APPLIED ENGINEERING AND TECHNOLOGY DIRECTORATE

 

Flight Dynamics Analysis Branch 

End of Fiscal Year 2002 Report 

Written by members of the Flight Dynamics Analysis Branch 

Edited by D. Mangus, C. Mendelsohn, S. Starin, T. Stengle, S. Truong 

 

National Aeronautics and 

Space Administration 

Goddard Space Flight Center 

Flight Dynamics Analysis Branch/Code 572 

Greenbelt, Maryland 20771

December 2002 

 

Flight Dynamics Analysis Branch 

End of Fiscal Year 2002 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 Division (GNCD) Flight Dynamics Analysis Branch (FDAB) at NASA/Goddard 

Space Flight Center (GSFC). 

The primary editors of this report were: 

David Mangus 

Chad Mendelsohn 

Scott Starin 

Tom Stengle 

Son Truong 

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

management: 

Thomas Stengle, Head, Flight Dynamics Analysis Branch, Code 572 (Thomas.H.Stengle@nasa.gov) 

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://fdab.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) 2002. 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 including navigation, spacecraft trajectory design, attitude 

analysis, attitude determination and attitude 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 Aquarius 

2.1.2 Earth Observing System (EOS) Aura 

2.1.3 Global Precipitation Measurement (GPM) Mission 

2.1.4 InFocus 

2.1.5 Solar Dynamics Observatory (SDO) 

2.1.6 Space Technology 5 (ST-5) 

2.1.7 Space Technology 7 (ST-7) Disturbance Rejection System (DRS) 

2.1.8 Triana 

2.2 Operational Missions 

2.2.1 Earth Observing System (EOS) Aqua 

2.2.2 EOS Terra 

2.2.3 Earth Radiation Budget Experiment (ERBS) Decommissioning 

2.2.4 The Far Ultraviolet Spectroscopic Explorer (FUSE) 

2.2.5 Geostationary Operational Environmental Satellite – N (GOES-N) 

2.2.6 Landsat-5: End of Life? 

2.2.7 Meteor/SAGE 

2.2.8 Microwave Anisotropy Probe (MAP) 

2.2.9 Rossi X-ray Timing Explorer (RXTE) 

2.2.10 TDRS-I Support 

2.2.11 Tropical Rainfall Mapping Mission (TRMM) 

2.2.12 Mission Services Program Office (MSPO) 

3.0 Study Mission Support 

3.1 Constellation X 

3.2 Geospace Electrodynamic Connections (GEC) 

3.3 James Webb Space Telescope (JWST) 

3.4 Leonardo 

3.5 Laser Interferometer Space Antenna (LISA) 

3.6 Lunar Science Explorer 

3.7 Magnetospheric Multiscale Mission (MMS) 

3.8 Themis 

3.9 Venus Sounder for Planetary Exploration (VESPER) Discovery Proposal 

3.10 General Orbit Determination Error Analysis 

3.11 Integrated Mission Design Center (IMDC) 

4.0 Technology Development Activities 

4.1 Advanced Mission Design 

4.1.1 Optimization 

4.1.2 Unique Orbits 

4.1.3 Integrators 
4.2 Autonomous Navigation Technologies 

4.2.1 Precise Relative Navigation 

4.2.2 Celestial Navigation 

4.2.3 High Altitude GPS Navigation 

4.2.4 Magnetometer-based Navigation 

4.3 Formation Flying Technologies 

4.3.1 Autonomous Formation Flying Control of EO-1 

4.3.2 Tethered Formation Flying 

4.3.3 6DOF Nonlinear Control for Formation Flying 

4.3.4 Maxim Pathfinder Testbed 

4.3.5 Decentralized Estimation and Control of Distributed Spacecraft 

4.4 Attitude Determination Technology 

4.4.1 PiVoT Attitude Determination 

4.4.2 GPS Attitude Sensor 

4.4.3 Advanced Attitude Determination Methods 

4.5 Flight Dynamics Automation Studies 

4.5.1 Flight Dynamics Automation with The University of Maryland 

4.5.2 TRMM Reengineering 

4.6 Nanosat Technology 

4.7 GMSEC Architecture Development 

5.0 Branch Infrastructure 

5.1 Flight Dynamics Tool Program 

5.2 Guidance, Navigation and Control Lab 

5.3 Flight Dynamics Models 

 5.3.1 SKYMAP Maintenance 

 5.3.2 Schatten Predicts 

 5.4 FDAB Website 

6.0 Interagency Activities

6.1 X43-A Failure Investigation 

6.2 SCISAT 

6.3 NASA Technical Standards Program 

7.0 Employee Development 

7.1 New Employee Profiles 

7.2 Professional Intern Program 

7.3 Cooperative Education Program 

7.4 Professional Development Program 

7.5 In-house Employee Classes 

7.6 Attitude Control System Handbook 

7.7 TableSat 

8.0 Outreach Activities 

 8.1 SAMPEX University Operations 

 8.2 PREST Program 

 8.3 Graduate Student Research Program (GSRP) 

 8.4 Mission Design and Navigation Workshop 
Appendix A – Goddard and NASA Awards 

Appendix B - University Grants 

Appendix C – Conferences and Papers 

Appendix D- Acronyms and Abbreviations 
List of Figures and Tables

Figure 2-1. The Aquarius Spacecraft 

Figure 2-2. The Aura Spacecraft 

Figure 2-3. Aura Ground Track 

Figure 2-4. The GPM Spacecraft 

Figure 2-5. InFocus Gondola and Balloon Launch 

Figure 2-6. The ST5 Constellation 

Figure 2-7. ST-5 Spacecraft Separation 

Figure 2-8. ST-5 Nutation Damper 

Figure 2-9. Conceptual Diagram of the ST-7 Disturbance Reduction System 

Figure 2-10. Terra DSC Timeline 

Figure 2-11. Terra DSC Geometry 

Figure 2-12. Landsat-5 Image With “Caterpillar Tracks” Present 

Figure 2-13. Trajectory From SK3 Balanced With SK4 Performed on November 5 2002 

Figure 2-14. MAP System Momentum History 

Figure 2-15. Rossi X-ray Timing Explorer (RXTE) 

Figure 2-16. TRMM Attitude Error 

Figure 2-17. TRMM Ephemeris Error 

Figure 3-1. Constellation X 

Figure 3-2. GEC Spacecraft Formation 

Figure 3-3. JWST L2 Trajectory 

Figure 3-4. Laser Interferometer Space Antenna (LISA) 

Figure 3-5. LSE Lunar Capture 

Figure 4-1. Primer Vector Analysis Tool (PVAT) 

Figure 4-2. Heteroclinic Connection 

Figure 4-3. Dynamical Systems Utility – Generator 

Figure 4-4. Quasi-stationary Locations 

Figure 4-5. Autonomous Navigation Supported by GEONS 

Figure 4-6. AutoCon Control Flow 

Figure 4-7. Tethered Formation Flying Satellite Dynamics Tool (SDT) – Satellite Tool 

 Kit (STK) Connection 

Figure 4-8. True and Measured Sun Angles for a Nutating Spacecraft 

Figure 5-1. Predicted and Actual Solar Activity for 2002, 2002 

Figure 6-1. X-43 Air-Breathing Hypersonic Aircraft 

Figure 7-1. Momentum Wheel Sizing Utility 

Figure 7-2. TableSat Side View 

Figure 7-3. TableSat Top View 

Table 2-1. Planned versus actual results for the three stationkeeping burns. 
U1.0 Introduction

This is the fourth annual report produced by members of the Flight Dynamics Analysis 

Branch (FDAB) at the Goddard Space Flight Center (GSFC). As part of the Guidance, 

Navigation and Control Division (GNCD), the branch is responsible for providing 

analytic expertise for trajectory and attitude systems. This includes dynamics and control 

analyses and simulations of space vehicles. The Branch creates and maintains state-ofthe-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 spacecraft mission 

design, on-orbit sensor calibration, and launch/early orbit operations. An active 

technology development program is maintained, 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 follows an outline similar to one used in past annual reports. It 

summarizes the major activities and accomplishments performed by the FDAB in support 

of flight projects and technology development initiatives in Fiscal Year (FY) 2002. 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 engineers 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 FY2002 are: 

• Successful branch support for the launch and early operations of the Aqua 

spacecraft. As part of this support, the branch designed and implemented orbit 

raising maneuvers.

• Successful insertion of the MAP spacecraft into its operational orbit about L2. 

Branch engineers continued their support of the MAP mission that was launched in 

July of 2001. This supported included attitude control system monitoring and orbit 

maneuver computations.

• In-house development efforts begin for the Global Precipitation Measurement 

(GPM) mission and Solar Dynamic Observatory (SDO). With these two major in-
house spacecraft development projects, the branch will play a key role in the 

development of flight and ground orbit and attitude systems.

• Space Technology (ST)-5 completes Critical Design Review. FDAB engineers 

successfully completed this milestone with presentation of the final design for the ST5 attitude control system and design of the orbit constellation.

• Patent application submitted for the “User Involved Star Identification and Attitude 

Estimation Technique.” This technique for star identification was implemented in 

institutional branch attitude estimation software used for current and future 

operational missions.

• Goddard proposal for the ST-7 Disturbance Reduction System (DRS) selected. 

Branch members will have a key role in the development of algorithms for this 

technology pathfinder.

• TRMM spacecraft lifetime extended. Branch engineers continued active support of 

TRMM operations and analysis. The TRMM lifetime was extended 2-5 years as a 

result of a boosting of the orbit from 350 km to 400 km.

• Autocon selected as Goddard software of the year. This software was also runner-up 

in the NASA-wide competition. Used to automate orbit maneuvering for the EO-1 

spacecraft, this software will be used for operational support of numerous future 

missions.
2.0 Flight Project Support 

This section summarizes FDAB support to GSFC flight projects during FY02. For 

purposes of this report, these projects are classified as: 

• Development Missions: Approved mission under development. 

• Operational Missions: Missions that were in-flight in FY02. This includes missions 

that were in the final stages of development and were successfully launced in FY02 

(e.g. AQUA). 

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 Aquarius (planned launch 2006) 

Figure 2-1. The Aquarius Spacecraft 

Aquarius is a satellite that will address NASA Earth Science Enterprise questions about 

the global cycling of water and the response of ocean circulation to climate change. 

Salinity is the only surface parameter not currently measured from space. Aquarius will 

provide this measurement by monitoring global ocean radiometric emission, which is 

influenced by surface salinity. 

The primary science objectives of the mission are to measure global sea surface salinity, 

monitor freshwater cycling at the ocean surface, understand the response of ocean 

circulation to buoyancy forcing, assess the impact of buoyancy forcing on the ocean 

thermal feedback to the climate (e.g., El Niño prediction), and improve the ability to 

estimate the air-sea exchange of CO2. The required mission duration is 3 years, with a 

goal of 5 years. 
Over the past few years, the FDAB has provided a variety of analysis support for the 

Aquarius proposal team. The support includes assisting in the tasks of orbit selection, 

launch vehicle selection, devising an orbit maintenance strategy, surface coverage 

analysis, and evaluation of various sensor configurations. Global coverage will be 

achieved each week from a 600 km near-polar orbit. From this altitude, a 3-meter antenna 

will produce a 300 km wide swath with 100 km resolution. 

The Aquarius proposal team was recently informed that their mission concept has been 

selected by NASA Headquarters to proceed to the formulation phase. 

[Technical contact: Frank Vaughn] 

2.1.2 EOS Aura (planned launch 2004) 

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

Figure 2-2. The Aura Spacecraft 

The Aura mission is planned for launch in January 2004 on a Delta 7920 rocket from the 

Western Test Range. The planned mission lifetime is six years. The Aura mission is 

composed of four complementary instruments: the High Resolution Dynamic Limb

Sounder (HIRDLS), the Microwave Limb Sounder (MLS), the Ozone Monitoring 

Instrument (OMI), and the Tropospheric Emission Spectrometer (TES). 

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 (Figure 2-2) is 3-axis stabilized and will operate in a 

near-circular, sun-synchronous polar orbit at an altitude of approximately 705 km, with 

ascending nodal crossings at approximately 1:45 PM mean local solar time. Aura will fly 

in constellation with EOS Aqua (launched in May 2002) on an adjacent World Reference 

System (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. Aura will follow Aqua with an 

along-track separation between 15 and 22 minutes. Figure 2-3 shows a schematic of how 

this is accomplished. In this example, Aqua is in an orbit with an ascending node mean 

local time (MLT) of 13:30; Aura will be in an orbit with MLT between 13:38 and 13:45. 

