APPLIED ENGINEERING AND TECHNOLOGY DIRECTORATE

 

Flight Dynamics Analysis Branch 

End of Fiscal Year 2001 Report 

A. Barnes, T. Stengle, S. Truong 

 

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 

1.0 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

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 stateof-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 Landsat7). 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

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

2.1.1 Triana

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

Triana Trajectory Design

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)

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 onboard 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)

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 standalone 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/

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, variableemissivity 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. Figure 2.1 ST-5 Spacecraft 

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, M. Woodard, M. Concha] 

2.1.3 Aqua Earth Observing System (EOS) (launch 3/02)

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] 

2.1.4 Aura EOS (launch 11/2003)

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. 

Aura: 15 to 22 minutes

behind Aqua

Aqua 

Aura ground track 

Aqua ground track

Aura MLS looking forward to

where Aqua had been

looking a short time before

Aura MLS field of

view intersects

Aqua ground track

at the Earth’s limb 

Figure 2-2. Relative Positions of Aqua and Aura Showing MLS Viewpoint 

[Technical contact: L. Newman] 

2.2 Operational Missions

2.2.1 Microwave Anisotropy Probe (MAP)

http://map.gsfc.nasa.gov

Background

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. 

CSS (6) 

RWA (3) DSS (sun side) 

IRU (2) 

AST 

(2)

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

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 L2 lunar shadows, lunar shadow avoidance at L2, 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 L2 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 L2 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 L2, the spacecraft will 

maintain a Lissajous orbit such that the MAP-Earth vector remains between 0.5° and 10.5° off the SunEarth 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 L2 Lagrange point for a period of two years. 

Figure 2-4: Representation of the MAP 3-Loop Trajectory 
Table 2-1: MAP Maneuvers 

Maneuver 

Location Description

Maneuver Start 

UTC Date & Time 

Delta-V 

(m/s) 

Statu

s 

A1 Calibration of 

Thrusters July 04 13:22:38 2.01 

P1 Apogee Raising 08 04:43:40 20.19 

A2 Engineering Burn 12 16:11:54 0.25 

P2 Apogee Raising 17 03:38:25 2.51 

A3 Engineering burn 21 18:54:43 0.30 

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 Sep. 14 16:37 0.04 

L2 LOI Dec. 14 16:37 TBD 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; S. Andrews] 

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, R. Luquette, C. Mendelson] 
2.2.3 Geostationary Operational Environmental Satellite (GOES)

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 sevenorbit 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] 

2.2.4 Landsat-4 Decommissioning

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.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] 
2.2.5 Tropical Rainfall Measuring Mission (TRMM) 

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-oflife 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, J. Morrissey] 

2.2.6 InFocus Balloon Program 

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. 

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. 

Figure 2-5 InFocus During Ascent 

[Technical contact: D. Olney] 

2.2.7 General Space Operations Management Office (SOMO) Support

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] 
3.0 Study Mission Support

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)

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/. 

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

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 (RE) and apogee near 

lunar distance (≈ 60 RE). 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 LunarAssisted 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.
Figure 3-1 KRONOS Direct Transfer HEO (5 year

Mission), View in Earth – Moon Rotating Coordinates

(June 1)

Figure 3-2 KRONOS Direct Transfer HEO (Post Lunar

Encounter Apogee), View from North Ecliptic Pole in

Solar Rotating Coordinates (June 1)

Figure 3-3 KRONOS Direct Transfer HEO (Post Lunar Encounter

Apogee), View in ecliptic Plane in Solar Rotating Coordinates (June 1)

Lunar

Encounter

Perigee

Apogee

S/C Trajectory

Lunar Orbit

Lunar

Encounter

Sun Line

(ecliptic plane)

Apogee

Sun Line

(ecliptic plane)

Lunar Orbit

S/C Trajectory

Apogee Lunar

Encounter

E

E

[Technical contact: Steven Cooley] 

3.3 Constellation X

Constellation X is a study mission that uses 2 

(possibly 4) x-ray telescopes in constellation at the 

Earth’s L2 libration point to study black holes and 

galaxy formation. The instrument consists of a large 

area X-ray mirror with 100-meter focal length. This 

year the FDAB has 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 lowradiation 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. 

