaper to
build, since the individual spacecraft can be built with much lower
individual levels of reliability than the supervisor satellite in a
centralized architecture.

When we began this research program two years ago, the basic principles
and technology concepts of decentralized control for satellite
formations had only just been observed and formulated. To bring this
promising technology to a level of readiness feasible for use in
upcoming formation flying missions, our research has focused on
investigating implementation issues and testing in a relevant
environment. In the former area, we studied fault detection, isolation,
and recovery, command and data handling system and communications
channel noise and latency, and modeling issues related to the handling
of nonlinearities inherent in the satellite cluster problem. In the
latter area, we augmented existing resources at Goddard, the University
of California at Los Angeles (UCLA), the Naval Postgraduate School
(NPS), and the Massachusetts Institute of Technology (MIT) to integrate
algorithms into closed-loop avionics testbeds, with actual sensor
hardware in the loop. As a result of our research activities, the
concepts and critical functions relevant to utilizing decentralized
architectures for precise formation flying missions have been proven
analytically and experimentally. We have also validated many of the
necessary algorithms and components, and in some cases, subsystems, in
laboratory and flight environments. We have identified collaborations
over the next few years that will take the technology through system
prototype demonstrations in a space environment. Below, we highlight our
key accomplishments, summarize the additional research and development
necessary to bring the remainder of the key components and algorithms to
the subsystem validation level, and describe the outlook for flight
qualification and utilization of decentralized control architectures for
future precise formation flying missions. Figure 4-14 shows distributed
spacecraft configuration.

## *Accomplishments*

Our research program was divided into three parallel efforts. The first
effort aimed to apply the basic principles and technology concepts of
decentralized control to satellite formation flying. The second effort
concentrated on developing and validating technology concepts for fault
detection, isolation, and recovery to decentralized architectures. The
final component of our research was to validate these theoretical
results in relevant, hardware-in-the-loop and flight environments.

We have presented some of our research findings at several conferences.
Our first annual report included papers presented at the 1999 AIAA GNC
conference in Portland, OR, the 1999 AAS Astrodynamics Specialists
Meeting in Girdwood, AK, the 2000 ION National Technical Meeting in
Anaheim, CA, and the 2000 IEEE Aerospace Conference in Big Sky, MT, and
the 2000 CNES International Symposium on Spaceflight Dynamics in
Biarritz, France. We have subsequently presented more of our work at the
2001 Goddard Flight Mechanics Symposium, the 2001 SIAM Conference on
Control, 2001 American Control Conference, 2001 IEEE Conference on
Decision and Control, and the 2001 AAS Astrodynamics Specialist
Conference. A version of our final report has been accepted for
presentation at the 2002 IEEE Aerospace Conference.

Our work has also been accepted for publication in several refereed
technical journals. The Girdwood paper was published in the AIAA Journal
of Guidance, Control, and Dynamics. A paper on decentralized fault
detection has been published in the ASME Journal of Dynamic Systems,
Measurement, and Control. A paper on the extended decentralized
controller has been accepted for publication in the International
Journal of Robust and Nonlinear Control. The following paragraphs
summarize our results.

### **Application of Decentralized Control Algorithms to Satellite Formation Flying**

###### 

###### *Natural Motion of Satellite Clusters*

Creating a single "virtual" satellite out of several "smallsats"
requires placing a cluster of these smallstats in orbit. The relative
positions of the satellites within the cluster must be maintained to
some degree of precision, depending upon the particular mission
requirements. A naïve approach to configuring a satellite array might be
to imagine the satellites flying side-by-side through space, circling
the earth in parallel orbits. Some reflection on this idea, however,
reveals that this is not a natural satellite motion, and therefore
cannot be maintained without continuous control force to counteract some
portion of the natural forces of gravity. The only practical
configuration in which the relative positions of the satellites never
changes has them following each other in a single line along a single
orbit. For many missions, this is not an acceptable configuration, and
thus non-coplanar orbits must be considered. Satellites on non-coplanar
orbits will move with respect to each other, and the necessity is to
understand that motion, and use it to maintain a cluster formation.

A promising approach to this problem is to consider the motion of
satellites about an ideal, spherical planet, in orbits with small
eccentricity and slight differences in inclination. The only firm
constraint needed to keep the satellites together is that the periods of
the orbits be identical (the semimajor axes of all orbits be the same).
The paper analyses this behavior in more detail, and examines the use of
linearized equations to examine the satellite motion. The limitations of
the equations due to linearization are discussed. As the paper deals
with idealized circumstances, perturbations are not discussed. This
behavior of neighboring satellites can be exploited to create formations
of great size and complexity, with no danger of satellite collisions. So
long as the semimajor axes and eccentricities of all orbits are
identical, the satellite cluster will appear to revolve about its common
center, and the satellites will remain very nearly in the same relative
positions within the cluster.

Extending these results to orbits with large eccentricities means that
the cluster will change in size and distribution during the orbit. It is
therefore necessary to allow for this in the control law. If this is not
taken into account, the controller may expend large amounts of fuel in
an unnecessary effort to eliminate these cyclic variations. It should be
noted that even with these variations, there is still no danger of
satellite collisions due to natural motion.

###### 

###### *Development of an extended decentralized controller* 

The original decentralized controller that forms the basis for our work
required that a linear model of the system dynamics, actuators, and
measurements be available. This requirement is particularly restrictive
for realistic formation flying applications, since realistic
measurements such as GPS and crosslink ranges can almost never be
assumed to be linear functions of the state variables. We have developed
extensions to the linear decentralized controller that are similar to
the commonly used extended Kalman filter, that allows the system to be
partitioned in such a way as to exclude the nonlinearities from the
essential algebraic relationships that allow the estimation and control
to be optimally decentralized. The extended form of the decentralized
controller can be used with the linear quadratic regulator (LQR)
formulated as a model-based

tracking law, in which minimal data transmission is assured, or as a
decentralized estimator for more complex controllers. We have
investigated both the LQR-type controllers and a controller of the type
that was used for EO-1's formation flying experiment with Landsat-7.

###### ![](media/image9.wmf){width="5.277777777777778in" height="2.8472222222222223in"}

###### 

###### Figure 4-15. Extended Decentralized Controller

###### 

###### *Decentralized Control with Communication Constraints*

Control algorithms were derived for decentralized control of dynamic
teams with minimal amount of communication between the satellites and
irregular control schedule. Two relaxed data exchange patterns were
studied. It was determined that optimal control laws can be obtained in
both cases if linear quadratic Gaussian payoff is used in conjunction
with discrete dynamics. It was concluded that the minimal communication
necessary for affine optimal controls involves exchanging of the control
values between the satellites.

*Control and State-Estimate Sharing:* In this paradigm, all nodes
(satellites) exchange both their local state-estimates and the control
values every time a control has been activated at any node. The global
state estimates are determined by algebraically combining the local
state estimate of each satellite. This creates a uniformly shared global
state estimate. The control schedule, however flexible, has to be known
in advance to all satellites to permit the dynamic programming solution.
The control law for each satellite is an offline function of the last
shared global state-estimate, and the local measurements collected since
the last communication. Dynamic programming is used to derive the cost
to go at each step where controls are engaged. This cost to go at each
control actuation in the backward dynamic programming algorithm is
manipulated into the form of a static LQG team problem, and the optimal
static team solution is used to derive the control gains.

*Control Sharing:* In this case, only the control values are exchanged
noiselessly between the satellites (i.e. after a node executes a control
it broadcasts it to all other nodes). Again, knowing the control
schedule in advance allows for the construction of the control gains
using the dynamic programming backward propagation. The static team
solution is used at every control action of the dynamic programming. The
control laws are also affine, but now they are functions of the initial
condition (*a priori* state value at the initial time), the local
measurement sequence collected since the initial time, and all previous
control efforts, executed by all nodes. This appears to be the minimal
information exchange that retains affine control laws.

*Simulations:* The controls were derived for a scalar 10-step example
with two nodes, using both data exchange modes. The results were then
compared to a centralized solution, which was obtained using a two-state
controller instead of two independent nodes. The comparison shows that
the performance in both non-classical cases is only moderately degraded
from the centralized solution. There is also a moderate cost increase in
the control sharing method over the control and state-estimate sharing
solution, which is consistent with the stochastic theory of having
better results with more information. The control and state-estimate
sharing algorithm will next be applied to a satellite cluster, which is
to be controlled with minimal effort so that its geometry is maintained
relative to a reference orbit. The reference orbit with a given
eccentricity is the center of the coordinate frame to be used by every
member in the formation and moves as if it were a physical satellite. A
schedule of station keeping burns and data transmission are specified *a
priori* to permit a solution by dynamic programming.