Figure 2-3. Aura Ground Track 

During FY2002, the FDAB completed a preliminary ascent maneuver plan, stationkeeping maneuver analysis, and constellation flying analysis including Aura, Aqua and 

other spacecraft. The FDAB conducted a peer review of the new COTS orbit 

determination system that is planned for incorporation into the Flight Dynamics System 

(FDS) in the Mission Operations Center (MOC). The FDAB also presented flight 

dynamics material at the Aura Mission Operations Working Group meeting, provided 

updates to the Mission Specific Requirements Document (MSRD), refined specifications 

for products, and completed development of draft Interface Control Documents (ICDs) 

with our external interfaces. 

[Technical contacts: Lauri Newman, David Tracewell] 

2.1.3 Global Precipitation Measurement (GPM) Mission (planned launch 2007) 

(http://gpm.gsfc.nasa.gov)

The Global Precipitation Measurement (GPM) mission is a follow on to the Tropical 

Rainfall Measuring Mission (TRMM) with the objective of covering a region of the earth 

up to 65 degrees in latitude while providing sufficient measurement data for short-term

rainfall accumulations. The GPM plan for achieving this increased coverage is to launch 

a constellation of satellites, each carrying radiometers. The first of the GPM satellites,

labeled the primary spacecraft, is being built as a Goddard Space Flight Center (GSFC) 

in-house mission in partnership with the National Space Development Agency 

(NASDA), the space agency of Japan. The GPM primary spacecraft is currently 

scheduled for launch at the end of 2007 aboard an H2A-202 launch vehicle. Its mission 

orbit will be at an inclination of 65 degrees and circular at 400 km. For science taking, 

the spacecraft will be 3-axis stabilized and earth pointing. The primary spacecraft will 
carry two radiometer science instruments and a GPM Microwave Imager (GMI), and will 

serve to calibrate and verify the standards to be used for the constellation.

Figure 2-4. The GPM Spacecraft 

The Flight Dynamics Analysis Branch (FDAB) is providing support for the GPM mission 

by providing orbit analysis, for both the primary spacecraft and the constellation, and

attitude analysis in designing the attitude control system (ACS) for the primary 

spacecraft. The primary spacecraft was accepted as an in-house new start with the idea of 

introducing new technologies. These new technologies include those that the FDAB has 

provided analysis support as part of the Guidance, Navigation and Control Division 

(GN&CD). An important new technology is the autonomous generation of spacecraft 

time tagged position and velocity ephemeris using the Goddard GPS receiver, PiVoT. 

Another new technology flown on a previous mission is the autonomous generation of 

orbit maintenance commands (Autocon). Both of these new automated technologies will 

reduce ground operation costs significantly. 

The FDAB responsibility for designing the primary spacecraft ACS began early this year. 

The FDAB worked with other branches and other subsystems to do trade studies to 

identify sensors and actuators to be used onboard. One trade study recently completed 

was a comprehensive look at balancing the sizing of the reaction wheels with the 

magnetic torquers. Each of the other sensors and actuators are undergoing evaluation at 

this time to determine the best complement that will meet the spacecraft’s pointing and 

knowledge requirements for each mode of operation. New technologies being considered 

for the GPM primary spacecraft sensor and actuator complement, developed within the 

GNCD, are the Active Pixel Star tracker (APS) and a new type of reaction wheel. 

The FDAB responsibility for the design of the ACS also involves the determination of 

modes that will take the spacecraft from separation of the launch vehicle to the earth 

pointing science mode. Branch attitude analysts are currently undertaking many studies 

to determine the transition philosophy, from mode to mode, as well as the ability of

sensor and actuator configurations to meet the requirements of each mode. The most 
stringent requirements are for the science mode, which requires knowledge of attitude to 

within 0.1 degrees, per axis, 3-sigma, and the ability to point to within 0.2 degrees per 

axis, 3-sigma. A trade study of error budgets for different sensor and actuator 

complements is currently being worked to determine which is best suited to meet these 

requirements. Two other requirements of concern that will affect the design of the 

science control laws are the necessity to do 180 degree yaw maneuvers about every 

month and the requirement to maintain science pointing requirements during all orbit 

maintenance maneuvers. Other analysis being conducted now are the design of the 

control algorithms for each mode, to be implemented in a high fidelity simulator and 

validated before becoming part of the onboard flight software. 

Figure 2-4 shows the current configuration of the GPM primary spacecraft. 

[Technical contacts: Joseph Garrick, Chad Mendelsohn]

2.1.4 InFocus (planned flight mid-2003) 

 

Figure 2-5. InFocus Gondola and Balloon Launch 

The Flight Dynamics Analysis Branch has, and continues to provide support in analysis, 

development, and operation of pointing systems used by the GSFC-developed InFocus Xray Telescope. The nine-meter instrument is intended to be regularly flown on 
stratospheric balloons at altitudes of 40 km. The pointing accuracy goals of the 

telescope range from arcminutes to an eventual goal of arcseconds. Even at stable 

stratospheric float altitudes, disturbances faced by the control system far exceed those 

experienced by a telescope in orbit. These disturbances are transmitted to a gondola via 

its 250 feet long load train, (the cable and parachute structure that connects it to the 

balloon). Much of the analysis work in the FDAB is focused on the load train dynamics 

and its effects on pointing performance. It is intended that knowledge gained in the 

development of this InFocus be shared with the entire scientific balloon community. 

Two balloon flights, in August 2000 and July 2001, have been completed for the InFocus 

Project. The first flight, a detector test flight, carried attitude sensors on the gondola and 

load train but had no pointing control. These included a high accuracy gyro and several 

magnetometers on the gondola and along the load train. The second flight carried the 

completed telescope and had an azimuth/elevation control as well as gyros and 

magnetometers (in the gondola only). It also carried a star camera and the Pivot GPS on 

the telescope receiver that provided enhanced position, velocity and time information. 

Pivot is under development by the GNCD. 

In the past year, branch efforts have been underway to analyze the data of both flights to 

characterize the nature of the load train disturbances and formulate improvements for the 

next flight expected either in May or September 2003. 

These improvements are needed because of inadequate performance on the second flight. 

High winds at float altitude shortened the flight duration to several hours and wind gusts 

acting on the balloon excited the load train and gondola, leading to highly stochastic 

swinging motion that exceeded the capacity of the elevation control loop. Elevation and 

cross elevation motion also coupled into the azimuth control loop. Overall pointing errors 

often reached 20 arc minutes. The stochastic gyro rate signals implied that acceptable 

pointing for a telescope the size of InFocus, may not be feasible unless the wind speed is 

reduced. Future flight dates may be restricted to spring and fall periods of low-wind 

“turnaround”. 

In addition, examination of the post flight data showed that low frequency oscillations 

were the largest contribution to pointing error and their mitigation would substantially 

improve the pointing. These low frequency modes are usually associated with the entire 

flight configuration, i.e., balloon/load train/gondola and have characteristic periods 

lasting as long as one minute. High fidelity simulations were prepared in the FDAB 

using software capable of formulating equations of motion for complex systems, i.e., 

AUTOLEV. The resulting model was substituted into a Matlab Simulink control loop in 

order to devise solutions to the long period modes. 

Even in calm conditions, swinging motions generated by control motions have plagued 

many balloon flights. Solutions to the problem have ranged from attempts to dampen the 

swing behavior to isolating the telescope instrument using gimbals etc., from the gondola 

base motion. One proposed approach that has been worked by the InFocus science team 

is to replace the azimuth/elevation control with a small spherical ball rigidly attached to 
the telescope and floated on an oil-filled cup attached rigidly to the gondola base. 

Analysis by the FDAB showed this to be a good solution for a later flight, but not 

practical due to budget and schedule constraints for the next flight. For this flight, a less 

expensive alternative was sought. The proposed solution is to actively keep the gondola 

level, even if there is swinging from the balloon. 

Approaches using large momentum wheels and shifting weights were studied. The 

results of simulations concluded that the shifting weight approach was better. Here, the 

horizontal location of two masses is varied to keep the center of mass aligned with the 

force from the load train so that its torque on the gondola is kept near zero, even though 

the gondola is translating through a swing motion. In this scheme, stepper motors move 

60 kg weights along two horizontal tracks in proper phase with the elevation and cross 

elevation motion. Efforts are continuing to establish the balance-weight loop parameters. 

This scheme is dependent on an accurate knowledge of the tilt of the gondola base. 

Because the flight software is being recomposed in XML- based software, a capability is 

forthcoming that will allow the use of Kalman filtering to estimate the tilt and other 

pertinent variables. 

Also, the data suggests that there is a correlation of swinging excitations with altitude 

changes that occur with the occasional dropping of ballast. This could indicate a 

nonlinear effect for further study. 

[Technical Contact: Dave Olney] 

2.1.5 Solar Dynamics Observatory (SDO) (planned launch 2008) 

http://lws.gsfc.nasa.gov/sdo.htm

The goal of the SDO mission is to observe the Sun's dynamics to increase understanding 

of the nature and sources of solar variability. The focus areas for the observations are:

 * Changes in the solar magnetic field 

 * Relationships between the Sun's magnetic field and mass and energy releases 

This past year has seen the start of the Solar Dyamics Observatory (SDO) mission. Initial 

design and concepts were put forward, and in the summer of 2002, an initial spacecraft 

concept was presented. 

Flight Dynamics Analysis 

As we begin to understand the requirements for developing the ground system and 

operational concepts for the Solar Dynamics Observatory, there are numerous mission 

analysis items that must be studied. The Flight Dynamics Analysis Branch at GSFC has 

been working many Flight Dynamics related analysis items since the spring of 2002. 

These include: 
• A trade study to determine whether the SDO propulsion system should 

include a solid Apogee Kick Motor (AKM) and an auxiliary hydrazine system 

or consist of a restartable Liquid AKM fueled by a bipropellant propulsion 

system. 

• The Launch and Early Orbit (L & EO) mission profile for the bipropellant 

system (the system of choice in the trade study) was generated including orbit 

maneuver placement, ground station and TDRS coverage and a timeline for 

Flight Dynamics operations. 

• The orbital profile that permitted analysis of the amount of radiation that SDO 

would experience in the L & EO phase was computed. 

• Mission orbit studies have also been worked in these areas: East-West station 

keeping, solar Radio Frequency Interference (RFI) at ground stations where 

science data will be dumped, and a right ascension of the ascending node trade 

study to help minimize orbital eclipse times and optimize High Gain Antenna 

gimbal angle operations. 

• To aid in the gimbal angle studies, the FDAB worked with the SDO 

mechanical design engineers and the engineers from Analytical Graphics, Inc. 

to develop a computer graphics model of SDO. With this model, we were 

able, using AGI’s Satellite Tool Kit software, to determine the time history 

when the SDO HGAs were obscured by a portion of the spacecraft from 

“seeing” a desired ground station. 

• The amount of ground station tracking data that would be required for orbit 

determination in the mission orbit was analyzed using orbit determination 

error analysis. 

Attitude Analysis 

In August of 2002, the instrument suite was selected, and an SDO Project Kickoff 

Meeting was held the first week in September. The GNC attitude team has started 

putting together low fidelity simulations so we can start designing and modeling the 

control modes. Spacecraft jitter has been identified as one of the control system’s tall 

poles, so we are working with the scientists to firm up the requirement, so we understand 

what we have to do. Component sizing and accuracy requirements are being worked. 

This includes reaction wheel momentum and torque capability, as well as thruster size 

and placement. In addition, nutation studies and analysis are ongoing, as this may be a 

large design driver, especially with a spinning upper stage to get to Geostationary Tansfer 

Orbit (GTO). Discussions have already occurred with the flight software group (Code 

582) to scope the automatic code generation and automatic algorithm documentation 

tasks, as well as the amount and type of flight software testing that will be required. 

The above analyses comprise only a small beginning to the work that FDAB will perform 

in aiding the design, testing, and operations of SDO. 