Figure 3-4 Constellation X Concept 
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 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. Figure 3-5 Sample Constellation X 

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

[Technical contact: Lauri Newman, Dave Folta]

3.4 Global Precipitation Mission (GPM) 

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 Coop 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 . 

Figure 3-6 Global Precipitation Mission

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. 

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. 

Figure 3-7 Results of Optimization Tool

[Technical contacts: Chad Mendelsohn, David Folta] 

3.5 Next Generation Space Telescope (NGST)

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. 

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. 

Figure 3-8 Next Generation Space Telescope 
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, L2. 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 L2 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. 

Solar Rotating Coordinates

Ecliptic Plane Projection

L2

+x-axis

C3 Energy km2/s2

 0.00

- .25

- .50

- .75

-1.00

Maneuvers 1,3

+y-axis

Maneuvers 2, 4

Solar Rotating Coordinates

x-z plane

Solar Rotating Coordinates

y-z plane

C3 Energy

Ecliptic Plane Projection in

Solar Rotating Coordinates

Ecliptic Plane Projection in

Solar Rotating Coordinates

Y-Z axis View in Solar

Rotating Coordinates

Maneuvers in blue

Figure 3-9 NGST Trajectory Planning

[Technical Contact: Peiman Maghami, David Folta ] 

3.6 Living With A Star: Solar Dynamics Observatory (SDO)

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. 
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. 

Figure 3-10 Solar Dynamics Observatory Trajectory 

[ Technical Contact: David Folta]

3.7 Magnetospheric Multiscale Mission (MMS)

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. 

[Technical contact: Charles Petruzzo] 

3.8 Laser Interferometer Space Antenna (LISA)

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. 

 Figure 3-11 Laser Interferometer Space 

[Technical Contacts: Peiman Maghami, Steve Hughes] 

3.9 Leonardo

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. 
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. 

Figure 3-12 Leonardo Mission Concept [Technical Contact: Steve Hughes] 
4.0 Technology Development Activities

4.1 Advanced Mission Design

4.1.1 Generator Software Tool

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). 

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

[Technical Contact: David Folta] 
4.1.2 SBIR Phase II Research: Optimal Orbit Transfer Analysis for Advanced 

Space Systems

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]

4.1.3 Optimal Mission Design

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. 

0 200 400 600 -100

-80

-60

-40

-20

0

Function Evaluation

Cost Function, days

F(x) vs. Count

0 200 400 600 0

5

10

15

Function Evaluation

Maximum Eclipse Duration, hrs

Max Eclipse vs. Count

0 200 400 600 0

1

2

3

Function Evaluation

Minimum Periapsis, Re

Minimum Periapsis vs. Count

0 200 400 600 16

17

18

19

20

21

Function Evaluation

Maximum Apoapis, Re

Maximum Apoapis vs. Count

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.

0 0.5 1 1.5 2

4.9

4.95

5

5.05

5.1

x 10

9 Range vs. Time

Time, years

Range, m

Arm12

Arm23

Arm31

0 0.5 1 1.5 2

-10

-5

0

5

10

Range Rate vs. Time

Time, years

Range Rate, m/s

Arm12

Arm23

Arm31

0 0.5 1 1.5 2

59

59.5

60

60.5

61

Angle vs. Time

Time, years

Angle, deg.

Angle12

Angle23

Angle31

0 0.5 1 1.5 2

-1

0

1

x 10

-7Angle Rate vs. Time

Time, years

Angle Rate, deg/s

Angle12

Angle23

Angle31

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] 
4.2 Autonomous Navigation Technologies

4.2.1 Autonomous Onboard Navigation Systems 

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, 

Ground-Based Navigation Scenario Autonomous Navigation Scenario

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 

Satellite Operations Ce nter

P red icted

Position Data

6

Trackin g

 Data

4

2 3

1

S /C to S un

direct ion al

 measurement

Earth

S /C to E arth

directional

 measurement

G roun d Station

 to S/C Doppler

 measurement

GPS
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. 