###### 

###### *Selection of Mechanizations* 

In our initial work leading up to the Explorers award, we used a very
simplified mechanization of the satellite dynamics, actuators, and
measurements in order to utilize the original linear decentralized
control algorithm. With the development of the extended version of the
algorithm, we have been able to begin to add more realism to the
mechanization. A significant accomplishment in this area is the
development of mechanizations that allow for realistic non-linear
measurements, but retain a linear time-invariant model for the regulated
variables. This allows us to avoid the need for an online backward sweep
of the controller Riccati matrix to determine the optimal controller
gain, which is a very computationally complex proposition. Disturbances
such as higher-order gravity, drag, etc., can also be accommodated. We
have developed such mechanizations for near-circular orbits, and for
halo or Lissajous trajectories about the colinear equilibrium points of
the three-body problem, e.g. *L*~1~, *L*~2~, and *L*~3~. We have also
developed a mechanization for highly elliptical orbits that utilizes
true anomaly instead of time as the its independent variable. This leads
to a time-varying, but periodic controller gain that is only a function
of eccentricity, and may be precomputed and stored online.

###### 

###### *Reliability vs. Cost Study* 

We performed a back-of-the-envelope type of cost/benefits study,
focusing on reliability vs. cost. For this study, we developed several
tools, including an *Excel* spreadsheet, in which one can manipulate
various assumptions including total number of spacecraft, minimal number
of spacecraft required to perform the mission, number of redundant
"strings" per spacecraft, probability of unrecoverable faults on
spacecraft of varying complexities, cost to build or replace spacecraft
of varying complexities, etc. We examined various cases that covered
various levels of required reliability, numbers of spacecraft (up to
30), etc. In many cases we found that for a given total mission
reliability, decentralized architectures are cheaper, so long as
string-level reliability is not free.

###### *Development of generic Matlab code*

We have developed an event-driven implementation of the decentralized
controller that we call *Pluribus* that is suitable for real-time
embedded systems, using the *Matlab* commercial-off-the-shelf
development environment. Using *Matlab* tools, *Pluribus* can be
auto-coded directly into *C* or *C++* for implementation on flight
microprocessors. *Pluribus* is of generic design, so that users upload
model and boundary condition information specific to their own
scenarios. *Pluribus* is being used by the AFRL/NASA University Nanosat
program as part of a program of on-orbit demonstration of formation
flying, and the AFRL Techsat-21 Program has expressed in interest in
utilizing it as part of its formation flying demonstration mission.

###### ![](media/image10.wmf){width="5.194444444444445in" height="3.5277777777777777in"}

###### 

###### 

###### 

###### Figure 4-16. Decentralized Controller Data Flow Diagram

###### 

###### 

###### 

###### 

###### *Satellite cluster simulations* 

We have developed several computer programs to model the dynamics of
satellite clusters in orbit around the earth. These simulations are
suitable for modeling clusters of arbitrary numbers of satellites. The
simulations are designed to be extensible, to meet future simulation
needs. In all of the simulations, the cluster is modeled by integrating
the motion of each satellite separately. The dynamic model of satellite
motion includes the effects of solar pressure, atmospheric drag, and the
non-spherical terms in the gravitational field of the Earth. In some of
the simulations, Sun and Moon point mass perturbations can also be
included.

![](media/image11.png){width="3.9in" height="2.9277777777777776in"}

***Mechanically motivated control and orbit analysis***

An alternate approach to the control of satellite clusters is based on
the geometry of the orbit. It can be shown that any non-collision
elliptic orbit can be uniquely described by the Runge-Lenz and the
angular momentum vectors of the orbit (*non-collision* here means that
the angular momentum is non-zero). As these vectors are natural
constants of the satellite motion, and appear naturally in the solution
of the differential equations of Kepler motion, they occupy a space on
which the usual algebraic structures of Euclidean space can be applied.
It then becomes straightforward to define a \"distance\" between a
current and desired orbit, and to construct Lyapunov functions based on
this distance. From these functions, controllers are derived which can
be rigorously proven to be stable and convergent. This leads to easily
implemented controllers that can be used for arbitrary orbit transfers.
Such controllers have been investigated for orbit transfers of a single
satellite. The implementation of these controllers to cluster control
and use with an oblate Earth model will require further investigation.

### **Fault Detection, Isolation, and Recovery**

The focus in this area has been on defining the potential scenarios that
can take place when considering a cluster of satellites flying in
formation. Identification of possible absolute and relative sensor
systems that will be used in such networks is an essential part of our
current task. Moreover, the set of all *a priori* sensor, actuator, and
plant faults to be assumed for these scenarios is being constructed. A
bank of robust single fault detection filters is required to monitor,
detect and identify a set of possible *a priori* faults. The importance
of the extension to time-varying systems is that they are applicable
through dynamical system maneuvers and if the motion is sufficiently
quick, these time-varying detection filters should be more accurate than
using gain scheduling on a set of time-invariant detection filters.
Asymptotic characterization of these detection filters near this limit
shows that the dimension of these single fault filters can be greatly
reduced.

A particular important property of these single-fault detection filters
is that their gains are computed from Riccati equations. This property
is a sufficiency condition for the computation of a decentralized
detection filter algorithm. The notion is to build local detection
filters related to their sensors and *a priori* fault direction set. A
reduced state space for the local detection filters is constructed by
choosing a minimal realization associated with the local measurements,
the global dynamics, and the local fault direction set. The residual for
a global fault detector associated with mutual measurements can then be
constructed from the local detection filters. This residual is used to
detect faults in the mutual (or relative) measurements.

Currently, detection filters are being developed for individual
satellites. A set of possible *a priori* faults associated with
actuation for attitude control and stationkeeping propulsion, and a set
of measurements associated with attitude determination and navigation is
constructed. These local detection filters will form the basis for
constructing the global residual for detection and identification of
faults in the relative measurements between spacecraft.

An approach to reconstructing the sensor and actuator faults is
developed based on the structure of the fault detection filter. This
approach allows for control system reconfiguration in a very
straightforward manner by correcting the corrupted instrument or
explicitly adjusting the remaining actuators to compensate for the
faulty actuator. Fault reconstruction is very important because it
greatly increases the flexibility of the system's reaction to the sensor
and actuator faults.

### **Integrated Hardware-in-the-Loop Testing**

###### 

###### *Testbeds* 

An associated benefit of our research is that, in order to demonstrate
our algorithms in a flight-like environment, we significantly enhanced
the capabilities of hardware-in-the-loop formation flying testbeds at
Goddard and UCLA (figure 4-18). We used the GSFC Formation Flying Test
Bed (FFTB) to demonstrate a fully decentralized, operational version of
the extended decentralized controller, implemented in the *Pluribus*
generic software program. As a result of improvements made to the FFTB
through our work, it has become a major institutional resource at
Goddard, and has already begun to be utilized by a pre-proposal phase
mission that requires precise formation flying capabilities, the Stellar
Interferometer (SI) mission. We used the UCLA formation flying testbed
to validate the UCLA Formation Flying Instrumentation System (FFIS), a
cm-level GPS relative navigation system that forms the basis of our
aircraft flight tests, which we discuss further below. The UCLA testbed
has since been used to validate relative navigation and control sensors
and algorithms for a formation flying flight test using F-18s sponsored
by NASA Dryden.

###### 

###### 

###### ![](media/image12.wmf){width="2.8055555555555554in" height="2.986111111111111in"}

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### 

###### Figure 4-19. Integrated Hardware-in-the-Loop

###### 

###### *Aircraft flight test* 

A series of flight test campaigns in conjunction with the Naval
Postgraduate School were conducted. Existing unpiloted air vehicles
(UAVs) at UCLA and NPS (known as the "Mule" and the "Frog") were flown
with UCLA's FFIS, which consists of a dual-frequency GPS receiver,
inertial measurement unit, and a flight computer, onboard each aircraft.
These flights were conducted at the Camp Roberts flight test range,
approximately 40 miles north of Fresno, California.

To begin the campaign, the FFIS was extended to include the ability to
read the control surface positions, and to generate pulse-width
modulated control signal to the UAV actuators. Control laws were derived
using the available aerodynamic models of the Mule and the Frog. These
models were somewhat crude, but conservative controllers with healthy
safety margins were defined for early use. The test plan called for
autonomous flight of the Mule first, then the Frog. Only after it was
determined that the FFIS was capable of individual autonomous flight
were dual flights to be undertaken.

![](media/image13.png){width="4.19375in" height="2.18125in"}

###### 

###### 

###### *MIT "SPHERES."* 

We began collaborating with MIT in the use of small, student-built
vehicles called "SPHERES" that are designed to fly inside the Space
Shuttle middeck and the International Space Station as research
platforms for satellite formation flying. The SPHERES had their first
two flight tests on NASA's KC-135 zero-gravity research aircraft at the
Johnson Space Center in Spring of 2000, and have been manifested to fly
on the International Space Station flight 9-A, currently manifested for
late 2002.

![](media/image14.wmf){width="5.875196850393701in"
height="2.480708661417323in"}Figure 4-21. MIT "SPHERES"

## 

## 

## *Next Steps*

Although we have accomplished nearly all of our original research
objectives, we describe here at a high level what remains to be
accomplished to complete subsystem and system level development and
validation studies that must precede space flight demonstration. We
currently anticipate that continued funding from the Space Base Core
Technology Program will enable us carry this work forward to the point
of departure for a space flight validation opportunity, such as the New
Millenium Program offers.