[Technical contacts: Stephen Andrews, Bob DeFazio] 
2.1.6 Space Technology 5 (ST-5) (planned launch 2004) 

http://nmp.jpl.nasa.gov/st5/

Figure 2-6. The ST5 Constellation 

Space Technology 5 (ST-5) 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 

shown in Figure 2-6 will validate new technology for spaceflight and demonstrate science 

grade data acquisition from a high precision three-axis magnetometer. The new 

technologies include a miniature cold gas thruster, x-band transponder, lithium ion 

battery, variable-emissivity coatings, 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 ST-5 mission design team is designing a maneuver sequence that will validate the 

miniature cold gas thruster and allow for science grade magnetic field measurements. 

Over the last year the cross link antenna originally proposed to fly on the spacecraft has 

been eliminated. As a result, the proximity requirement of 1000 km at apogee has been 

lifted. The new requirement is to achieve a 0.5 hour Mean Local Time (MLT) separation 

between spacecraft in a string of pearls configuration no more than 45 days after 

deployment. The required configuration is shown in Figure 2-7. Limited fuel capacity 

dictates that the deployment of the three spacecraft must be planned as efficiently as 

possible to meet this new requirement. Furthermore, as a launch vehicle selection for ST5 continues and the launch parameters remain unknown, designing a constellation scheme

robust to unknown orbital parameters and compliant to the spacecraft operating 

constraints is crucial for mission success. The maneuver plan incorporates an orbit 

phasing scheme that takes advantage of small variations in the orbit periods between the 

spacecraft to place the spacecraft in the required configuration. The orbit phasing scheme

will require that the leading and trailing spacecraft perform a 1 m/s maneuver in opposite 

directions along their directions of motion. 
Figure 2-7. ST-5 Spacecraft Separation 

The onboard ACS hardware consists of a sun sensor mounted perpendicular to the spin 

axis, the three-axis magnetometer, and the cold gas thruster. The challenge was to 

provide an ACS that uses simple algorithms to minimize onboard processing and work 

with the limited sensor and actuator complement. Algorithms developed for the ST-5 

ACS use Rhumb line precession to keep the spacecraft spin axis aligned with the ecliptic 

pole and to reorient the spin axis when orbit adjust maneuvers are required. These 

algorithms have been tested using a hi-fidelity simulation that models the spacecraft 

orbital and rotational dynamics, sensors, actuators, and space environment. Passive 

nutation damping will be achieved using the fluid-filled ring damper shown in Figure 2-8. 

In the past, fluid-filled ring dampers have required a fluid reservoir to reduce the internal 

pressure of the damper, which increases as the fluid expands with increasing temperature. 

The small size of the ST-5 spacecraft has allowed a smaller fluid-filled ring damper to be 

designed that does not require a fluid reservoir and functions at internal pressures up to 

10,000 psi. This provides simpler design, integration, testing and lower weight. The 

damper will be tested at NASA-GSFC by mounting the damper on the platform of a 

torsional pendulum and observing the damping of the platform rotational motion with 

rates telemetered via an RF link. 

Figure 2-8. ST-5 Nutation Damper 
As an experimental phase of constellation control, the ST-5 GN&C team is proposing to 

control the relative position of each ST-5 spacecraft thru actively controlling the ballistic 

properties of each spacecraft thru perigee. Small changes in each spacecraft’s attitude 

will result in variations in the cross sectional area, thus the drag profile, experienced 

through atmospheric passes. These differences can be calibrated as small maneuvers to 

further tune the constellation’s relative formation. This control scheme will be attempted 

at the end of the mission; after all other mission goals have been met. 

[Technical contacts: Marco Concha, Jim Morrissey] 

2.1.7 Space Technology-7 (ST-7) Disturbance Rejection System (DRS) 

(planned launch 2006) 

Figure 2-9. Conceptual Diagram of the ST-7 Disturbance Reduction System 

Design of the ST-7 spacecraft and its experiment packages is a joint venture of ESA, JPL, 

NASA/GSFC, Stanford University, and Busek Co., Inc. The NASA/GSFC contribution 

will be dynamical modeling, controller design, and flight code generation for the ST-7 

Disturbance Rejection System in the LISA Test Package. FDAB effort on the ST-7 DRS 

in FY2002 has focused on dynamical model development and preliminary controller 

design. 

The ST-7 DRS will include two free-floating test masses that are shielded from solar 

radiation pressure. Thrusters would then be used to establish disturbance-free (usually 

called “drag-free”) motion by moving the spacecraft to center the test masses in their 

respective sensor cages. Models have been created for the disturbance noise spectra of

solar radiation pressure and test mass accelerations. Design requirements for the colloidal 

propulsion devices and appropriate thruster quantization levels have been determined. A 

7-degree-of-freedom model of the dynamics of the spacecraft and two test masses—two 

translational DOF for each body, plus rotation of the main spacecraft—has been 
developed and validated against a Simulink model. This model has both linear and 

nonlinear versions, and it includes star tracker attitude measurements and preliminary 

control algorithms for attitude control, DRS control, and Gravity Reference Sensor 

(GRS). 

ST-7 spacecraft control requires a complex design, with two main control loops. First is 

the translational controller, which controls the position of the spacecraft to establish dragfree motion the first test mass. The second controller is the spacecraft attitude control, 

which is currently designed to orient the spacecraft inertially in the low frequency band 

(DC and near DC). However, it is also designed to center the spacecraft about the second 

test mass (establish drag-free motion about the second test mass) in the ST-7 science 

measurement band (1 to 10 mHz). 

[Technical Contact: Scott Starin ] 

2.1.8 Triana 

Triana is a mission dedicated to helping scientists construct more accurate models of 

Earth's climate and obtaining a detailed understanding of how solar radiation affects 

climate. Triana is designed to study the Earth, for the first time, from the vicinity of the 

Sun-Earth L1 Lagrange Point, a vantage point 1.5 million kilometers from Earth 

Sunward, and promises to offer new insights into how our planet's climate works as an 

integrated system. Triana will use its science instruments to meet the following science 

objectives: 

• Make direct measurements of the radiant power emitted by the illuminated side of 

the Earth to increase the understanding of how much of the Sun's energy is 

absorbed in the atmosphere and thus improve the understanding of global climate. 

• Observe the vegetation canopy structure and evolution to monitor the health of the 

Earth's vegetation and measure ozone and cloud coverage to study their affect on 

the amount of UV radiation that reaches the ground. 

• Measure global aerosol optical thickness to increase our knowledge of how 

pollution, generated by both human and natural causes, affects the Earth. 

• Improve our understanding of the characteristics of the solar wind and magnetic 

field and provide an early warning system for communication satellites and 

ground based systems susceptible to solar-related disturbances. 

Triana is currently in storage, and NASA is investigating a number of launch vehicle 

options. Triana and its upper stage will nominally be deployed by a Space Shuttle 

Orbiter from low Earth orbit (LEO). From LEO the upper stage burn will increase the 

Triana velocity by approximately 3.1 kilometers per second sending Triana on a 

trajectory to the Sun-Earth L1 orbit. Approximately 6 months after the upper stage burn, 

the Triana propulsion system will be used to achieve the nominal L1 mission orbit. In the 

nominal L1 mission orbit, the Sun-Earth-Triana angle will nominally be maintained 

between 4.0 and 15.0 degrees for a period of 5 years. After mission orbit insertion at L1, 

small maneuvers will be required to maintain the Sun-Earth L1 orbit. 
The Flight Dynamics Analysis Branch of the GSFC Guidance Navigation and Control 

Center leads the Triana flight dynamics team, which consists of NASA civil servants, 

Purdue University personnel, and contractor personnel. The Triana flight dynamics team 

is responsible for the nominal trajectory design and for providing the nominal trajectory 

data to the Johnson Space Center in preparation for deployment from the Space Shuttle 

Orbiter. The Triana flight dynamics team is responsible for the development of a 

trajectory correction maneuver plan, for pre-mission analysis of the trajectory correction 

maneuvers, for pre-mission orbit determination error analysis, for determining the 

required tracking schedule for the Deep Space Network (DSN) and Universal Space 

Network (USN) stations which will support Triana, and for contingency analysis. During 

the early mission phase, the Triana flight dynamics team will be responsible for 

generating the maneuver commands, for orbit determination, and for product generation. 

The Triana flight dynamics team, with support from the Mission Applications Branch, 

added automation functionality to Purdue University's Generator software, which greatly 

reduced the time required for nominal trajectory generation. In conjunction with 

Analytical Graphics, Inc. (AGI), the Triana flight dynamics team is developing software, 

which will automate the manually intensive process of determining the trajectory 

correction maneuver fuel requirements. The Triana flight dynamics team, in concert with 

the Attitude Control System (ACS) Flight Software team, completed testing and 

verification of the final attitude control system flight software build for delivery to the 

spacecraft prior to storage and completed a preliminary delivery of the attitude 

determination ground system. Triana flight dynamics team personnel have successfully 

supported Triana Project reviews including the trajectory design peer reviews. 

[Technical contacts: Greg Marr, Steve Cooley, Rick Harman] 
2.2 Operational Missions 

2.2.1 Earth Observing System (EOS) Aqua 

http://eos-aqua.gsfc.nasa.gov/

The focus for the Aqua Project is the multidisciplinary study of the Earth's Interrelated 

Processes (atmosphere, oceans, and land surface) and their relationship to earth system 

changes. The global change research emphasized with the Aqua instrument data sets 

include: atmospheric temperature and humidity profiles, clouds, precipitation and 

radiative balance; terrestrial snow and sea ice; sea surface temperature and ocean 

productivity; soil moisture; and the improvement of numerical weather prediction. 

The Aqua spacecraft was launched May 4, 2002 on a Delta II 7920-10L rocket with 9 

strap-on solid rocket motors and a 10-foot composite fairing from the Western Test 

Range at Vandenberg. The planned mission lifetime is six years. Aqua is the lead 

spacecraft in the PM portion of the EOS AM & PM constellations. 

The Aqua Flight Dynamics Team (FDT) provided pre-mission analysis support, 

participated in the pre-mission testing/simulations, supported the spacecraft launch, and 

planned the ascent maneuvers to achieve the desired mission orbit. The FDT also 

provided post-launch attitude sensor calibration products, daily planning products, critical 

activity planning/monitoring, and Flight Operations Team (FOT) training during the 120-

day checkout period. 

The FDT planned the ascent maneuvers necessary to achieve the desired Aqua mission 

orbit while also phasing with the EOS AM constellation to minimize potential ground 

station conflicts (and possible collisions) between constituents of the two constellations. 

Although the two constellations reside in separate orbit planes, the intersection of the two 

planes occurs near the northern EOS Polar Ground Network (EPGN) sites and could have 

led to EPGN resource conflicts. 

The FDT also provided maneuver information to the Aqua Project management and Air 

Force representatives from Cheyenne Mountain to assess the potential for close 

approaches with other spacecraft (or tracked space debris) during the ascent maneuvers. 

Backup maneuvers were planned by the FDT for each ascent maneuver and these were 

also assessed for close approaches. The timing of the later maneuvers in the ascent 

scenario was very critical to achieve the correct phasing with the AM constellation. 

Backup maneuvers were usually planned within a few orbits of the prime maneuver. 

The repeatability of the thruster performance during the first four of six ascent maneuvers 

was negatively affected by flight software (FSW) changes and altered inertia parameters 

implemented by the manufacturer to correct unexpected thruster duty cycle behavior. 

These spacecraft changes resulted in the first four FDT maneuvers not meeting their 

targeting goals. Once the critical spacecraft parameters were updated and used for the 

last two maneuvers, the FD maneuver team achieved the final orbit within several meters. 

 The FD attitude sensor calibration team provided parameter updates for the three-axis 

magnetometers (TAMs), gyros, and star trackers. The initial FDT star tracker calibration 
showed a very large offset in the boresight separation (~0.2 degrees), much larger than 

that expected due to launch shock (typically <0.02 degrees). This apparent boresight 

offset was too large for the onboard star identification algorithm to work, so Fine Point 

Mode (FPM) transition was delayed. The large discrepancy in the star tracker boresight 

positions was finally traced to a sign error in a star tracker alignment matrix in the FSW 

and the simulator. FPM had already been achieved using the first FDT star tracker 

calibration (~0.02 degrees offset from nominal), but a new calibration update was 

required once the sign error was corrected in the FSW. The new computed offset from 

nominal was ~0.01 degree (~40 arcseconds). A large bias was discovered in TAM #1 by 

the FD calibration team. A coarse, early mission TAM-1 bias update was provided to 

improve attitude control in modes that use the TAM. Fine TAM calibrations were done 

later in the 120-day checkout phase. An error in the onboard magnetic field calculation 

was also identified by FD calibration team and later corrected in the FSW. Although it 

was only a one-character error in the FSW code, the resulting error in the onboard 

reference field and associated attitude was quite significant. Earth sensor calibration 

parameters were not uplinked since the noise on the sensor was greater than the potential 

corrections. 