 

Figure 4-5 PiVoT Receiver 

Flight Experiments 

The FDAB navigation team

provided extensive analysis of the 

autonomous navigation accuracy

being achieved for the EO-1

mission using GPS, in support of

the Extended Formation Flying

(EFF) experiment. We performed 

post-facto processing of the raw 

measurement data to evaluate 

achievable accuracies, and 

filtering of the Loral Tensor GPS 

receiver’s point solutions and 

filtered solutions to assess the 

resultant improvement in these 

solutions. We provided a detailed

summary of resultant navigation

accuracies that can be achieved

using the various solutions. We 

also assessed the quality of the 

operational S-band solutions. 

Based on this analysis, we 

prepared the technical report 

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. 

Figure 4-6. Formation Flying 

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 rootmean-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. 

GPS orbits

Main lobe

First side lobe

42.6 degrees (L1)

GEO

LEO (below

3000 km)

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, David Folta, Cheryl Gramling]
4.2.2 Adaptive Kalman Filter for Autonomous Navigation 

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 autocorrelated 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. 

 

 

N AVIG A TIO N

(EKF)

T U N IN G SU B SY STE M

(Fuzzy S ys tem +

P reproc ess or)

G P S

D ata

S et

Filter

[Q ],

[R ]

G P S E phem ,

M easurem ents

Estim ated States

Tuning P aram eters

T R A IN IN G SYS TE M

(N euro-Fuzz y S ys tem +

P reproc ess or)

T R A IN IN G

EKF

T R AIN IN G

D A TA

P R O C E SSIN G

[P], [M ], Filter Status

Tuning P aram eters

[Q ], [R ]

[P ], [M ], Filter Status

Training +

C hec k ing D ata

P aram eters to build

Fuzzy S ys tem

G P S E phem ,

M easurem ents

 

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 retuning 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. 

and

or

not

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] 
4.2.3 Magnetometer Based Navigation 

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. 

 

GPS Signals

To ground 

To control system

Sensor MAGNAV 

Magnetic Field 

Measurements 

Attitude, Orbit, Rate 

‘Black Box’

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. 
Table 4-1. MAGNAV Performance 

Sensor Combination Orbit Attitude Rate 

MAG+GYRO 15-25 km 0.2-1.4 deg Gyro dependent

MAG+SUN 10-40 km <1 deg <0.003 deg/sec

MAG+GPS ~meters <0.3 deg <0.003 deg/sec

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, Rick Harman] 

4.3 Formation Flying Technologies

4.3.1 EO-1 Formation Flying Experiment 

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. 
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. 

Figure 4-11. EO-1 Following Landsat-7

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.

 FF Start

Velocity

FF Maneuver

Formation Flying Spacecraft

Reference S/C

In-Track Separation (Km)

Observation Overlaps

Radial Separation (m)

Ideal FF Location

Nadir

Direction

I-minute separation

in observations

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, EO1 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] 

4.3.2 SBIR Phase II Research: Autonomous Unified Orbit and Attitude Control for 

Formation Flying using GPS and LQG/LTR Controller

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. 

EO 1

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. 

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.

Figure 4-13. Spacecraft in Formation Flying

[Technical Contact: David Folta ]

4.3.3 Decentralized Estimation and Control of Distributed Spacecraft 

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 Figure 4-14. Distributed Spacecraft
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. 

z-1

z-1

Yi

j

Xi

ui

j

α i

lj

,l = 1,2 ,...,K; l ≠ j

αi

jj hi

j

xˆ i

Dj xi

Dj

Xi

C

 

ui−1

l , l =1, 2, ...,K ;l ≠ j

αi

jl

To Network

From

Network

Glob. Data

Reconstr.

Glob. Opt.

Control

 X R

σi

j

 

σ i

l

From

Network To Network

Loc. Opt.

LKF

z-1

z-1

Actuator Plant

Sensor

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 stateestimates 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 stateestimate 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. L1, L2, and L3. 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.