The most important next steps in our work will be to integrate the
decentralized control algorithms we have developed with a high-fidelity
orbit determination filter, such as GSFC's GEONS, and refining the level
of detail in our controller mechanizations. The latter effort will
include continued investigations of alternative control approaches than
the linear quadratic regulator, which formed the core of the present
work.

Next, we plan to incorporate the work we have begun in the fault
detection area, by integrating a bank of single-fault filters into the
decentralized control algorithm. We are also investigating the
development of numerical algorithms for generating the optimal filter
gains for robust multi-fault detection filter. Our goal is to put the
required minimization problem into a form that is convex so that linear
matrix inequality techniques may be used. This research is aimed at
developing a practical, single, multi-fault filter that could replace
the bank of single-fault filters.

This fall, we plan to accomplish a two-week campaign of
three-dimensional (6 DOF) flight tests in the microgravity environment
aboard NASA's KC‑135. The research goal is to quantify the performance
of formation-flying algorithms by testing their applicability for
different maneuvers. These flights will also investigate improvements to
the global metrology subsystem. Pending continued funding from the Space
Base Core Technology Program, we expect to fly with the SPHERES on ISS
Flight 9-A sometime in the next year or so.

## 

## *Outlook*

We will continue to work with the University Nanosat and Techsat-21
programs to specifically tailor our algorithms to their missions. These
technology demonstration missions will begin to pave the way for precise
formation flying to be adopted by a multitude of future missions. We are
hopeful that NASA is able to create an opportunity for a full-fledged
precise formation flying demonstration mission, perhaps through the New
Millenium Program. We would eagerly anticipate the chance to compete for
a slot on such a mission.

\[Technical contacts: [J. Russell
Carpenter](mailto:russell.carpenter@gsfc.nasa.gov)\]

**[4.4 Attitude Determination and Modeling Techniques]{.underline}**

**[4.4.1 Attitude and Orbit Model Support]{.underline}**

This year, there was an update of the SKYMAP Web site (see
<http://cheli.gsfc.nasa.gov/dist/attitude/SKYMAP_021201_page.html>),
which makes available to the public and to interested professionals the
SKYMAP Master Star Catalogs and the various mission-specific SKYMAP
ground and on-board star catalogs. The latest version of the Master
Catalog (Version 3) is available for download from this page, but has
not yet been released through the Astronomical Data Centers (ADC). The
preceding version, which has been widely released, is presently the
sixth most frequently downloaded catalog from the Vizier database
maintained by the Centre de Donneès Stellaires in Strasbourg, France
(see <http://vizier.u-strasbg.fr/cats/Usage.htx>). The mission-specific
catalogs are now available from the SKYMAP Web site accompanied for the
first time by the delivery memoranda provided with the catalogs to each
mission that had commissioned one. The delivery memoranda describe
catalog format and contents, and contain recommendations for the use of
the catalogs.

This task completed necessary paperwork and provided necessary
documentation to begin the process of commercially releasing the SKYMAP
System utility MCDUMP. This program allows a user working with the
SKYMAP Master Catalog to create subsets of the 299,160-entry catalog
based upon one or more user-specified criteria (e.g., brightness in the
Johnson *V* passband, position on the Celestial Sphere, or known or
suspected variability).

Work continued to produce the SKY2000 Version 4 Master Catalog, which
will be improved by the global and comprehensive replacement of all
variable star identifiers and data fields from the following major
variable star data sources: The New Catalogue of Suspected Variable
Stars (NSV), the Supplement to the New Catalogue of Suspected Variable
Stars (NSVS), the General Catalogue of Variable Stars (GCVS), and the
76^th^ Namelist of Variable Stars (NL76), all available from the Web
site of the Sternberg Astronomical Institute (see
<http://www.sai.msu.su/groups/cluster/gcvs/gcvs/new.htm>). This update
will provide variable star identifiers and data for many Master Catalog
stars not previously identified in the MC as variable. This in turn will
assist missions using SKYMAP star catalogs in guide star selection,
allowing more reliable avoidance of variable stars where desired. In
addition, since much of the more recent work done in the area of
variable star astronomy has involved research into the variability of
stars at longer wavelengths (e.g., red or infrared), updated variability
data in the SKYMAP MC will often be for longer-wavelength passbands.
These passbands correspond more closely to the passbands of the CCD
chips used in many star trackers than the visual or photographic
passbands in which variable stars were often measured in the past.

\[Technical contact: [David
Tracewell](mailto:david.tracewell@gsfc.nasa.gov)\]

**[4.4.2 Advanced Attitude Determination and Sensor
Calibration]{.underline}**

A new gyro calibration utility was developed for a gyro quadruplet
(Figure 4-22) that will be flown on the EOS-AQUA mission. This utility
consists of a Kalman Filter that estimates the gyro scale factors,
misalignments, biases for all four gyros as well as offers a refined
rate and attitude estimate. The refined rates and attitudes are possible
by the use of the gyro data and star tracker data itself as a
measurement. Figure 4-23 demonstrates the successful determination of
gyro parameters for Aqua.

![](media/image15.wmf){width="0.3251968503937008in"
height="0.25275590551181104in"}

Figure 4-22. The Gyro Configuration of the EOS-AQUA Satellite.

> ![](media/image16.png){width="4.6194444444444445in"
> height="2.68125in"}

Figure 4-23. AQUA Gyro Scale Factor Estimates (bold) versus Truth (thin)

An advanced Real Time Attitude Determination System (RTADS) was
developed for the Microwave Anisotropy Probe (MAP) mission. Besides
providing the attitude state at any given time, the system was designed
to provide a backup for MAP in case of a gyro failure when the star
trackers were inoperable due to radiation belts. In that case, the
system's attitude state would be used to reinitialize the onboard
attitude estimate. The rate estimation tool consisted of a Kalman Filter
that estimated rate and attitude using available gyro data, the sun
sensor data, and star tracker data. This filter was run twice every
cycle, once for each gyro configuration. The missing rate information
for each gyro was supplemented using the sun sensor data. In case of a
gyro failure in the radiation belts, the last computed attitude with a
star tracker measurement would be propagated using the two rate axes
from the surviving gyro along with the rate derived from the sun sensor
measurements for the missing axis. The onboard attitude would be
overwritten with the ground computed attitude once the spacecraft
attitude profile had stabilized, and the spacecraft would shift to its
backup rate algorithm and continue operations. Figure 4-24 shows four of
the RTADS displays with the rate estimate and attitude estimate displays
on the left, the star tracker display on the upper right, and the star
status display in the lower right. The left-most sensor status plot
shows the relative star tracker alignment errors. The error can be seen
dropping to zero. An alignment correction was uplinked at that time.

![](media/image17.png){width="5.875in" height="3.5520833333333335in"}

Figure 4-24. The MAP RTADS Displays

The performance analysis of two Lockheed Autonomous Star Trackers (ASTs)
was performed using the EO-1 and IMAGE missions. The IMAGE mission is a
spin stabilized spacecraft rotating at approximately 0.5 RPM. The noise
characteristics were derived by modeling the spacecraft as a rigid body
and comparing the attitude profile derived by dynamics to that computed
from the AST. The error between the two (neglecting systematic effects)
was taken as the sensor noise. The EO-1 spacecraft is a three-axis
stabilized spacecraft. The noise characteristic was easier to derive by
computing an attitude using a batch least squares estimator, which used
the AST data along with gyro data.

\[Technical contact: [Rick Harman](mailto:rick.harman@gsfc.nasa.gov)\]

**[4.4.3 Multi-Mission Attitude Determination System]{.underline}**

NASA/GSFC recently filed for a provisional patent for the Multi-Mission
Attitude Determination System (ADS). This software system analyzes and
processes spacecraft attitude sensor and actuator data. This processed
data is used to compute spacecraft attitudes and sensor calibrations.

The system architecture, implemented with MATLAB^^, offers a flexible
environment for adapting and augmenting the configuration to meet
specific mission requirements. MATLAB^^ provides an efficient,
straightforward programming language, and compatibility with multiple
computer platforms. The ADS, as implemented with MATLAB^^, is
henceforth referred to as ADS-MATLAB. However, the system design and
architecture is independent of MATLAB^^, and may be implemented with
any appropriate interpretive programming language.

The ADS is comprised of an extensive set of specialized MATLAB^^
function files, integrated through a series of function calls from one
component of the system to another. The operator interacts with the
system through a set of graphical user interfaces (GUI's), organized in
the natural sequence of operations for a typical analysis task.

The overall system architecture is depicted in Figure 4-25 and an
example main GUI is depicted in figure 4-26.

![](media/image18.png){width="5.999305555555556in"
height="3.5194444444444444in"}

**Figure 4-25. ADS-MATLAB Overview**

The first button on the main GUI is the Telemetry Processor (TP). This
module is a mission unique set of routines, designed to interpret raw
telemetry data from the spacecraft. Typically, the attitude sensor data
is ingested from large ASCII files containing a significant amount of
unnecessary telemetry. The TP extracts the required data, and stores it
in arrays for subsequent operations.