The FDT provided significant assistance to the FOT when investigating several 

anomalous transitions from FPM to one of the spacecraft safe modes—Earth Point Mode 

(EPM) or Sun Point Mode (SPM). The FDT provided specific analytical results, FD 

product interpretation, and technical support for numerous meetings. The FDT also 

provided daily planning products, routine attitude support, and trained the FOT in the 

operation of the FD System (FDS) for the duration of the mission. The transition to Aqua 

FOT routine operation support was completed on 9/30/02. 

 [Technical contact: David Tracewell] 

2.2.2 Earth Observing System (EOS) Terra 

http://terra.nasa.gov/

Terra has been on orbit since December 1999 and is operating nominally. In FY2002, the 

FDAB supported the Earth Science Mission Operations (ESMO) Project in planning for a 

Terra Deep Space Calibration (DSC) attitude maneuver. The purpose of the DSC is to 

provide the science instruments with calibration opportunities using the cold background 

of deep space and the stable lunar surface. Baseline DSC maneuver is a constant rate (Yaxis) pitch maneuver with an inertial 0.122 degree/sec rate. DSC is initiated at the subsatellite point of the Earth surface terminator, and is completed before spacecraft day, as 

shown in Figure 2-10. 
Figure 2-10 Terra DSC Timeline 

Figure 2-11 illustrates the Earth-Moon geometry for Terra DSC maneuvers, as Viewed

from the North celestial pole. For the DSC maneuvers without lunar viewing, the Moon 

must be between last quarter and first quarter phase to avoid potential impingement of 

reflected sunlight from the moon into MODIS or the other instruments. For the DSC 

maneuver with lunar viewing, the maneuver is to be executed ~2 days before full moon 

with a lunar phase angle of ~22.5 degrees (specific time depends upon exact Descending 

Node crossing time.) 

Figure 2-11 Terra DSC Geometry 
ESMO personnel presented a briefing of an integrated plan for Terra and Aqua Deep 

Space Calibration to NASA HQ on August 28, 2002. The presentation was received very 

favorably, and subsequently, Code Y asked Code Q to provide an assessment of the 

maneuver and its associated risks. To enable Code Q to comply with the Code Y request, 

the ESMO Project is responsible for implementing a review and is requesting GSFC 

Code 300 support in this endeavor. The review panel will include FDAB personnel. The 

FDAB will also be involved with the actual DSC maneuver planning, maneuver 

simulations and validation, and maneuver execution. 

[Technical Contact: Mark Woodard] 

2.2.3 Earth Radiation Budget Satellite (ERBS) Decommission 

The Earth Radiation Budget Satellite (ERBS) mission was intended to be a 2-year 

mission; eighteen years after it's 1984 launch it is still operating and producing useful 

science data. While ERBS continues to collect science data with the SAGE II and ERBE 

instruments, the technology is more than 20 years old and much more sophisticated and 

accurate instruments are available. In the summer of 2002, the NASA HQ Earth Science 

Program Office decided to decommission the ERBS satellite to ensure funding for future 

missions and more up-to-date and functional Earth science missions. The GSFC ERBS 

Mission Manager Vickie Moran was tasked to produce a decommission plan for the 

ERBS satellite which maximized the safe disposal of the satellite while meeting science 

and budget constraints. In a remarkably short amount of time, Vickie and her team, 

which included 572 branch members Frank Vaughn, Sue Hoge and Jim Morrissey, were 

able to produce a viable ERBS Decommission Plan that NASA HQ approved on June 20, 

2002. Per NASA directive, a Peer Review of the Decommission Plan was held on July 

8th; the Peer Panel included Karen Richon (chair), Dave Mangus and Greg Marr of 572. 

The Panel was impressed with the work the team had produced in such a short time and 

approved the plan. 

The primary goal of the Decommission Plan was to lower the perigee altitude as much as 

possible so that the ERBS orbit decayed more quickly, decreasing the predicted lifetime 

from 18 to 9 years (nominal predicted lifetime). A secondary but important goal was to 

complete all maneuvers by the end of Fiscal Year 2002. Hardware constraints limited the 

duration and number of maneuvers which could be performed from July 14-Sept 30 to 34 

maneuvers, leaving fuel in the tanks, so the plan was to empty the tanks after the last 

maneuver on Sept 30. 

[Technical Contact: Sue Hoge, Frank Vaughn] 

2.2.4 The Far Ultraviolet Spectroscopic Explorer (FUSE) 

http://fuse.pha.jhu.edu/overview/mission_ov.html

FUSE is part of the NASA Origins Program. The purpose of the Origins Program is to 

answer two fundamental questions: Where do we come from? Are we alone? FUSE gives 
astronomers the unique capability of observing the universe’s far ultraviolet portion of 

the electromagnetic spectrum (approximately 90 to 120 nanometers). Studying this light, 

astronomers are able to better understand the conditions just after the big bang, as well as 

the chemical evolution of galaxies and interstellar gas clouds. 

In the fall of 2001, the FUSE spacecraft lost two reaction wheels. In May 2001, one gyro 

failed. Based on intensity warnings, it is anticipated that all of the remaining gyros will 

also fail. With the mission in jeopardy, the Johns Hopkins University Applied Physics 

Laboratory (JHUAPL) and Orbital Sciences Corporation (OSC) requested the FDAB to 

review the recovery procedures, attitude determination methods, and control system 

designs. Working closely with OSC, the FDAB studied various attitude and rate 

determination methods to determine the best fit for the mission in the event of gyroless 

operations. The FDAB also developed a simple safe-hold design that will maintain a 

power-positive attitude in the event that attitude determination and all of the gyros are 

lost. 

Two attitude and rate determination approaches were tested. One relies on a kinematics 

model for propagation—a method used in aircraft tracking—and the other is a traditional 

extended Kalman filter (EKF) that utilizes Euler’s equations in the propagation of the 

estimated rate. Both methods compare the measured geomagnetic field with an onboard 

model of the expected magnetic field to update the attitude and rate estimates. No other 

attitude sensors are used. The traditional EKF was selected by OSC for gyroless 

operations, primarily since the flight software is currently running an EKF. The attitude 

and rate estimates must be within 2 degrees and 20 arcsec/sec, respectively, for the star 

sensor to locate target stars for science operations. Whether or not the system will remain 

stable in a closed-loop sense during slew maneuvers is still under investigation. 

The new safe-hold algorithm is required to point the solar arrays at the Sun during the 

daylight portion of the orbit and hold the instrument out of the orbit plane without the use 

of gyros. The algorithm makes use of a physical consequence of "B-dot control," which 

simply controls according to the difference between consecutive magnetic field 

measurements. If B-dot control is applied to a body that has an internal momentum, that 

momentum will tend to precess away from the orbit plane. By controlling a wheel with 

its axis parallel to the instrument to hold the wheel at near constant speed (providing 

internal momentum), the wheel and instrument are made to precess away from the orbit 

plane. The wheel is then slightly modulated to maintain Sun pointing. 

During this development effort, the FUSE spacecraft has continued to perform worldclass science. Some of the discoveries have been a galactic corona that is much larger 

than expected and the discovery of hydrogen on Mars. The “Science Summaries” section 

at the web site has a list of highlights. 

[Technical contacts: Dave Mangus, Julie Thienel, Rick Harman] 
2.2.5 Geostationary Operational Environmental Satellite – N (GOES-N) 

In a program stretching back for more than twenty-five years, GOES-N will be the first 

satellite in the fourth series of NOAA geostationary equatorial weather satellites. Built 

by Boeing Satellite Systems (BSS), El Segundo, CA, this spacecraft is scheduled for 

launch in the first quarter of calendar 2004. Unlike the earlier series, the Flight 

Dynamics analysis and operations are not under the direct control of Goddard’s Flight 

Dynamics Analysis Branch (FDAB) and its predecessors. Instead, the FDAB is serving 

as a consultant to the GSFC GOES Project in reviewing BSS Flight Dynamics mission 

analysis and operational planning. The GOES-(N-Q) contract specified that these 

spacecraft be delivered to their checkout longitude with BSS performing and responsible 

for all Launch and Early Orbit (L&EO) operations. This is still the case; however, a 

recent proposal by BSS would include GSFC Flight Dynamics personnel on the BSS 

Flight Dynamics team for L&EO. Training of these GSFC personnel will begin in early 

spring 2003. 

GOES-N will be launched on a Delta-III rocket from Kennedy Space Center into a very 

eccentric geosynchronous transfer orbit (GTO) having an apogee radius near 85,000 

kilometers and an inclination of 28.5 degrees. Using a restartable Liquid Apogee Motor 

(LAM), the orbit will be placed using approximately 6 orbit maneuvers into a 

geosynchronous equatorial orbit with an inclination below 0.5 degrees. While 

maneuvering toward geosynchronous orbit, the spacecraft will also be phased in 

longitude by specific timing of these burns. The GOES Program places its operational 

satellites at the longitude locations of 75.0 degrees and 135.0 degrees West longitude. If 

fully functional satellites are already in place at those locations, GOES-N may be kept in 

a storage condition near 105.0 degrees West longitude. 

Flight Dynamics personnel play a critical role in placing the spacecraft at its desired 

location during geosynchronous L&EO operations. A carefully constructed L&EO flight 

operations script depends heavily on the timing of events and products that are computed 

by Flight Dynamics personnel. Although the role of the GSFC Flight Dynamics Team is 

different from that of past missions, we are still ready to provide our best efforts and 

experience to make GOES-N a success. 

[Technical contacts: R. DeFazio] 

2.2.6 Landsat-5: End of Life? 

Not so fast! It is true Landsat-5 suffered a loss of its primary imaging mode, but the 

mission is far from over. Near the end of 2000, the Landsat-5 instrument shutter began to 

lose synchronization with its scan mirror so that the shutter would periodically obscure 

the Earth over part of the image area, resulting in what has come to be known as 

“caterpillar tracks” in some of the images. At first, the instrument was used sparingly to 

save it for specific images, but eventually the effect came to manifest itself in more and 

more of the images. This eventually forced consideration of either a more effective means 
of dealing with the problem or declaring the mission over and commencing de-orbit 

procedures on what was an otherwise healthy spacecraft. Since we aided USGS (who 

currently runs Landsat-5) in the de-orbit of its sister spacecraft, Landsat-4, Code 572 was 

asked to be present at several discussions regarding the option of de-orbiting Landsat-5. 

Meanwhile, Flight Operation engineers devised a brilliant scheme to continue the mission 

by controlling the scan time to a nominal value that is within the control envelope of the 

calibration shutter rather than the nominal Scan Angle Monitor (SAM) mode, in which 

the calibration shutter passes in front of the scan mirror during the mirror’s non-imaging 

period. This so called “Bumper Mode” comes at the price of losing the precision 

knowledge of the mirror position (as would normally be reported in the SAM mode) but 

this deficiency can be accommodated on the ground; as a result, Landsat-5 may be 

expected to provide useful data for years to come. The fix to this problem has permitted 

Landsat-5 the continued ability to outlast its designed mission life—launch was in March 

of 1984, and the mission was to be five years long—and has long term implications for 

Landsat-7, whose Enhanced Thematic Mapper can be expected to suffer the same

problem as it ages. 

Code 572 is charged with creating and maintaining contingency de-orbiting plans, which 

would be engaged should the spacecraft mission ever be declared over for any reason. 

Since that now appears to be pushed off until further notice, we have run a long-term

reentry analysis assuming an end-of-life declaration every year for the next five years and 

remain ready to respond should we get the call. 

Figure 2-12. Landsat-5 Image With “Caterpillar Tracks” Present 
[Technical contacts: David Quinn, Bob DeFazio] 

2.2.7 Meteor/SAGE 

The SAGE III/Meteor-3M satellite mission is a joint partnership between NASA and the 

Russian Aviation and Space Agency (RASA). It was initiated by the Gore-Chernomyrdin 

Commission in 1994 and extends a long-term working relationship between the United 

States and Russia to understand Earth's environment. 