Off

Standby

Wait

Count

down

Wait

Rcv

Inst

Rcv

State Vect

Rcv

Meas

Xmt

Semf

Rcv

Semf

Rcv

Ctrl

Rcv

Alph Xmt Ctrl

Xmt Alph

Clock

Rcv

Sched

Rcv

Re-init

inst

init re-init

meas

semf

prop only

(if nothing else happens & timeout)

ctrl

alpha

sched

(if timeout, produce suboptimal

Control using any available alphas)

(if timeout, assume no meas. 

@ other nodes this cycle)

S/W State

Trigger

S/W Method

Producing

State Transition

Legend:

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. 
Figure 4-17. Formation Motion Relative to Reference 

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. 

Spacecraft Servers GPS Receivers

Environment

Client

GPS Simulator

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. 

In this campaign, we completed

autonomous flights and fully

equipped the Frog for autonomous

formation flight with the Mule. 

The Mule was successfully flown 

by the FFIS several times, and the 

controller successfully navigated 

the aircraft through a flight path

defined by waypoints, pre-defined 

positions relative to the launch

point of the plane. Unfortunately, 

a landing mishap (while the Mule 

was under the control of the

human operator) damaged the

Mule beyond repair just as the 

formation flight segment of the 

program was about to begin. 

Using program reserve funds, an 

additional Frog is currently being equipped to replace the "Mule," and we plan to complete the formation 

flying test campaign this fall. We plan to issue a supplement to this final report at the conclusion of the fall 

campaign. 

Figure 4-20. Aircraft Flight Test 

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 zerogravity 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. 
 

Left: A free-flying SPHERE Breadboard Satellite on the KC-135 Zero Gravity Trainer Aircraft; 

Right: Two of the SPHERES Satellites in formation flight on the KC-135. 

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] 

4.4 Attitude Determination and Modeling Techniques

4.4.1 Attitude and Orbit Model Support 

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 userspecified 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 76th 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] 

4.4.2 Advanced Attitude Determination and Sensor Calibration

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. 
 

X 

Z

Y G1

G2

G3

G4

γ

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

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. 

 

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] 

4.4.3 Multi-Mission Attitude Determination System

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. 

 

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. 
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 ADSMATLAB 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] 

4.5 Flight Dynamics Automation Studies

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] 
5.0 Branch Infrastructure

5.1 Flight Dynamics Tool Program

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] 

5.2 Flight Dynamics Lab

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

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 . 

[Technical contact: Sue Hoge] 

5.4 2001 FLIGHT MECHANICS SYMPOSIUM

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

6.1 GSFC Standards Program

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 agencyendorsed 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

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) PegasusXL 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 controlsstructures-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.

 

 Figure 1 Hyper-X attached to the Booster Rocket 

[Technical Contacts: Josephine San, Peiman Maghami, Jim Morrissey] 
7.0 Outreach Activities

7.1 SAMPEX University Operations

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] 

7.2 Undergraduate Student Research Program (USRP) 

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 microNewton 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 

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] 

7.4 Graduate Student Research Program (GSRP) 

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 

• “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] 

7.5 TEAMS Competition 

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. 

Figure 7-1 TEAMS Robotic Competition at Goddard 

[Technical Contact: Tom Stengle] 
7.6 Public Education/Community Outreach 

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

Team Awards 

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

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 

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] 

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] 

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] 

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] 

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] 

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] 

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

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: 

24th Annual AAS Guidance and Control Conference, Breckenridge Co, Jan. 31-Feb. 4, 2001

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

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

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

Flight Mechanics Symposium, NASA Goddard Space Flight Center, Greenbelt, MD, June 19-21, 2001

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

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

2001 American Controls Conference, Arlington, June 25-27

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.

Carpenter, J. Russell, "Distributed Spacecraft Control Architectures" (invited presentation) 
AIAA Guidance, Navigation, and Controls Conf., Montreal, Quebec, Canada, August 6-9, 2001

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

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 GammaRay 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

David Quinn, R.E. Farley, “Tethered Formation Configurations: Meeting the Scientific Objectives of Large

Aperture and Interferometric Science” 

Institute of Navigation 14th International Technical Meeting, Salt Lake City, UT, Sept. 11-14. 2001

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," 

40th IEEE Conference on Decision and Control, Orlando, FL, December 4-7, 2001

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 

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