The Data Adjuster converts the raw attitude sensor and actuator data
from the TP into observation vectors in the satellite body axes and
generates corresponding reference vectors in an inertial coordinate
frame (such as J2000). Supported sensors and actuators include: Star
Trackers with star position angle/magnitude output; Star Trackers with
quaternion output; Digital Sun Sensors (High and Low Fidelity); Coarse
(Analog) Sun Sensors; Three Axis Magnetometers; Inertial Reference
Units/Gyro (rate or accumulated angle output); Earth Sensors; Gimbaled
sun sensor; Reaction Wheels; Magnetic Torquer Assembly.

![](media/image19.png){width="5.217361111111111in"
height="3.542361111111111in"}

**Figure 4-30. ADS MATLAB Main GUI for the MAP Mission**

The Star Identification (STARID) module, which is not shown on the MAP
ADS-MATLAB, matches observed stars with reference catalog. STARID
supports a star catalog loaded as a MATLAB .mat file, or as a SKYMAP MMS
Run Catalog. The module provides Direct Match Star Identification,
Pattern Match Star Identification, and User Star Identification with a
GUI interface.

The Quaternion and Rate Estimator, Single Frame Estimator, Batch Least
Squares Estimator, and the Extended Kalman Filter provide the suite of
tools for analyzing the sensor and actuator data. Specifically for
spinning spacecraft, an azimuth estimator and a despun platform
estimator and a predicted-observed attitude data utility are included.
Based on operator selection of the analysis tool and sensor/actuator
data, the system estimates the spacecraft attitude, spacecraft rate,
gyro bias, and/or magnetometer bias.

The Utilities are a collection of tools designed to assist the analyst.
The Attitude Validation utility compares ground and On Board Computer
(OBC) computed attitude history files, outputting results to both
reports and plots. Other utilities provide reference vectors, ephemeris
file preview and selection, ephemeris comparison, and quaternion
comparison. The system supports all commonly used ephemeris file
formats. The GUI is readily adaptable to include additional, mission
specific utilities.

The Sensor Calibration module contains a set of utilities to calibrate
the sensors for various effects such as field of view errors, biases,
scale factors, and misalignments. The calibration utilities include:
Magnetometer Calibration (magnetometer alignment, scale factors, and
bias, as well as the coupling matrix between the torquer bars and
magnetometer); Attitude Dependant Alignment (estimates sensor
alignments); Attitude Independent Alignment (computes alignment
corrections); IRUCAL (estimates gyro scale factors, alignments and
biases); BICAL (estimates gyro scale factors, alignments and biases with
the ability to add batches later on to refine your estimation); FSSFOV
(Estimates field of view calibration coefficients for a fine sun
sensor); SSPPCAL (Estimates gimbal misalignments on which the SSPP is
mounted).

The Real Time Attitude Estimation (RTADS) estimates the spacecraft
attitude as well as other desired parameters in real time. This system
option executes all the ADS-MATLAB functions (DA, STARID, SF, EKF, etc.)
as well as user defined functions and displays in real time. The user
selects which processes to execute via a GUI and develops scripts (if
desired) that can be plugged into the RTADS processing by defining the
name of the script in the ADS-MATLAB namelist. The major processing
function that is required is the front-end. Again, only the name of the
front-end script needs to be defined in the ADS-MATLAB namelist.

The Load and Save module allows the operator to save and reload data,
allowing resumption of an analysis session without data reprocessing.
The module also provides the operator with a variety of options to save
output data arrays and reports, and system configuration settings.

The ADS-MATLAB provides a complete set of portable tools for attitude
determination for most spacecraft with a highly adaptable architecture
to conform to specific mission configuration requirements. The
estimation algorithms represent the current state-of-the-art technology,
and are easily modified or extended as methodologies improve. In
general, the routines are coded to optimize performance with the
MATLAB^^ environment. Also, many of the individual ADS-MATLAB routines
(m-files) form a function library that independently serve as analysis
tools, or provide a resource base for building similar systems.

\[Technical contact: [Rick Harman](mailto:rick.harman@gsfc.nasa.gov)\]

**[4.5 Flight Dynamics Automation Studies]{.underline}**

The University of Maryland Department of Aerospace Engineering has
continued to act as a test bed for researching ground system automation
techniques for the SAMPEX mission. Work completed to date includes the
pre-processing and uploading of tracking data. Work continues on the
following: parallel testing of old and new systems; improvement of user
interface; enhancement of tracking data conversion methodology; orbit
determination based on tracking data; post-processing of orbit
determination results into various products and sending the products to
their intended recipients. Planned work includes the following: complete
the development of a graphical user interface accessible through a web
browser which will allow a user to observe the status of the process,
obtain the latest results, and to modify the products produced including
where certain products should be sent to, as well as when this should
occur. This will be done so as to automate the current manual process as
well as being flexible enough to address future products

\[Technical contact: [Joe Toth](mailto:rick.harman@gsfc.nasa.gov)\]

**[5.0 Branch Infrastructure]{.underline}**

**[5.1 Flight Dynamics Tool Program]{.underline}**

The Flight Dynamics Analysis Branch was engaged in activities to
maintain and enhance the capabilities of flight dynamics software used
for mission feasibility, analysis and operations support. Work was
performed in the technical areas of attitude estimation and sensor
calibration, attitude control subsystem (ACS) analysis and design,
navigation and orbit determination, and trajectory design and mission
planning. These activities included identifying and correcting software
errors; development, implementation,

testing and validation of new software algorithms; documentation of
existing software systems; and, evaluation of commercial off-the-shelf
(COTS) products and in-house capabilities. The large systems maintained
by this effort include the Attitude Determination Error Analysis System
(ADEAS), Orbit Determination Error Analysis System (ODEAS), Goddard
Trajectory Determination System (GTDS), General Maneuver (GMAN) Program,
Swingby, Multi-mission Spin Axis Stabilized Spacecraft System (MSASS)
and Multi-mission Three-Axis Stabilized Spacecraft System (MTASS).

\[Technical contact: [John Lynch](mailto:rick.harman@gsfc.nasa.gov)\]

**[5.2 Flight Dynamics Lab]{.underline}**

The Flight Dynamics Lab continued to provide support for the
development, test, integration and operation of software systems as well
as analysis for the performance of flight dynamics functions for
operational and new missions during FY 2001. Two new Windows 2000 based
machines were added to the lab equipment for use by 570 personnel. The
hardware in the FD Lab includes 6 NT/Windows 2000 workstations and 4
UNIX based machines for general use. The lab also houses the prime and
backup GNCC servers, the GNCC WEB servers and online storage in excess
of 1 terrabyte. The lab has tape back-up capability for this equipment.

The FD Lab began a consolidation of GNCC computing resources during FY
2001. This consolidation will allow for more efficient management of
GNCC IT infrastructure, expands access to Lab resources to all GNCC
personnel and provides for more uniform security procedures.

Plans for the coming year include the completion of the consolidation of
GNCC computing resources and the upgrade of the prime and back-up NT
based servers. This upgrade will provide additional online storage space
for the GNCC users.

\[Technical contact: Sue Hoge\]

**[5.3 FDAB WEB Page]{.underline}**

The FDAB WEB page was upgraded this year to include the latest
information on projects that the branch is supporting, links to other
branch related pages and pictures of all branch members. There is also a
flight dynamics WEB tool available through the WEB page. This tool
allows the user to perform some very basic flight dynamics analysis
functions. The URL for the FDAB WEB page is
[http://fdab1.gsfc.nasa.gov](http://fdab1.gsfc.nasa.gov/) .

\[Technical contact: Sue Hoge\]

# **[5.4 2001 FLIGHT MECHANICS SYMPOSIUM]{.underline}**

The Flight Mechanics Symposium, sponsored by the Guidance, Navigation
and Control Center (GNCC), Code 570, was held in the Building 3
Auditorium on June 19-21, 2001. The symposium provided an opportunity
for specialists in spacecraft flight dynamics to present, discuss, and
exchange information on a wide variety of topics such as attitude/orbit
determination, prediction and control; attitude simulation; attitude
sensor calibration; theoretical foundation of attitude computation;
dynamics model improvements; autonomous navigation; constellation design
and formation flying; estimation theory and computational techniques;
Earth environment mission analysis and design; and, spacecraft re-entry
mission design and operations. Forty-two technical papers were presented
by participants from NASA, other U.S Government agencies, private
industry and academia.

\[Technical contact: John Lynch\]

**[6.0 Interagency Activities]{.underline}**

## [6.1 GSFC Standards Program]{.underline}

The FDAB supports the GSFC standards program, the Data Standards
Steering Council (DSSC), and the Consultative Committee for Space Data
Systems (CCSDS).

The GSFC standard program aims to expand the scope of best practices,
and to develop an agency-endorsed database of preferred technical
standards for NASA.

The Data Standards Steering Council (DSSC) is the hub of the NASA Data
Systems Standards Program and is sponsored by the SOMO Chief Engineer.

The Consultative Committee for Space Data Systems (CCSDS) is an
international organization of space agencies interested in mutually
developing standard data handling techniques, to support space research
conducted exclusively for peaceful purposes.