SAGE III was successfully launched onboard a Meteor-3M spacecraft on December 10, 

2001 at 17:18:57 UTC from the Baikonur Cosmodrome in Kazakhstan. The satellite is in 

a sun-synchronous orbit with an ascending node time of about 9 AM. 

The primary navigation system for SAGE III was a GPS/GLONASS receiver. Shortly 

after launch, it was determined that problems with the receiver could not be rectified and 

an alternative navigation source was needed. FDAB met with project personnel to discuss 

alternatives. It was decided to use ground laser ranging for orbital tracking since no twoway transponder was onboard the spacecraft. Code 926 was asked to provide orbit 

determination based upon the laser ranging data. SAGE III has continued with its 

successful mission. 

[Technical contact: Mark Beckman] 

2.2.8 Microwave Anisotropy Probe (MAP) 

http://map.gsfc.nasa.gov

FY 2002 has been another productive and busy year for the MAP team. MAP was 

successfully launched on June 30, 2001 at 19:46:46 Z from the Eastern Range at Cape 

Canaveral, FL, aboard a Delta II 7425 expendable launch vehicle. MAP is currently at its 

mission orbit, a Lissajous orbit about the L2 Sun-Earth Lagrange point, which is about 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 is being maintained such that the MAP-Earth vector remains between 0.5° and 

10.5° off the Sun-Earth vector to satisfy communications requirements while avoiding 

eclipses. The MAP mission lifetime is 2 years with a goal of 4 years, which will make 

MAP the first spacecraft to orbit about the L2 Lagrange point for up to 2 years. The MAP 

satellite arrived at L2 on January 2, 2002. The maneuver team has successfully 

completed a number of important milestones during this year. 
Figure 2-13. Trajectory From SK3 Balanced With SK4 Performed on 

November 5, 2002 

MAP support during FY02 has mainly consisted of planning and supporting maneuvers 

to keep MAP in its nominal Lissajous orbit. To date, there have been three such 

stationkeeping maneuvers (named SK1, SK2, and SK3). The trajectory from SK3, 

balanced with an SK4 performed on November 5, 2002, can be seen in Figure 2-13. For 

each of the maneuvers the trajectory team was responsible for planning each maneuver 

and determining what would happen in the case of certain contingencies. The planning 

process involves several critical steps. First, the team determines the direction and 

magnitude of the burn. This is accomplished by targeting a maneuver at least three 

months downstream that will meet the requirements of being less than 1 meter per 

second. Once the magnitude of the maneuver is determined, the trajectory team

calculates the desired attitude for that maneuver. The spacecraft attitude must meet 

several criteria during the burn. First, it must place the spacecraft z-axis 19 degrees off 

the Sun line, balancing thermal, power and attitude control concerns. Second, during the 

maneuver, the spacecraft z-axis must move away from the Sun line (the motion is due to 

the torque on the spacecraft caused by the canting of thrusters 1 & 2). Third, the attitude 

is chosen such that the delta-V direction lies as close to the velocity vector as possible. 

This third constraint allows for the most efficient maneuver given the other two 

constraints. Once the attitude is determined, the finite maneuver is planned. 

For the SK3 maneuver, a new procedure was added that allows the maneuver planning 

team to more accurately match what the HiFi and FlatSat simulators predict for the burn, 

which improves the accuracy of the burn. This additional iteration consists of generating 
a maneuver plan using predicted thruster duty cycles that are typical of stationkeeping 

maneuvers. The maneuver support has been extremely successful. All three maneuvers 

have been well within requirements (see table below). The next maneuver (SK4) is 

scheduled for November 4, 2002. The trajectory team has also refined and updated the 

stationkeeping maneuver procedures and has recently started training the spacecraft 

controllers so that they can support future stationkeeping maneuvers. 

Delta-V (cm/s) Duration (sec) Mass Loss (kg) Scale Factor 

Maneuver Pred Actual Pred Actual Pred Actual Pred Actual 

SK1 42.94 43.50 73.20 72.92 0.186 0.188 0.935 0.950 

SK2 34.80 35.06 54.08 53.84 0.162 0.159 0.970 0.982 

SK3 45.96 46.60 72.48 71.84 0.213 0.211 0.970 0.983 

Table 2-1. Planned versus actual results for the three stationkeeping burns. 

In addition to the maneuvers described above, there was a Safehold incident at about 10 

pm EST on November 5, 2001. A “Hardware-Cold” reset had occurred, which originated 

in the Power On Reset (POR) circuit of the Mongoose V processor card, and was caused 

by a single event transient (SET) due to heavy ions from a very intense solar storm that 

occurred November 5 & 6, 2001. This condition was discovered at the onset of MAP's 

ground pass at 9 am Nov. 6. Both sets of attitude control electronics (ACE) were in 

correct configuration, with the ACE Safe-Hold controlling to within 1 degree of the Sun 

line. The particular configuration indicated the reason for Safe-Hold was some reset of 

the Mongoose processor, rather than a failure of any ACS component. 

Fortunately, there was no evidence of permanent damage to any components, and we 

were able to exit Safehold by 12:45 pm and reconfigure back to Observing Mode by 1:40 

pm. We watched MAP for a full precession period (one hour) and then declared a return 

to nominal operations and terminated our Spacecraft Emergency status. The whole MAP 

Team responded quickly and efficiently to the anomaly, and rapidly and safely returned 

the MAP Observatory to Observing Mode. The Flight Software, ACS, Command 

Control & Communications, Systems Engineering and Spacecraft Mission Operations 

Center control teams, and the Code 561 Radiation Group were instrumental in the 

investigation of this anomaly. MAP has operated without incident since this event. 

System momentum buildup on-orbit has been less than 0.01 newton-meter-seconds 

(Nms) per day. This means that MAP can easily go three or four months between 

stationkeeping burns with its system momentum remaining below the 3-Nms 

performance limit. The maximum system momentum on orbit was 1.45 Nms, seen just 

before the first stationkeeping maneuver, as shown in Figure 2-14. 
Figure 2-14. MAP System Momentum History 

The MAP GNCD support team wrote a significant numbers of papers during this fiscal 

year, and was presented with two team awards. The list of papers and awards may be 

found in the Appendix of this document. 

[Technical contacts: Osvaldo O. Cuevas, Stephen Andrews] 

2.2.9 Rossi X-ray Timing Explorer (RXTE) 

http://guinan.gsfc.nasa.gov/docs/xte/xte_1st.html

Since 1995, the RXTE has been observing bursts of X-rays that come from high-energy 

phenomena including black holes, neutron stars, and X-ray pulsars. The RXTE performs

multiple slew maneuvers, to point to the various ground selected targets. RXTE can dwell 

on a target with arcsecond pointing accuracy. This tight pointing can be accomplished 

using high-precision gyros and star trackers. The star tracker updates any type of drift 

that the gyros might produce by using a Kalman filter. 

On November 25, 2001, the RXTE star tracker failed to establish lock onto a particular 

set of stars. One of the stars had a variable, near-neighbor star that was not in the onboard 

catalog. At that time, the near neighbor star was bright enough to pull the spacecraft off 

its known attitude. As RXTE slewed to new targets, the star tracker algorithms were 

unable to get an attitude fix and, therefore, did not correct for any gyro drift. To recover, 

the Flight Operations Team (FOT) sent a set of commands that had previously been 

developed by the FDAB and the FOT. The FOT also attempted to use the Real Time

Attitude Determination System (RTADS) that was developed by the FDAB. Neither 

method worked, so the FOT contact the FDAB for support. 
Figure 2-15. Rossi X-ray Timing Explorer (RXTE) 

The FDAB was able to determine that the commands were initially sent while the Earth 

occulted the star tracker field of view. It was also determined that the RTADS in the FOT 

had been set to a wrong configuration. After the FDAB reconfigured the RTADS and 

worked with the FOT to determine the proper orbit geometry, the RTXE was successfully 

reset to its proper attitude. 

The lessons learned covered three areas. First, it was made clear to the FOT that attitude 

knowledge loss is not part of normal operations. The urgency within the control room

should be heightened. Second, a launch-style checklist was developed. This will verify all 

ground systems and spacecraft configurations before the recovery commands are sent. 

Third, recovery procedures must be rehearsed with the oversight of an independent test 

conductor. 

[Technical contact: Dave Mangus] 

2.2.10 TDRS-I Support 

http://tdrs.gsfc.nasa.gov/tdrsproject/

Several months before the launch of TDRS-I, concerns were raised about the satellite’s 

Hemispherical Resonator Gyro (HRG). HRGs ingest helium, and if contaminated with 

too much they can take a significant amount of time to start up. Because the TDRS-I 

HRGs were going to be powered off and then back on soon after launch—to switch from 

the high-rate mode needed at launch to the more accurate low-rate mode used for the rest 

of its operations—the concern was for the health of the spacecraft if the HRGs failed to 

restart in the nominal amount of time. The TDRS-I contractor addressed these concerns 
by running a dry nitrogen purge to slow the degradation of the HRGs before launch, and 

by identifying a power- and thermal-safe, spin-stabilized orientation that the spacecraft 

could adopt if the HRGs did not start up correctly. Flight Dynamics Analysis Branch 

analysts Jim O’Donnell and Steve Andrews were asked to verify that this orientation 

would keep TDRS-I safe without gyros during its geosynchronous transfer orbit (GTO). 

By collecting information on its mass properties and the planned spin-stabilized gyroless 

control mode, the FDAB analysts were able to quickly put together a simulation of 

TDRS-I in its GTO orbit. Using this simulation, the analysts verified that the gyroless 

control mode would allow the spacecraft to remain power-safe through a number of 

orbits, well past the worst-case HRG restart time. 

TDRS-I launched on March 8, 2002, at 5:59pm, and there were no problems with the 

restart of its HRGs. The spacecraft did suffer an on-orbit anomaly after launch related to 

its propulsion system. At the time of this writing, work continues to be done to try to get 

TDRS-I as close as possible to its planned geostationary orbit. 

[Technical contact: James O’Donnell] 

2.2.11 Tropical Rainfall Mapping Mission (TRMM) 

http://trmm.gsfc.nasa.gov/

http://trmm-fot.gsfc.nasa.gov/

TRMM has executed 69 Delta V maneuvers since a boost to a 402 km altitude orbit in 

August 2001. In addition, 14 yaw maneuvers have been performed in the past year. 

During the orbit boost, the Earth Sensor Assembly performance degraded to the point that 

it could no longer be used to control the spacecraft attitude. The Kalman filter was 

enabled during the boost phase, and has been running continuously since then. 

It was found that the Precipitation Radar (PR) data from TRMM could be used to 

measure the roll attitude error, and this data showed that TRMM was not meeting 

pointing requirements after the Kalman filter was turned on. There were several 

improvements made to the flight software and the ground system data processing that 

greatly improved the accuracy of the Kalman filter. Some of the changes were to 

onboard software, and other changes were to the ephemeris processing; a detailed 

description of the changes made may be found in the following reference: 

S. Andrews, S. Bilanow, “Recent Flight Results Of The Trmm Kalman Filter”,AIAA2002-5047, AIAA GNC Conference, Monterey, CA, August 5-8, 2002. 

The result of the changes is that the attitude error, as determined by the PR data, is now 

less than 0.2 degrees, peak-to-peak, as shown in Figure 2-16. The ephemeris errors also 

have been substantially reduced, as indicated in Figure 2-17. 
Figure 2-16. TRMM Attitude Error 

Figure 2-17. TRMM Ephemeris Error

Reentry planning got underway in earnest on TRMM this year. A TRMM Reentry 

GN&C Peer Review was held on June 11, 2002. Brent Robertson (GNCD System

Engineering Branch Head) had been coordinating and leading this effort. 

Re-entry Maneuver Planning (F. Vaughn)
A baseline maneuver sequence to de-orbit TRMM was developed. Analysis was 

performed to characterize the impact footprint based on variations in thruster 

performance and ballistic coefficient. The impact of partial or missed maneuvers on the 

maneuver sequence was assessed to develop contingency procedures. 