The CCSDS Sub-Panel P1J is specifically chartered to investigate and
recommend Navigation Data standards. P1J has a membership representing
several international agencies. The work of P1J is accomplished
primarily at workshops, conducted at least twice a year, at facilities
coordinated by the hosting member agency. The main task of P1J is to
develop preferred standards for the exchange of navigation data. The
fiscal year 2001 workshops were conducted at the European Space Agency
(ESA) Vilspa facility, Spain, in October, and the ESOC facility,
Germany, in May.

P1J completed a green book (technical report), titled \"Navigation
Definitions and Conventions\", which was formally released for
distribution in July 2001; and a red book, titled \"Orbit Data
Messages\", which proposes a recommendation for space data systems
standards for the exchange of spacecraft orbit information. This red
book was released for official review by all CCSDS member agencies, in
July 2001, and is expected to be approved by the end of the calendar
year 2001. Following approval the red book will be promoted to blue book
status, formalizing the recommendation as an accepted preferred
standard.

Future work of P1J will involve developing new technical reports and
recommendations for navigation data exchange in support of proximity
operations, tracking, attitude, time services, environmental models and
astrodynamic constants.

For information about CCSDS and the GSFC standards program please refer
to

http://www.ccsds.org/

http://joy.gsfc.nasa.gov/GTSP/

\[Technical contact: Felipe Flores-Amaya\]

######## 

######## **[6.2 X-43A Anomaly Investigation]{.underline}**

<http://www.dfrc.nasa.gov/Projects/hyperx/developments.html>

Hyper-X is a NASA multi-year hypersonic flight research program seeking
to advance the state of the art through air-breathing hypersonic flight.
The goal of the Hyper-X program is to flight validate key propulsion and
related technologies for air-breathing hypersonic aircraft. The program
consists of three X43 vehicles, which will fly at speeds of Mach 7 and
10. Each of the vehicles is 12 feet long with a span of about five feet
(Figure 6-1). The first X43 and its modified Orbital Sciences
Corporation (OSC) Pegasus-XL booster rocket were launched on June 2,
2001 at about 1:43 p.m. from NASA\'s B-52 launch aircraft flying at
about 24,000 feet altitude. The flight was terminated when a major
malfunction occurred about eight seconds into the boost phase, causing
the X43-A vehicle to lose control. In support of the failure
investigation board, FDAB personnel provided support in (a) linear and
nonlinear analysis and implementation of the longitudinal autopilot, (b)
assessing controls-structures-interaction issues, and (c) validating
sensor models.

The OSC's Autopilot longitudinal design and linear analysis model were
examined. An independent Simulink and INCA model for one time step were
developed to mitigate the integrity of OSC's model and analysis.
Concerns of digital implementation of the integrator and filter,
insufficient frequency range for analysis and the sensitivity of gain
margin to rate gain were raised to the board of investigation.
Recommendations for future improvement were suggested.

The OSC 6 DOF non-linear model and the flight data load were examined.
In the process of reviewing, discrepancy of control gain profile of path
steering guidance between linear model and MDL was uncovered which
raised a question on the accuracy of the implementation of the design in
the actual flight data. Digital implementation of the autopilot in the
flight code was also examined and they appeared to be adequate. It was
recommended to improve the digital implementation method of the
integrator (use Tustin to replace forward transform) to ensure
stability.

One of the fault trees considers that the loss of vehicle control may
have resulted from a controls-structures-interaction (CSI) brought about
by the mis-modeling of the vehicle structural dynamics or an incorrect
implementation of the structural dynamic properties into the linear and
nonlinear analysis. The flexible body dynamics of the vehicle were
independently derived for both the longitudinal and lateral/directional
loops. Errors in modeling were identified and corrected in the new
model. A technique based on the balanced realization algorithm was used
to reduce the order this system for analysis. To assess potential
controls-structures interaction issues with Hyper-X a discrete-time
linear stability analysis was performed using the independently
developed flexible dynamics models. This analysis was followed by a
limited perturbation analysis to assess the effects of variations in
modal parameters and time delay. The results of these analyses indicated
that errors in structural dynamics modeling and implementation were not
a factor in the Hyper-X mishap. Recommendations for flexible body model
improvement, model reduction, and further investigation of time delays
were made to the board.

![](media/image20.jpeg){width="5.752777777777778in"
height="3.301388888888889in"}

Figure 1 Hyper-X attached to the Booster Rocket

\[Technical Contacts: Josephine San, Peiman Maghami, Jim Morrissey\]

**[7.0 Outreach Activities]{.underline}**

**[7.1 SAMPEX University Operations]{.underline}**

# The University of Maryland Aerospace Engineering Department completed its second full year of sole responsibility for flight dynamics support of the Solar Anomalous and Magnetospheric Particle Explorer (SAMPEX) spacecraft. In this role, a team of University of Maryland undergraduate and graduate students provides routine spacecraft orbit determination, attitude determination, attitude sensor analysis, and flight dynamics product generation. This effort is sponsored and supported by the FDAB, which provides consultation support as needed and periodically reviews the overall program status. This has been a very successful outreach initiative and gives the student team practical experience and training in spacecraft flight dynamics computations, the use of several commercial ground support tools and analysis of flight data. The operation also serves as a test bed for researching ground system automation techniques. 

\[Technical Contact: [Tom
Stengle](mailto:thomas.stengle@gsfc.nasa.gov)\]

**[7.2 Undergraduate Student Research Program (USRP)]{.underline}**

Code 572 sponsored Teddie Brinkley, a Junior from NC State University,
as part of the pilot NASA Undergraduate Student Research Program (USRP)
at Goddard. Teddie is a Mechanical Engineer who came to GSFC to perform
research in the area of formation flying technology. During his 3-month
stay at GSFC, Teddie was able to use both his programming and hardware
skills. He coded algorithms for the Formation Flying GPS Testbed under
the direction of Dr. Russell Carpenter and participated in the testing
of the Propulsion Branch\'s microNewton Thruster Test Stand with Chuck
Zakrzwski as his mentor. The thruster test stand is Code 574\'s
technology development project enabling the accurate testing of
micro-Newton thrusters to be used on nano satellites for constellations
and formations. Teddie especially enjoyed this opportunity to do
hands-on work with one of NASA\'s cutting edge technology development
efforts, as seen in the photo below. Code 572\'s experience with the
pilot USRP program was very positive. The USRP is definitely a benefit
for GSFC if we can get students of high quality and ability like Teddie
Brinkley.

\[Technical Contact: Karen Richon\]

**[7.3 PREST Program]{.underline}**

During FY01, the FDAB supported Nicholas Hamilton (USAF) under a grant
with the George Washington University Program of Research and Education
in Space Technology (PREST). This student is currently in residence at
the GSFC and is working with branch members on research of formation
flying control techniques.

\[Technical Contact: [Tom
Stengle](mailto:thomas.stengle@gsfc.nasa.gov)\]

**[7.4 Graduate Student Research Program (GSRP)]{.underline}**

# The FDAB continued its long standing support of the GSRP program. In FY01, the following GSRP efforts were underway: 

# 

-   "Decentralized Control of Distributed Satellite Networks."
    Researcher: Belanger, UCLA

```{=html}
<!-- -->
```
-   "Feasibility of Atmospheric Penetration for Satellite Formation
    Flying Experiment." Researcher: Joseph Schultz, University of
    Maryland

-   " Adaptive Satellite Attitude Control." Researcher: Kevin Walchdo,
    University of Florida

-   "Investigation of Libration Orbits in the Earth-Moon System."
    Researcher: Raquel Jarabek, University of Maryland

\[Technical Contact: [Tom
Stengle](mailto:thomas.stengle@gsfc.nasa.gov)\]

**[7.5 TEAMS Competition]{.underline}**

The Technology Education Alliance with Middle Schools (TEAMS) Program is
supported through an Educational Grant with NASA's Distributed
Spacecraft Technology Development Program. The focus of the TEAMS
program is on teamwork and the development of effective teaming skills
in middle school students. This focus is accomplished through teams of
students who design, develop, and operate teams of robots. The
student/robot teams compete in contests that stress the elements of
effective teaming (planning, communication, cooperation, and
coordination), engage the students through exciting hands-on
technologies, and challenge the students with realistic, non-trivial
problems to be solved. This year, hundreds of students from local middle
schools competed in robotic soccer. Branch volunteers served as
competition coordinators, judges and timekeepers in the one-day
competition that was held at Goddard Space Flight Center.

![](media/image21.jpeg){width="6.348611111111111in"
height="4.289583333333334in"}

\[Technical Contact: Tom Stengle\]

**[7.6 Public Education/Community Outreach]{.underline}**

A number of branch employees supported a variety of outreach activities.
These include:

**Science Fair & Engineering Judging or Career Day-**

-   Future City National Competition, Hyatt Regency, Capitol Hill, 400
    NJ Ave, NW, 2/20/01 (Dr. Aprille Ericsson)

-   District of Columbia Citywide Science Fair, Howard University, WDC,
    3/17/01 (Dr. Aprille Ericsson)

-   (Boys) Choir Academy of Harlem, 2005 Madison Ave., Manhattan, NY,
    5/18/01 (Dr. Aprille Ericsson)

-   South Carroll Covenant Keepers Homeschool Group, 3/01, (Richard
    Luquette)

**Student Engineering Design Projects-**

-   Take Our Daughter to Work Program, Egg Drop Contest, NASA GSFC,
    Greenbelt, MD, 4/26/01 (Dr. Aprille Ericsson)

-   NASA GSFC SISTER Program, \"Rocket Building and MAP Discussion\",
    Greenbelt, MD, 6/26/01 (Dr. Aprille Ericsson)

-   High School Botball Tournament, Wakefield High School, Arlington, VA
    (John Downing)

**Career Presentations-**

-   Buford Middle School, Charlottesville, VA, 10/17/00 (Dr. Aprille
    Ericsson)

-   Windows to the Universe, Family Science Night, Ballou Sr. HS, WDC
    10/25/00 (Dr. Aprille Ericsson)

-   2 HS, & Teachers/Staff mtg., Stockton, CA, 11/9/00 (Dr. Aprille
    Ericsson)

-   Junction City Middle School & Manhattan Middle School, Junction City
    & Manhattan, KS 1/18/01 (Dr. Aprille Ericsson)

-   Los Alamos Middle School, Los Alamos, NM, 1/23/01 (Dr. Aprille
    Ericsson)

-   District of Columbia SEED Charter School, 4300 C St, SE, WDC,
    2/12/01 (Dr. Aprille Ericsson)

-   QEM/MSE National Conference, "You meet the Scientist", JW Marriott
    Hotel, WDC, 2/3/01(Dr. Aprille Ericsson)

-   Engineers week Family Science night, National Bldg. Museum, WDC,
    2/21/01 (Dr. Aprille Ericsson)

-   Huron High School, A-A History & Goals Class, Ann Arbor, MI, 3/23/01
    (Dr. Aprille Ericsson)

-   Windows to the Universe, High School Classroom Visits, Montgomery,
    AL, 4/30/01-5/2/01 (Dr. Aprille Ericsson)

-   Windows to the Universe, Family Science Night, Tuskegee Univ.,
    Tuskegee, AL, 5/1/01 (Dr. Aprille Ericsson)

-   3 Philadelphia MS Schools, PENNLincs, Institute for Research in
    Cognitive Science, U.Penn, 5/11/01 (Dr. Aprille Ericsson)

-   (Boys) Choir Academy of Harlem, 2005 Madison Ave., Manhattan, NY,
    5/18/01 (Dr. Aprille Ericsson)

-   TRIO Math and Science Program, Howard University, WDC, 7/17/01 (Dr.
    Aprille Ericsson)

**Speeches (Open, Award, Closing & Graduation Ceremonies, Breakfast,
Luncheon, Dinner) -**

-   University of Virginia, Women 2000: Shapers of the World
    celebration, Charlottesville, VA 10/17/00 (Dr. Aprille Ericsson)

-   African American Chamber of Commerce Gala Awards Ceremony, Stockton,
    CA, 11/10/00 (Dr. Aprille Ericsson)

-   University of the Pacific, AMP, Stockton, CA, 11/10/00 (Dr. Aprille
    Ericsson)

-   National Alliance of Black School Educators, Convention Center,
    Philadelphia, PA, 11/17/00 (Dr. Aprille Ericsson)

-   Fernbank Science Ctr/NASA SEMAA, Imagine This: Science Here, There &
    Everywhere, ATL, GA 11/18/00 (Dr. Aprille Ericsson)

-   Mother/Daughter Tea, Celebrating the Past,Creating the Future,
    Andrew Jackson MS,Forestville,MD, 3/16/01 (Dr. Aprille Ericsson)

-   University of MD, 1st Annual Pre-College Program Conf., Adele Stamp
    Union, College Park, MD 4/21/01 (Dr. Aprille Ericsson)

-   NYU Summer Conf. on Urban Science and Math Teaching, \"Sharing Our
    Success\", NY, NY, 5/24/01 (Dr. Aprille Ericsson)

-   Howard Univ., Upward Bound Mathematics & Science Initiative,
    Blackburn Ctr. HU, WDC, 7/27/01 (Dr. Aprille Ericsson)

**Program Visits NASA GSFC/Mentor Student/Fellow-**

-   Sunbeams Program, NASA GSFC, Greenbelt, MD (Dr. Aprille Ericsson)

-   University of Maryland, Eastern Shore, Summer Program, MCC, NASA
    GSFC, Greenbelt, MD, 7/5/01 (Dr. Aprille Ericsson)

-   University of Maryland, Baltimore County, Myerhoff Program, NASA
    GSFC, Greenbelt, MD, 7/6/01 (Dr. Aprille Ericsson)

-   Delaware State University, Summer Program, NASA GSFC, Greenbelt, MD,
    7/13/01 (Dr. Aprille Ericsson)

-   FBI Conference Students visiting NASA GSFC, Greenbelt, MD, 7/31/01
    (Dr. Aprille Ericsson)

-   United Negro College Fund/NASA HQs Harriet Jenkins Fellows, NASA
    GSFC, Greenbelt, MD 8/4/01 (Dr. Aprille Ericsson)

## Seminar Presentations-

-   Lawrence Livermore National Laboratories, Science 2001 Lecture
    Series, LL HS, CA, 9/28/00 (Dr. Aprille Ericsson)

-   Kansas State Univ, Following Your Dreams: Life Lessons from a NASA
    Engineer, Manhattan, KS 1/18/01 (Dr. Aprille Ericsson)

-   Los Alamos National Laboratories, Los Alamos, NM, Dr. MLK Jr.
    Celebration, 1/23/01 (Dr. Aprille Ericsson)

-   Patent & Trade Office, "African-Americans in Technology",
    Alexandria, VA, 2/20/01 (Dr. Aprille Ericsson)

-   John Hopkins Univ.- APL Colloquim, "African-Americans in
    Technology", MD, 2/23/01 (Dr. Aprille Ericsson)

-   US Library of Congress, Inspiring Stories of Vision of Courage,
    James Madison Mem. Bldg., WDC, 3/6/01 (Dr. Aprille Ericsson)

-   MD Aviation Administration\'s, \"Women\'s History Month\", BWI
    Airport, Baltimore MD, 3/22/01 (Dr. Aprille Ericsson)

-   University of Michigan, IMPACT Program, Ann Arbor, MI, 3/23/01 (Dr.
    Aprille Ericsson)

-   Women In Technology Expo, World Bank Headquarters, Washington, DC,
    5/30/01 (Dr. Aprille Ericsson)

-   NASA GSFC Summer Interns, "How to make an Outstanding Technical
    Presentation" GSFC 6/12/01 (Dr. Aprille Ericsson)

**Education/Career Conference or Panel-**

-   NSBE-AE Region II Professional Development Conference, Gaithersburg
    Hilton, MD, 10/28/00 (Dr. Aprille Ericsson)

-   George Mason Univ. African American Studies, \"Cookies, Java Beans,
    and Other Digital Delicacies: Closing the Digital Divide\",
    George W. Johnson Ctr, Fairfax campus, VA, 11/4/00 (Dr. Aprille
    Ericsson)

-   Women's of Washington, TLT, Women & Power Tools\" Four Seasons
    Hotel, WDC, 3/14/01 (Dr. Aprille Ericsson)

-   Black Issues in Higher Learning National Policy Summit on Science,
    Mathematics, and Technology for African American Students, Reston,
    VA, 6/15/01 (Dr. Aprille Ericsson)

**Television/Radio/Magazine/Website/Newspaper Interviews-**

-   Woman Engineer, "To Give is to Receive", Anne Baye Eriksen, 10/00
    (Dr. Aprille Ericsson)

-   Emerging Markets Magazine, Lillianne Sy, Traci Jones, 10/00 (Dr.
    Aprille Ericsson)

-   The Stockton Record, Stockton CA, Sarah Grunder, 11/10/00 (Dr.
    Aprille Ericsson)

-   Kansas State Univ. Radio Station, Manhattan, KS, Dr. Suzanne E.
    Franks, 1/18/01 (Dr. Aprille Ericsson)

-   Current Biography, Christopher Luna, Bronx, NY, 2/01 (Dr. Aprille
    Ericsson)

-   Space Day, Air & Space, PA Radio Interview, Devilier Assoc.,
    Gretchen Fox , (Dr. Aprille Ericsson)

#  **Proposal/Application Reviewer-**

-   Intel Science Talent Search, 1719 & 1723 N St, NW, WDC, 12/00 (Dr.
    Aprille Ericsson)

-   NASA CSTEA Review, Howard University, Founders Library, WDC, 2/5/01
    (Dr. Aprille Ericsson)

-   HU GSAS Mechanical Engineering review, Howard University, Founders
    Library, WDC, 3/12/01 (Dr. Aprille Ericsson)

-   HU Science, Engineering & Mathematics Program Advisory Board, HU
    School of Engineering, WDC, 4/4/01 (Dr. Aprille Ericsson)

-   HU Graduate School, Responsive Ph.D. Initiative Task Force, HU GSAS,
    WDC, 4/4/01 (Dr. Aprille Ericsson)

**[Appendix A -- Goddard and NASA Awards]{.underline}**

[Team Awards]{.underline}

EO-1 Flight Dynamics Launch Support Team Outstanding Teamwork Award

GSFC COE group achievement award to the GNC team for the GRO re-entry

Outstanding Teamwork Award - EOS AM Project Team

MAP Trajectory Team Customer Service Excellence Award

MAP Monte Carlo Tool Development and Implementation Customer Service
Excellence Award

MAP Thermal Vacuum/Thermal Balance Test Team Customer Service Excellence
Award

MAP Integration & Test Team Customer Service Excellence Award

MAP Comprehensive Performance Test Team Customer Service Excellence
Award

NASA Group Achievement Award for the CGRO Reentry Team

[Individual Goddard/NASA level Awards]{.underline}

Goddard Civil Service Excellence (Lauri Newman)

Customer Service Excellence Award for EO-1 support (David Quinn)

Goddard Award of Merit (Robert DeFazio)

**[Appendix B - University Grants]{.underline}**

The following university grants being administered by FDAB engineers
were in place in FY00:

1.  GRANT NAG5-9961 with the University of Maryland Department of
    Aerospace Engineering titled "Precise Virtual Rigid Body Control of
    a Satellite Constellation." This grant is developing a possible
    control strategy for formation flying.