ACS Analysis / Simulations (J. Morrissey) 

Attitude control system analysis for the TRMM re-entry has been performed using the 

existing TRMM HiFi Simulator. Simulation runs are performed for every aspect of the 

TRMM re-entry to verify desired pointing and system stability. The simulator is also 

used to predict fuel usage for the ACS that goes into the overall fuel budget for the reentry operation. A high fidelity aerodynamic model has been added to the simulator in 

order to increase the accuracy of performance predictions at low altitudes where higher 

atmospheric density will produce large torques on the spacecraft during re-entry. Prior to 

the final burn that will re-enter TRMM the spacecraft must survive a low perigee pass 

(150 km altitude) in the presence of these high aerodynamic torques. It was found using 

the TRMM HiFi Simulator that the reaction wheels could not store enough angular 

momentum to maintain nadir pointing through this perigee pass. This problem was 

solved by modifying the thruster-based momentum unloading mode to perform as a PD 

controller that maintains nadir pointing. The modifications were added to and verified on 

the TRMM HiFi Simulator. A software patch that implements this modification will be 

uploaded to the spacecraft before re-entry operations. 

Reentry Operations Timeline (S. Andrews) 

A timeline and Absolute Time Sequence (ATS) structure was developed to perform the 

controlled re-entry. The existing ATS and timeline for stationkeeping maneuvers were 

modified to allow the spacecraft to perform the following functions: 

Delta H mode test 

Burn #1 

Burn #2 

Perigee Delta H attitude hold 

 Burn #3 

A more formal timeline and complete ATSs will be developed next. 

FDC Configurations (D. Mangus) 

Due to the criticality of the reentry maneuvers, the nominal Failure Detection and 

Correction (FDC) configuration is not adequate for this phase of the mission. The 

maneuvers have to go as planned, so failures or anomalies that do not directly affect the 

maneuvers must not be allowed to abort the burns. Detection limits and thresholds have 

been examined, as well as actions by the spacecraft and ground. The goal is a 

configuration that minimizes risk to the maneuvers and the spacecraft, while at the same 

time maximizes the chance of a successful controlled reentry. 

Risk Analysis (D. Mangus) 

In the course of the FDC analysis, a risk management approach was started. The idea is 

to look at risks to successful completion of the reentry, and to try to quantify the 
probability and criticality of those risks. Then the effort is put into managing the high 

probability or high criticality risks. Some of the risks identified are component failures, 

inadequate performance, aborted maneuver, and the spacecraft reconfigurations necessary 

to perform the reentry. 

Aron Cooper, a co-op student, was available during the summer of 2002 to do some work 

on TRMM. He was tasked with using flight data from the orbit maintenance maneuvers 

in combination with equations of motion to try to estimate how much fuel was remaining 

on the spacecraft. The goal to help verify the propulsion and flight ops teams’ fuel 

estimates, since this is critical to determining when reentry will start. Unfortunately, 

Aron found that the fuel mass estimate was extremely sensitive to other spacecraft 

parameters (such as thruster force for the maneuver), and a satisfactory fuel mass was not 

found. 

The –Y solar array drive assembly (SADA) seized during a large array slew on 

September 4, 2002. This occurred after a long period of array inaction due to solar array 

feathering (initiated to try to increase mission life by reducing the effects or aerodynamic 

drag). The SADA eventually started moving again during the slew period, but the 

deployables subsystem engineer felt it best to fix the array at a 0° rotation angle 

(feathered with respect to velocity) to prevent a lock up at some other position. This has 

led to a lot of analysis and simulation to show that the ACS and power systems can 

operate with one array fixed and one array feathered, both for nominal mission activities, 

and for the controlled reentry. Code 572 engineers have been instrumental in simulating 

the proposed configuration, and determining the software changes required to do the 

change. A peer review was held September 19, 2002, and a plan was put in place to fix 

the –Y array, and to fully operate the +Y array. This work will be the focus of the 

TRMM engineers for the next few months. 

[Technical contacts: Stephen Andrews, Jim Morrissey, Frank Vaughn, Dave Mangus] 

2.2.12 Mission Services Program Office (MSPO) 

The FDAB supported the MSPO by providing consultative support for the initial 

planning of the move of the Flight Dynamics Facility (FDF). The move of the FDF will 

consolidate all of the operations facilities that the MSPO oversees and provide for more 

efficient system and facility security. In its consulting capacity, the FDAB assisted the 

MSPO in reviewing the high level plan for the move, provided input for the new facility 

modifications and layout and reviewed schedule and budget. Since the move to the new 

facility is expected to take about 12 months, the FDAB will be continue to provide 

consulting support to MSPO for the planning, contract monitoring and equipment 

procurement related to the move for fiscal year 2003. The FDAB will also assist in the 

implementation of the FDF move plan. 

[Technical contact: Sue Hoge] 
3.0 Study Mission Support 

3.1 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. The

baseline plan is to launch the spacecraft in pairs aboard an Atlas V launch vehicle. 

This year the FDAB has provided support to the Constellation X study team in the area of 

trajectory design. The baseline trajectory is an orbit about the Sun-Earth L2 libration 

point. FDAB provided input to the baseline Reference Mission Description document, 

which is being rewritten for release next year. The input included a delta-V budget, input 

on possible orbit determination systems that could be flown onboard, and evaluation of 

the orbit determination requirements. The team also participated in discussions regarding 

the mission timeline and separation mechanism from the launch vehicle. In addition, 

analysis of cross-link navigation capabilities that could be used to support Constellation

X are ongoing as a branch technology effort. For additional information on that effort, 

please see the Celestial Navigation topic in the Autonomous Navigation Technology 

section of this report (section 4.2.2). 

Figure 3-1. Constellation X 

[Technical contact: Lauri Newman] 
3.2 Geospace Electrodynamic Connections (GEC) 

The Geospace Electrodynamic Connections (GEC) mission is a multi-spacecraft mission 

managed out of NASA’s Solar Terrestrial Probe (STP) office at the Goddard Space Flight 

Center, Code 460. Currently in the formulation phase, GEC plans to use four spacecraft 

to study the Earth’s Ionosphere-Thermosphere (IT) system. While this region has been 

studied before, the coordinated use of four spacecraft will allow the scientists to: 

1. discover the spatial and temporal scales on which magnetospheric energy input to 

the IT region occurs 

2. determine the spatial and temporal scales for the response of the IT system to this 

input of energy 

3. quantify the altitude dependence of the response 

The four GEC spacecraft will be launched on a Delta-II 2920 expendable launch vehicle 

into a 185 km x 2000 km, 83° orbit no earlier than September, 2009. After separation 

from the launch vehicle, the GEC spacecraft will initialize a “pearls on a string” 

formation with uneven inter-satellite spacing that will be varied during the course of the 

mission. . The uneven spacing of the four spacecraft will allow GEC to be able to resolve 

six different temporal and spatial scales. The spacing will vary from a goal of 10’s of 

seconds up to a quarter orbit throughout the life of the mission. GEC also plans to 

perform periodic “deep dipping” campaigns where all four spacecraft will lower their 

perigee to an altitude near 130 km. Excursions to this altitude were performed by 

Atmospheric Explorer C (AE-C) in 1975, but the multi-satellite nature of GEC will 

collect more data while also carrying instruments that AE-C didn’t (eg, electric field 

booms). 

Figure 3-2. GEC Spacecraft Formation 
During FY02, Flight Dynamics supported an industry study of GEC conducted through 

the Rapid Spacecraft Development Office (RSDO). After an open competition of RSDO 

vendors, two spacecraft contractors were selected and received funding for a 100-day

study of the GEC concept. FDAB engineers supported the studies through consultation of 

the GEC mission during the study period and by providing technical reviews at the midterm and end-term presentations. 

FDAB work is continuing in the areas of the GEC operations concept. This includes 

investigations of the initial spacecraft formation after launch vehicle separation and how 

to initiate and terminate the dipping campaigns for the entire four-spacecraft formation. 

Further work will examine the possibility of establishing a form of a repeat cycle for 

GEC to enhance the science data return. 

[Technical Contact: Michael Mesarch] 

3.3 James Webb Space Telescope (JWST) 

On September 10, 2002, NASA selected TRW, Redondo Beach, Calif., to build the Next 

Generation Space Telescope (NGST). In addition, the space-based observatory will be

known as the James Webb Space Telescope, named after James E. Webb, NASA's 

second administrator. 

The James Webb Space Telescope is scheduled for launch in 2012 aboard an expendable 

launch vehicle. It will take about three months for the spacecraft to reach its destination, 

an orbit 940,000 miles or 1.5 million kilometers in space, called the second Lagrange 

Point or L2, where the spacecraft is balanced between the gravity of the Sun and the 

Earth. See Figure 3-3 below for nominal trajectory. 

Figure 3-3. JWST L2 Trajectory 
FDAB has been chosen by TRW to provide flight dynamics support of JWST. This 

support begins with a kick-off meeting at TRW the week of November 18, 2002. FDAB 

will provide trajectory design, mission planning, and orbit determination support for 

JWST. 

FDAB also participated in a significant investigation into the design of JWST at the 

request of NASA HQ. FDAB was asked to investigate scenarios that include launch from 

the U.S. Space Shuttle and the possible servicing of JWST. FDAB came up with several 

different options for launch from STS including: 

- Use of a bi-prop engine with 8+ perigee burns to reach L2 

- Use of different ion engine configurations to reach L2 via low thrust 

FDAB also determined several scenarios in which JWST could be serviced including: 

- Launch of an separate ion tug to bring JWST back from L2 to LEO which 

would include the first ever spacecraft rendezvous in LPO 

- Use of the bi-prop engine to bring JWST back to Earth with the use of 

aerobraking to capture into LEO 

The baseline mission remains the TRW proposal of an ELV launch with no serving 

option however. 

[Technical contacts: Mark Beckman, Dave Folta] 

3.4 Leonardo 

http://climate.gsfc.nasa.gov/~wiscombe/LeoBRDF/LeoBRDFhome.html

The purpose of the Leonardo mission is to define the Bi-directional Reflectance 

Distribution Function (BRDF) of sunlight off of Earth’s clouds. According to Warren 

Wiscombe, the Leonardo Principal Investigator (PI), 

“...one of the most important [uses of the BRDF] is to reduce the 

currently large errors in estimates of radiative forcing due to air pollution, 

dust outbreaks, biomass burning fires, hurricanes, and other natural and 

anthropogenic phenomena. Radiative forcing is the lingua franca of 

global change in general, and of global warming in particular. It provides 

the common measurement scale by which phenomena as disparate as CO2

increase, cloudiness changes, and desertification can be compared.” 

Leonardo consists of a constellation of spacecraft that simultaneously measure the 

sunlight reflectance off of a cloud from different perspectives. Analysis over the past 

year focused on a constellation of six spacecraft in near-equatorial low Earth orbits. The 

spacecraft look at clouds that are above the equator and centrally located with respect to 

the constellation. Constellation maintenance occurs once per month to account for orbit 

perturbations and the Sun’s position. A genetic search algorithm was used to find 

optimum constellation formations for each month. Optimality was based directly on the 
PI’s science algorithm. Future work includes calculating minimum propellant transfers 

from one formation to the next, as well as determining the propellant required for 

constellation initialization. 

[Technical contact: Alexander Barnes, Steve Hughes] 

3.5 Laser Interferometer Space Antenna (LISA) 

The primary objective of the Laser Interferometer Space Antenna (LISA) mission is to 

detect and measure gravitational waves from massive black holes and galactic binaries in 

the frequency range between 10P

–4P

 and 0.1 Hz. The LISA mission comprises three 

identical spacecraft, 500,000 km apart, which form an equilateral triangle (Figure 3-4). 

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 toward an 

identical assembly on each of the other two spacecraft (figure 1). 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 pressure. In order to be able to detect gravitational strain levels to the 

order of 10P

–23P

, tight pointing and positioning requirements are placed on the spacecraft 

and the proof masses (e.g., acceleration requirement on each proof mass: 3x10P

-15P

 m/sP

2

P

/HzP

-

1/2P

). 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 couple of areas: (1) Dynamics and 

control modeling and analysis; and (2) Design and analysis of Disturbance Reduction 

System (DRS) control. Each of theses contributions is described in the following 

paragraphs. 