> \[Technical Contact: [Thomas
> Stengle](mailto:thomas.stengle@gsfc.nasa.gov)\]

2.  GRANT NAG5-9890 with the University of Maryland Department of
    Aerospace Engineering titled "Rarefied Flow Aerodynamics for
    Stability and Control of Formation-Flying Satellites." This grant is
    researching problems and control strategies for spacecraft flying in
    formation with low perigee passes. This research may benefit the
    development of control approaches for the Geospace Electrodynamics
    Connections (GEC) mission.

> \[Technical Contact: [Marco
> Concha](mailto:marco.concha@gsfc.nasa.gov)\]

3.  GRANTS NAG5-8694 and NAG5-8879 with the University of California at
    Los Angeles titled "Decentralized Estimation and Control of
    Distributed Spacecraft," and "Precise Relative State Estimation and
    Control of Distributed Satellite Networks." These grants are
    developing and applying new decentralized control architectures for
    satellite formations.

> \[Technical Conract: [Russell
> Carpenter](mailto:russell.carpenter@gsfc.nasa.gov)\]

4.  GRANT NAG5-9829 with the University of Texas at Austin titled
    "Spacecraft Rendezvous Navigation with Integrated INS-GPS." This
    grant is focusing on GPS/INS software architecture development for
    relative navigation and attitude determination.

> \[Technical Contact: [Russell
> Carpenter](mailto:russell.carpenter@gsfc.nasa.gov)\]

5.  GRANT NAG5-9612 with Cornell University Sibly School of Mechanical
    and Aerospace Engineering titled "New Algorithms for Magnetometer
    Orbit and Attitude Estimation." This grant is studying the
    feasibility of a moderate precision navigation (\<10 km orbit, \<0.5
    degrees attitude) using Magnetometer data.

> \[Technical Contact: [Richard
> Harman](mailto:richard.harman@gsfc.nasa.gov)\]

6.  GRANT NAG5-9748 with Princeton University Department of Mechanical
    and Aerospace Engineering titled " Satellite Attitude Estimation
    with the Two Step Optimal Estimator." This grant is studying the
    ability of the two-step algorithm to out perform the standard
    Extended Kalman Filter currently used for spacecraft and ground
    attitude estimation.

> \[Technical Contact: [Richard
> Harman](mailto:richard.harman@gsfc.nasa.gov)\]

7.  GRANT NAG5-11331 with State University of NY at Buffalo titled,
    \"Attitude Determination Schemes for the CEGANS Sensor." The CEGANS
    concept is to perform spacecraft attitude determination by
    considering the sightline vectors of GPS SVs visible to each antenna
    element of a multi-element array fixed to the user spacecraft.
    Simulation data provided by NASA-Goddard will be analyzed at the
    University of Buffalo in order to investigate robust and optimal
    attitude determination schemes for the CEGANS sensor.

> \[Technical Contact: David Quinn\]

8.  GRANT NAG5-10563 with the University of Maryland Department of
    Aerospace Engineering titled \"Automation of SAMPEX Orbit
    Determination.\" This grant is researching the automation of the
    orbit determination of the SAMPEX satellite through the automation
    of the following phases: data acquisition, data processing, and data
    output.

\[Technical Contact: Joe Toth\]

**[Appendix C -- Conferences and Papers]{.underline}**

Given below is a list of journal papers, professional papers and
technical presentations that were prepared and delivered in FY01 by
branch members.

JOURNAL ARTICLES:

R. Azor, I.Y. Bar-Itzhack, J. Deutschmann (now Thienel), and R.R.
Harman, "Angular-Rate Estimation Using Delayed Quaternion Measurements",
*AIAA* *Journal of Guidance, Control, and Dynamics,* Vol. 24, No. 3,
May-June 2001, pp. 436-443.

J.K. Deutschmann (now Thienel) and I.Y. Bar-Itzhack, "Evaluation of
Attitude and Orbit Estimation Using Actual Earth Magnetic Field Data",
*AIAA Journal of Guidance, Control, and Dynamics*, Vol. 24, No. 3,
May-June 2001, pp. 616-623.

Carpenter, J. Russell, \"Decentralized control for satellite
formations,\" final draft accepted for publication in International
Journal of Robust & Nonlinear Control

Carpenter, J. Russell and Schiesser, Emil S. \"Semi-major axis knowledge
and GPS orbit determination\" to appear in upcoming issue of NAVIGATION

CONFERENCE:

***[24^th^ Annual AAS Guidance and Control Conference, Breckenridge Co,
Jan. 31-Feb. 4, 2001]{.underline}***

Julie Deutschmann (now Thienel), Itzhack Bar-Itzhack, and Rick Harman,
"A LEO Satellite Navigation Algorithm Based on GPS and Magnetometer
Data"

S. Hoge & F. Vaughn (GSFC), "Trajectory Design and Control f0r the
Compton Gamma Ray Observatory Re-entry"

J. Bolek, E. Holmes, J. O\'Donnell, P. Sabelhaus, S. Scott, and J.
Story, \"On-Orbit ACS Performance of the Landsat 7 Spacecraft\"

***[IEEE Aerospace Conference, Big Sky Montana, March 11-17,
2001]{.underline}***

Steven P. Hughes (GSFC) and Laurie M. Mailhe (a.i.solutions), \"A
Preliminary Formation Flying Orbit Dynamics Analysis for
Leonardo-BRDF\",

***[AIAA Aerodynamic Decelerator Systems conference, May 2001,
Boston]{.underline}***

Carpenter, J. Russell, \"Trajectory Reconstruction,\" (invited seminar)

***[Flight Mechanics Symposium, NASA Goddard Space Flight Center,
Greenbelt, MD, June 19-21, 2001]{.underline}***

Julie Thienel (GSFC) and R.M. Sanner Univ. of Md.), "A Nonlinear
Spacecraft Attitude Controller and Observer with an Unknown Constant
Gyro Bias and Gyro Noise"

Rich Luquette (GSFC) and Rob Sanner (Univ. of Md.), "A Nonlinear
Approach to Spacecraft Trajectory Control in the Vicinity of a Libration
Point"

D. Kelbel, T. Lee & A. Long (CSC), and R. Carpenter & C. Gramling
(GSFC)., "Evaluation of Relative Navigation Algorithms for
Formation-Flying Satellites"

D. Folta (GSFC), C. Youn (Univ of CO), A. Ross (Harvard Univ) , "Unique
Non-Keplerian Orbit Vantage Locations for Sun-Earth Connection and Earth
Science Vision Roadmaps*"*

D. McGiffin, M. Mathews (CSC), and S. Cooley (GSFC), "High Earth Orbit
Design for Lunar-Assisted Medium Class Explorer Missions"

S. Hughes (GSFC), L. Mailhe (aisolutions), "A Preliminary Formation
Flying Orbit Dynamics Analysis for Leonardo-BRDF"

S. Belur (CSC) & R. Harman (GSFC) , "Calibration of Gyros with
Temperature Dependent Scale Factors"

I. Bar-Itzhack (Technion) & R. Harman (GSFC) , "In-Space Calibration of
a Gyro Quadruplet"

I. Bar-Itzhack (Technion) & R. Harman (GSFC), "State-Dependent
Pseudo-Linear Filter for Spacecraft Attitude and Rate Estimation"

J. Chen, W. Morgenstern & J. Garrick (GSFC), "Triana Safehold: A New
Gyroless Sun-Pointing Attitude Controller"

J. O'Donnell, W. Morgenstern, M. Bartholomew (GSFC), "Using Automation
to Improve the Flight Software Testing Process"

S. Starin & J. O'Donnell (GSFC) , "A Two-wheel Observing Mode for the
MAP Spacecraft"

D Folta (GSFC) & A. Hawkins (AI Sol), "Preliminary Results of NASA's
first Autonomous Formation Flying Experiment: EO-1"

S. Hoge & F. Vaughn (GSFC), "Trajectory Design and Control f0r the
Compton Gamma Ray Observatory Re-entry"

N. Ottenstein, M. Challa, & A. Home (CSC), and R. Harman & R. Burley
(GSFC), "IMAGE Mission Attitude Support Experiences"

D.A. Quinn (GSFC), P. Sanneman, S. Shulman, J. Sager (Swales), "The
Integration, Testing and Flight of the EO-1 GPS."