A 19-degrees-of-freedom (DOF) model of a LISA spacecraft was developed. This model 

captures the complete rigid-body dynamics of a typical LISA spacecraft in formation 

with two others. Dynamically, it includes all rigid-body degrees of freedom, six for the 

S/C translation and rotation; six for each proof mass translation and rotation; and, finally, 

one rotational DOF for telescope articulation. A number of disturbance sources, both 

internal and external, are included. Measurement noise for quad detector and capacitive 

sensing, as well as actuation noise models for the Micro-Newton Thrusters and 

electrostatic suspension controls are included. Orbital dynamics are brought in via the 
ephemeris file, which was obtained using orbit optimization. For the purpose of this 

model, the direction of the incoming beam is simulated with the aid of the orbital 

ephemeris data file for LISA, where the nominal orbital positions of the three LISA 

spacecraft are provided. Nonlinear electrostatic forces and torques, as well as those from

self-gravity, are modeled via a linear time-invariant system. Moreover, electrostatic 

actuation and sensing cross-talks are also modeled. The five control systems that 

comprise the LISA Disturbance Reduction System (DRS) are included in the model. 

These are the attitude control system (ACS), to maintain the pointing of the two 

telescopes with respect to two incoming beams from the other spacecraft; the drag-free 

control (DFC), which commands the positioning of the spacecraft to center about the 

proof masses (PM); the proof mass suspension control, to maintain the position and 

attitude of the proof mass with respect to its caging; and the telescope articulation loop, to 

maintain the optical link between the spacecraft as the angle between the spacecraft 

varies according to the natural propagation of the orbits of the spacecraft. 

Figure 3-4. Laser Interferometer Space Antenna (LISA) 

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 in the sensitive axes 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. Digital control loops 

have been designed for each of the five control systems. Two designs have been 

considered for the drag-free loop. In the first design, a centralized approach is followed 

wherein the S/C position and the proof mass translation commands are computed in a 

centralized manner to achieve drag free motion of the test masses. In the second design, 

however, the spacecraft is positioned to maintain drag-free motion in the sensitive axes, 
while the proof mass suspension control ensures that the masses follow their respective 

housings in the transverse directions. The attitude control loop uses the measurements 

from the two quad detectors, and computes appropriate S/C attitude error and telescope 

articulation commands to ensure that the optical links between the spacecraft are 

maintained to the desired accuracy. 

[Technical Contact: Peiman Maghami] 

3.6 Lunar Science Explorer 

The Lunar Science Explorer (LSE) principle investigators visited the Integrated Mission 

Design Center (IMDC) the week of March 4, 2002 and again the week of April 2. The 

LSE is a Discovery class mission designed to obtain a detailed topography map of the 

moon. The mission will map the surface of the Moon over 2 years using laser altimeters. 

The mission orbit was chosen to be circular, with a 30 km mean altitude above the Lunar 

surface. This altitude was chosen to maximize science while minimizing the fuel budget, 

or dV cost, of orbit maintenance (14 m/sec per month). The direct transfer option was 

chosen after a comparison with Weak Stability Boundary (WSB) and low-thrust options. 

The direct transfer takes 4.7 days and requires 3 dV maneuvers to capture and lower the 

altitude about the Moon (see Figure 3-5). No significant savings could be identified with 

the WSB, and the low-thrust option was too power-intensive. The total mission dV will 

be 1460 m/sec. 

Figure 3-5. LSE Lunar Capture 

Orbit determination for LSE will be quite challenging. In order to provide the science 

quality requested, orbit determination accuracy to 1 meter (radial) is necessary. However, 
at 30 km altitude, the uncertainty in the lunar potential model gives a radial uncertainty of 

28 m (as determined by Lunar Prospector orbit determination results). The only way to 

meet the OD requirements is to use LSE Doppler data to generate an updated lunar 

gravity model specific to the 30 km polar orbit. Mars Global Surveyor OD accuracy 

improved three-fold, from beginning of mission to end, due to gravity model tailoring. 

[Technical contact: Mark Beckman, David Folta] 

3.7 Magnetospheric Multiscale Mission (MMS) 

http://stp.gsfc.nasa.gov/missions/mms/mms.htm

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. An Announcement of Opportunity for the instrument 

complement and principle investigator teams is scheduled for September, 2002, selection 

of multiple payload science teams in early 2003, and final selection of one team about a 

year later. 

The mission consists of four science phases. The main result of the past year’s effort has 

been to determine the general characteristics of the trajectories for these phases. Phase 3 

involves two lunar flybys to set up the trajectories for phase 4. The analysis effort is not 

complete, but much about the orbit dynamics for these phases has been learned and added 

to what we already knew about phases 1 and 2. After the Principal Investigator teams 

have been selected, the analysis for all four phases can become more specific. 

A paper, "The Double Lunar Swingby of the MMS Mission," authored by personnel 

from a.i. solutions, Inc, was presented in December 2001 at the16th International 

Symposium on Space Flight Dynamics in Pasadena, California. The complete paper is 

available at http://issfd.jpl.nasa.gov/sessions/10session/64_Edery.pdf. 

Two more papers are in preparation for a conference in early 2003. 

 [Technical contact: Charles Petruzzo] 

3.8 Themis 

Themis was one of four MIDEX missions selected on April 17, 2002 for Phase A studies. 

Down selection to two will occur by March 2003. If selected, THEMIS will be launched 

in the spring/summer of 2006. 

THEMIS's five identical probes measure particles and fields on orbits which optimize 

tail-aligned conjunctions over North America. Ground observatories time auroral breakup 
onset. Three inner probes at ~10Re monitor current disruption onset, while two outer 

probes, at 20 and 30Re respectively, remotely monitor plasma acceleration due to lobe 

flux dissipation. THEMIS is complementary to MMS and a science and a technology 

pathfinder for future STP missions. 

The FDAB is supporting the University of California in the design, development, 

integration, test, and launch of THEMIS. Specifically, the FDAB is consulting in the 

areas of attitude/orbit control and determination, autonomous navigation, and GN&CD 

systems. 

[Technical contact: Mark Beckman] 

3.9 Venus Sounder for Planetary Exploration (VESPER) Discovery Proposal 

The Flight Dynamics Analysis Branch of the GSFC Guidance Navigation and Control 

Center is supporting a Venus orbiter Discovery Proposal, Venus Sounder for Planetary 

Exploration (VESPER), being led by the Goddard Space Flight Center's Planetary 

Systems Branch. VESPER will integrate key measurements with atmospheric models to 

investigate the coupled processes of chemistry and dynamics in the Venus middle 

atmosphere; the goal being to conduct a tightly focused study of the Venus atmosphere as 

part of a larger NASA program of comparative planetology. VESPER consists of a 

spacecraft and an atmospheric entry probe and will nominally launch in 2008. The Flight 

Dynamics Analysis Branch has analyzed launch vehicle requirements, generated nominal 

trajectory data, and analyzed potential probe impact locations for a 2008 launch. 

[Technical contact: Greg Marr] 

3.10 General Orbit Determination Error Analysis 

The Flight Dynamics Analysis Branch of the GSFC Guidance Navigation and Control 

Center is supporting a range of orbit determination (OD) error analysis studies for future 

and current space missions. In the last year, orbit determination error analysis has been 

performed in support of the Triana, Constellation-X, and Gravity Probe-B (GP-B) future 

missions. The capability to model TDRSS differenced one-way Doppler tracking data 

was recently added and has been used to analyze critical early orbit support of the Galaxy 

Evolution Explorer (GALEX) mission and to analyze a 48 hour tracking campaign to 

generate precise orbit determination solutions in support of the Thermosphere, 

Ionosphere, Mesosphere Energetics and Dynamics (TIMED) mission. 

[Technical contact: Greg Marr] 

3-11 Integrated Mission Design Center (IMDC) 

http://imdc.gsfc.nasa.gov/
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. 

Engineers from the FDAB provided analysis and design support to the IMDC’s 

customers in the areas of Attitude Determination and Control (ACS) and Flight 

Dynamics. ACS products included 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. Many of these innovative 

solutions were synthesized to meet the need for high precision requirements in L2 and 

formation missions. Flight dynamics products included trajectory design and the 

computation of trajectory-dependent quantities needed by other subsystem engineers. 

Among the studies conducted, 7 were for near Earth orbit, 1 was a multipart study for 

geosynchronous orbit, 9 were for trajectories near L2, and 1 was for lunar orbit. Among 

these, five involved formation flying. 

In the past year, ACS and Flight Dynamics IMDC personnel have contributed to 

improving the design process with new tools and design methodologies. Much of this 

was accomplished by taking advantage of the breadth of experience found within the 

FDAB. Some new tools are in the nature of utilities and have been made available to the 

FDAB in general. 

[Technical contacts: Paul Mason, Charles Petruzzo] 
4.0 Technology Development Activities 

4.1 Advanced Mission Design 

The ultimate goal of Advanced Mission Design is to develop and integrate improved 

methods that allow us to design more complex missions and to minimize the cost of 

flying these missions. From a simple request to reduce the amount of fuel to achieve an 

orbit or to compute unique trajectories using new mathematical methods, this task aids us 

in helping spacecraft engineers and scientist to accomplish their goals. From this effort, 

we incorporate basic components of optimization methods into our mission design 

software tools. We also add capabilities to directly use a branch of mathematics called 

dynamical systems. Using these methods, new orbits were established that encouraged 

science proposals and allowed new missions. Besides designs of single trajectories, this 

activity also supports a suite of general design tools that enable optimal geometric 

designs that meet the constraints for Distributed Space Systems (DSS), which have 

multiple spacecraft in formations. 

This work crosses many projects at GSFC and NASA enterprises as it involves all orbit 

types, many spacecraft, and provides for new technologies. A portion of this work was a 

continuation of the Goddard Mission Services Evolution Center (GMSEC) and Earth 

Science Technology Office (ESTO) funded activities for applications to both Earth and 

Space Science Enterprises (ESE and SSE). The Technical Readiness Level (TRL) of the 

research varies, as some optimization techniques are clearly understood but how we 

should best apply them to orbit design is not. Recent successful optimization analysis has 

been performed in support of the Global Precipitation Measurement (GPM) constellation, 

the Solar Dynamics Observatory (SDO) orbit transfers, and the Laser Interferometer 

Space Antenna (LISA). The LISA formation and control optimization consists of three 

spacecraft flying 5 million kilometers (km) apart in the shape of an equilateral triangle. 

This work helps satisfy the charter of GSFC to reduce the cost of access to space, provide 

innovative technologies, build capabilities, and transfer this knowledge to the academic 

and commercial communities. The technical investigations and developments further 

support the resident expertise particularly within the context of libration point orbit 

analysis, transfer trajectory design, and general formation establishment and 

maintenance. The work enhances the theoretical understanding of the multi-body 

problem and offers the advantages available by incorporating the dynamical relationships 

into formation flying design. The models and techniques developed provide immediate 

results for mission support, thus enhancing GSFC participation in proposals while 

expanding capabilities. 

The Advanced Mission Design work described here covers flight dynamics areas 

important to all trajectory design. These include optimization of orbits to meet science 

and engineering requirements while minimizing maneuver impacts, application of new 

mathematical methods to ensure optimal design, investigation of unique orbit design, and 

the development of new utilities and algorithms to support GSFC missions. 
4.1.1 Optimization 

The FDAB researches and develops optimization methods for complex and challenging 

missions. In light of last year’s successes with the initial development of algorithms and 

general research, this year’s approach was planned to complete and validate these 

activities. This year a utility was developed that runs algorithms as an executive and 

provides some guidance in the selection of the correct optimization method given the 

orbit conditions and goals. Improved or new optimization methods continue to be 

included into this utility.

Primer Vectors and Sequential Quadratic Program: Figure 4-1 shows a Matlab GUI 

using primer vector theory developed this past fiscal year. The next step is to expand the 

optimization algorithms to enable all orbit designs, add multiple spacecraft capabilities, 

add conditional constraints (like shadows and coverage), and to ensure robustness. We 

improved computational speed by ‘Mexing’ code and investigated various action 

sequences to converge to different neighboring trajectories. That is, we have the 

capability to add simultaneously two impulses, add an impulse before moving the timeof-flight, add internal impulses, and etc. 

Optimization Tools: To provide a true capability to ensure the optimal trajectory design 

we require methods or a set of integrated tools that can perform both mission 

optimization perform sensitivity analyses in a very general way. It is general in the sense 

that the analyst can define performance measures, constraints, and independent variables 

flexibly according to different mission requirements. An approach where the mission 

sequence can be driven by an optimizer or Monte-Carlo function provides much of this 

flexibility. Because Matlab is quickly becoming an industry standard for engineering 

design, it seems an obvious choice to leverage off of its optimization and statistical 

capabilities. There are many other choices of direct methods that determine if an orbit is 

Figure 4-1. Primer Vector Analysis Tool (PVAT) 
optimal and indirect methods that post-process the orbital data to see if the trajectory is 

optimal. The goal is to baseline several methods with enough information to allow correct 

choices for the orbit type, constraints, or engineering aspects (propulsion system, etc.) 