***[Institute of Navigation 57th Annual Meeting, Albuqurque, NM, June,
2001]{.underline}***

J.L. Garrison, M.C. Moreau (GSFC), P. Axelrad (Univ. Colorado), \"
Tracking Loop Optimization for On-Board GPS Navigation in High Earth
Orbit (HEO) Missions,\"

***[2001 American Controls Conference, Arlington, June
25-27]{.underline}***

Starin (GSFC), R. K. Yedavalli (Ohio State University) & Andrew Sparks
(VACA/AFRL)}, "Design of a LQR Controller of Reduced Inputs for Multiple
Spacecraft Formation Flying"

***[SIAM mini-symposium on Control, Mission Design, and Satellite
Dynamics, July 2001, San Diego.]{.underline}***

Carpenter, J. Russell, \"Distributed Spacecraft Control Architectures\"
(invited presentation)

***[AIAA Guidance, Navigation, and Controls Conf., Montreal, Quebec,
Canada, August 6-9, 2001]{.underline}***

P. G. Maghami (GSFC), and D. E. Cox (LaRC), "Control of Flexible Systems
in the Presence of Failures"

Starin (GSFC) , R. K. Yedavalli (Ohio State University) & Andrew Sparks
(VACA/AFRL), "Spacecraft Formation Flying Maneuvers Using
Linear-Quadratic Regulation with No Radial Axis Inputs"

***[AAS/AIAA Astrodynamics Specialists Conference, Quebec City, Canada,
July 30 August 2, 2001]{.underline}***

Rich Luquette (GSFC/GNCC) and Rob Sanner (Univ. of Md.), "A Nonlinear
Approach to Spacecraft Formation Control in the Vicinity of a Collinear
Libration Point"

Robert L. DeFazio (GSFC), Skip Owens (a.i.solutions), Susan Good
(a.i.solutions) \"Follow that Satellite: EO-1 Maneuvers into Close
Formation with Landsat-7\"

Sue Hoge (GSFC) and Frank Vaughn (GSFC), \"Trajectory Design and Control
for the Compton Gamma-Ray Observatory Re-entry\"

Rich Luquette (GSFC) and Rob Sanner (Univ. of Md.), "A Nonlinear
Approach to Spacecraft Formation Control in the Vicinity of a Collinear
Libration Point"

***[Space 2001 Conference, Albuqurque, NM, August, 2001]{.underline}***

David Quinn, R.E. Farley, "Tethered Formation Configurations: Meeting
the Scientific Objectives of Large Aperture and Interferometric Science"

***[Institute of Navigation 14^th^ International Technical Meeting, Salt
Lake City, UT, Sept. 11-14. 2001]{.underline}***

Julie Thienel, Itzhack Bar-Itzhack (Technion Institute of Technology),
and Rick Harman, "GPS/Magnetometer Based Satellite Navigation and
Attitude Determination"

M.C. Moreau (GSFC) , P. Axelrad (Univ. Colorado), J.L. Garrison, M.
Wennersten (GSFC) , A.C. Long CSC), \" Test Results of the PiVoT
Receiver in High Earth Orbits using a GSS GPS

Simulator,\"

***[40^th^ IEEE Conference on Decision and Control, Orlando, FL,
December 4-7, 2001]{.underline}***

Julie Thienel (GSFC) and R.M. Sanner (Univ. MD), "A Coupled Nonlinear
Spacecraft Attitude Controller/Observer with an Unknown Constant Gyro
Bias"

**[Appendix D- Acronyms and Abbreviations]{.underline}**

This appendix gives the definitions of acronyms used in this document.

AAS American Astronautical Society

ACS Attitude Control System

ACT Attitude Control Thrusters

AETD Applied Engineering and Technology Directorate

AI Artificial Intelligence

ALI Advanced Land Imager

AO Announcement of Opportunity

APL Applied Physics Laboratory

AST Autonomous Star Tracker

ATMS Advanced Technology Microwave Sounder

CCSDS Consultative Committee for Space Data Systems

CETDP Cross Enterprise Technology Development Program

CGRO Compton Gamma Ray Observatory

COTS Commercial Off-the-Shelf

CPT Comprehensive Performance Test

CSOC Consolidated Space Operations Contract

CVS Concurrent Version System

DACC Distributed Active Archive Center

DoD Department of Defense

DSN Deep Space Network

DSS Digital Sun Sensor

DST Dynamical Systems Theory

EFF Enhanced Formation Flying

EMOS EOS Mission Operations System

EO Earth Observing

EOS Earth Observing System

ESA European Space Agency

ESSP Earth System Science Pathfinder

EUVE Extreme Ultraviolet Explorer

FAA Federal Aviation Administration

FDAB Flight Dynamics Analysis Branch

FDS Flight Dynamics System

FDSS Flight Dynamics Support System

FDF Flight Dynamics Facility

FOT Flight Operations Team

FSW Flight Software

FY Fiscal Year

GEO Geosynchronous Earth Orbit

GEODE GPS Enhanced Orbit Determination Experiment

GEONS GPS-Enhanced Orbit Navigation System

GINA Generalized Information Network Analysis

GNCC Guidance, Navigation, and Control Center

GOES Geostationary Operational Environmental Satellite

GPM Global Precipitation Mission

GPS Global Positioning Satellite

GRO Gamma Ray Observatory

GSE Ground Support Equipment

GSFC Goddard Space Flight Center

GSRP Graduate Student Research Program

GTDS Goddard Trajectory Determination System

GUS Gyroscopic Upper Stage

HD Henry Draper

HDS Hybrid Dynamic Simulator

HEO High Earth Orbit/ Highly Elliptical Orbit

HGA High Gain Antenna

HTML HyperText Markup Language

I&T Integration and Test

ICD Interface Control Document

IHS Inner Heliospheric Sentinels

IM Ionosphere Mapper

IMDC Integrated Mission Design Center

IMU Inertial Measurement Unit

ISU International Space University

ITAR International Traffic In Arms Regulation

ITSO International Telecommunications Satellite Organization

JPL Jet Propulsion Laboratory

LEO Low Earth Orbit

LOR Launch and Orbit Raising

LPT Low Power Transceiver

LQG Linear Quadratic Gaussian

LRR Lightweight Rainfall Radiometer

MAP Microwave Anisotropy Probe

## MARSAT Mars Areo-stationary Relay Satellite 

MC Master Catalog

MCC Mid Course Correction

MCO Mars Climate Observer

MIT Massachusetts Institute of Technology

MLS Microwave Limb Sounder

MMS Magnetic Multi-scale Mission

MOC Mission Operations Center

MOCC Mission Operations Command and Control

MODIS Moderate Resolution Imaging Spectroradiometer

MOPSS Mission Operations Planning and Scheduling System

MOST Mission Operations Support Team

MOWG Mission Operations Working Group

MSRD Mission Specific Requirements Document

NASA National Aeronautical and Space Administration

NGST Next Generation Space Telescope

NMM Normal Maneuver Mode

NOAA National Oceanic and Atmospherics Administration

NPB Navigation Processor Board

NPM Normal Pointing Mode

NRTS Network Resources and Training Sites

NSF National Science Foundation

NT New Technology

OAT Orbit Adjust Thrusters

ONS Onboard Navigation Systems

OSSM Ocean Surface Salinity Mission

PC Personal Computer

PI Principal Investigator

PLT Post Launch Testing

PREST Program of Research and Education in Space Technology

QuikSCAT Quick Scatterometer

R&D Research and Development

RBM Radiation Belt Mapper

RMS Root-Mean-Square

RPO Radiation Protection Office

RSDO Rapid Spacecraft Development Office

RTOD Real-time Orbit Determination

RWA Reaction Wheel Assembly

RXTE Rossi X-Ray Timing Explorer

SA Selective Availability

SAMPEX Solar Anomalous and Magnetospheric Particle Explorer

SMEX Small Explorer

SOHO Solar and Heliospheric Observatory

SOMO Space Operations Management Office

SPECS Evolution of Cosmic Structure

SPS Standard Positioning Service

ST Space Technology

TDRSS Tracking Data Relay Satellite System

TMM Thruster Maneuver Mode

TONS TDRSS Onboard Navigation System

TRACE Transition Region and Coronal Explorer

TRMM Tropical Rainfall Measuring Mission

URL Uniform Resource Locator

USN Universal Space Network

VCM Velocity Control Mode

VIIRS Visible Infrared Imaging Radiometer Suite

WAAS Wide Area Augmentation System

WIRE Wide-Field Infrared Explorer

WRS World Reference System

WWW World Wide Web