By permitting a level of generality, we can solve many diverse problems that we 

currently don’t have the capability to solve and vastly improve the time and quality of 

problems we currently handle. Below is a brief highlight of necessary capabilities that 

provide a design environment with the desired capability and flexibility: 

- Primer Vector analysis using the PVAT utility. 

- Sequential Quadratic Programming (SQP) and Projected Gradient Methods 

- Genetic Algorithm applications. 

- Constrained non-linear optimization for both impulsive and low thrust applications. 

- State Transition Matrix propagation to aid in methods or to compute perturbations. 

- Linear and non-linear controller capabilities, i.e. acceleration input from a controller 

for the spacecraft to steer itself on the proper orbit. 

Genetic Algorithms: Genetic algorithms (GA) are established methods for the 

determination of globally optimal conditions that can be used in a wide range of 

missions. By global, we mean that the orbit or trajectory computed is the best choice. 

Using this characteristic, we incorporated GA algorithms into our direct methods to 

ensure a globally optimal solution for minimizing fuel or other orbital conditions. We 

also considered formation flying mission dependent problems such as, finding the best 

launching orbit (subject to launch vehicle constraints) that would allow the minimization 

of fuel to initialize spacecraft in their final operational orbit and formation shape. 

4.1.2 Unique Orbits 

We analyzed unique orbits using a dynamical system approach. Orbits in the vicinity of 

the collinear libration points in the Sun-Earth and Earth-Moon systems serve as excellent 

vantage locations for scientific investigations involving the Sun, planetary, and 

Earth/Moon environments. We will continue to focus significant development and 

operations activities for NASA in support of such missions. GSFC missions involving 

libration point orbits include Constellation-Xray, Maxim, Stellar Imager and James Webb 

Space Telescope (JWST). The use of multiple spacecraft in a distributed approach to 

perform interferometry and optical measurements not achievable by a single spacecraft 

was one of the major drivers in this effort. Trajectory design and pre-launch analysis, as 

well as on-orbit operations and performance evaluation, for these missions is increasingly 

challenging as more complex missions are envisioned throughout the upcoming decades. 

Orbital Bifurcations: These are the expansion of “typical” libration orbits over an entire 

region called the weak stability boundary region about any planet. We examined all 

accessible regions about the Earth and Earth-Moon system. This effort provides orbits to 

meet unique science requirements that require, for example, dwell times over the poles 

and low energy transfers. As part of this effort, a branch of mathematics called 

combinatorics is being investigated for its applications to trajectory design, and also to 

the aforementioned optimization methods. 
Heteroclinic Connection: We began analysis of the regions for transfers of orbits to the 

Sun-Earth libration points via the Earth-Moon system. In more detail, there are invariant 

manifolds (a vast array of virtual winding tunnels and conduits around the Sun and 

planets) in the Earth-Moon system and similar manifolds in the Sun- Earth/Moon system. 

Utilizing dynamical systems theory, invariant manifolds associated with solutions in the 

vicinity of libration points of the Sun-Earth system and of the Earth-Moon system can be 

easily computed and analyzed. We have been involved with each system individually. 

Nevertheless, understanding the connections between these two systems is of paramount 

importance for new mission design. In fact, these new trajectories follow the natural 

system dynamics and allow the designers to develop new mission concepts to explore 

different regions of the Sun-Earth-Moon system and beyond. No one has yet established 

the analytical methods for computing the complete intersection of these manifolds, which 

can be used to easily transfer a spacecraft from one manifold to another. There is some

related work going on in academia but usually it is not to the fidelity required for GSFC 

use, or it is of an abstract nature and not applicable to orbit design. This work is very 

important to servicing missions such as JWST. 

Figure 4-2. Heteroclinic Connection 

Dynamical System using 4-body perturbations– Generator: This work involves the 

upgrades of algorithms to our Generator trajectory design code. The original algorithms

developed under this work have provided FDAB with unprecedented capabilities to 

design libration orbit transfers that are direct. These updates are focused on the 

application of lunar gravity assist to attain libration orbits, missions routinely supported 

by FDAB. These updates allow a more efficient and faster method for the computation 
of such trajectories. The basic code has been used for JWST and Triana design and will 

be very useful for all future missions such as Constellation-X and FKSI. 

Figure 4-3. Dynamical Systems Utility- Generator 

Quasi-stationary locations (QSLs): These are locations in the Earth-Sun system that 

remain relatively stationary and stable over a long period. It is unique in that there is no 

typical orbital motion like a halo orbit. QSLs have been shown to exist in the analytical 

and restricted three-body problem. The initial conditions for such locations shows 

promise and we must now research the control efforts, expand into elliptical restricted

motion, use a full ephemeris fidelity, and analyze all the locations about the Earth-Moon 

system. The work on Quasi-stationary locations is still new (no other center or 

commercial vendor is performing this analysis) and has only been attempted at GSFC. 

Some interesting results have been found under last year’s effort and will be addressed in 

a presentation to headquarters Office of Space Science. A method of control was 

investigated to maintain the orbit near its initial conditions.

Lunar Orbit

Sun Direction Earth

Trajectory

Earth

Sun’s Direction Spacecraft

Lagrange

Point, L2

Lunar Orbit

Exaggerated dimensions shown

Length, 6 km

Width, 80 meters

•Spacecraft location 4 degrees

above Earth-Sun line

Earth

Sun’s Direction Spacecraft

Lagrange

Point, L2

Lunar Orbit

Exaggerated dimensions shown

Length, 6 km

Width, 80 meters

•Spacecraft location 4 degrees

above Earth-Sun line

w jec o

Figure 4-4. Quasi-stationary Locations 
4.1.3 Integrators 

One arena that benefits from a generic availability of numerical integrators is modeling of 

multiple spacecraft flying in formation. This is unique in that we developed application 

algorithms that do not follow the standard thinking when propagating spacecraft 

(predicting their future orbit). The standard approach is to process multiple spacecraft 

individually, but in some cases it can be performed more efficiently if we mathematically 

construct multiple spacecraft orbital data as a single vector. 

Single Formation State Propagation: To allow tight formations or formations in 

locations that have similar perturbations, a method of solving a full state (3 or 6 DOF) to 

propagate either simultaneously or in sequential fashion as is widely done now was 

completed. This allows a vast improvement of formation geometry control. There are a 

number of issues that bear study when modeling the propagation of multiple spacecraft. 

Among these are the basic approach to integration of such weakly coupled systems, 

numerical precision issues that arise when the propagation step size is controlled by the 

most sensitive member of the formation, and issues of efficiency that can arise for either 

a linked propagation or independent propagation scheme. A key element required for any 

spacecraft mission design is a precision orbit propagator, designed to match the orbit of 

the spacecraft. Precision numerical integrators exist in a number of commercially 

available tools used for flight dynamics planning and operations. These integrators are 

tightly coupled to the systems that employ their services. Prototype systems that require 

integration services must either implement the propagation algorithms or make use of an 

existing system in its entirety in order to model a satellite’s orbit. 

We developed a suite of spacecraft propagators that share a common, easily integrated 

application-programming interface (API). The system includes C++ code implementing 

several Runge-Kutta integrators, a Bulirsch-Stoer integrator, and a predictor-corrector 

integrator based on the Adams-Bashford-Moulton method. Each of these integrators are 

designed to work with either a single spacecraft or with multiple spacecraft 

simultaneously. 

[Technical Contact: David Folta]

4.2 Autonomous Navigation Technologies 

Autonomous navigation provides highly accurate onboard inertial and relative navigation 

for multiple satellites. This enables many advanced mission concepts such as formation 

flying, solar sailing, and low-thrust orbit transfer. It also enhances autonomy for all 

aspects of mission operations including maneuver planning and execution, 

communication signal acquisition, real-time onboard attitude determination and control. 

The FDAB approach maximizes design flexibility by providing a single navigation 

software system, GEONS, for multiple mission scenarios. 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. This year, GEONS Release 1.2 was delivered, 

which primarily includes new celestial object and intersatellite line-of-sight measurement 

processing capabilities. This delivery, which includes both the configured flight software 

and the associated documentation, represents a major milestone in the development of the 

GEONS flight software. 

Mission Orbit 

GPS SV GPS SV

GPS SV

Groun -Statio -t - Satellite measuremen

GP

measuremen

GPS Crosslin, Doppler 

Celestial Crosslin, and Doppler 

Satellit-t -Sun direction

measuremen

Satellit-t -Eart direction measuremen Groun -Statio -t - Satellite measuremen

Crosslin measuremen

Crosslin measuremen

Su -Eart Libratio Point (e.g. Constellation X, Su

Eart

L

L 1.5M

150M

Low to Very High Earth Orbits (e.g. 

50R Moo

+

Figure 4-5. Autonomous Navigation Supported by GEONS 

4.2.1 Precise Relative Navigation 

Options to reduce the GEONS code size for minimal capability DSP-type 

implementations were implemented. A thorough analysis of options for implementing 

GPS carrier-phase measurement processing in GEONS was performed. 

Recommendations were developped for a low-risk, cost-effective implementation 

approach that should provide a relative navigation capability that meets the requirements 

of the AFRL TechSat-21 and XSS-11 flight experiments. These experiments will fly 

GEONS integrated with the ITT Low Power Transceiver. 

A technical paper that compares autonomous relative navigation performance for 

formations in eccentric, medium and high-altitude Earth orbits using Global Positioning

System (GPS) Standard Positioning Service (SPS), crosslink, and celestial object 

measurements was prepared. The paper demonstrates that, for close formations, the 

relative navigation accuracy is highly dependent on the magnitude of the uncorrelated 

measurement errors. This paper was presented at the International Symposium on 

Formation Flying in Toulouse France, October 29-31, 2002. 

[Technical contact: Russell Carpenter]
4.2.2 Celestial Navigation 

A detailed investigation of the autonomous navigation accuracy achievable for the 

Constellation-X libration point orbit using only ground-to-spacecraft Doppler 

measurements was completed. This study indicates that ground-to-spacecraft Doppler 

measurements are not sufficient to achieve/maintain a stable solution for this orbit using a 

realistic tracking scenario with realistic measurement errors. Additional analysis was 

performed of the accuracy that can be achieved using only line-of-sight observations of 

the Sun, Earth, and Moon. Based on realistic simulations using optical, star-tracker-like, 

and digital-sun-sensor-like sensor accuracies, the results indicate steady-state accuracies 

ranging from 10 to 30 kilometers are achievable depending primarily on the magnitude of 

the sensor biases and the spacecraft attitude errors. Adding in a pseudoangle 

measurement from North star to the moon significantly improves the accuracy to below 5 

kilometers. Relative position accuracies were also investigated using the pseudoangle 

observations listed above with the addition of intersatellite crosslink range and line of 

sight observations. These results indicate relative position accuracies on the meter-level 

can be achieved for a Constellation-X orbit. 

[Technical Contact: Cheryl Gramling] 

4.2.3 High Altitude GPS Navigation 

A study of GPS navigation for future GOES spacecraft was performed. Based on 

simulations of realistic GPS navigation scenarios that included momentum unloads and 

East-West and North-South station-keeping maneuvers, it was demonstrated that, if the 

maneuvers are modeled, GEONS can provide excellent performance both during and 

following the maneuvers. The GOES system engineering support contractor will use this 

information to develop the concepts for future GOES missions. 

Significant flight data returned from the AMSAT-OSCAR-40 (AO-40) high altitude GPS 

flight experiment was also analyzed. AO-40, an amateur radio satellite launched 

November 16, 2000, is currently in a low inclination, 1000 by 58,800 km altitude orbit. 

Previous experiences with GPS tracking in such orbits have demonstrated the ability to 

acquire GPS signals, but very little data were produced for navigation and orbit 

determination studies. The GSFC AMSAT experiment, developed jointly with the 

GN&CD Components and Hardware Branch, is the first to demonstrate autonomous 

tracking of GPS signals from within a HEO with no interaction from ground controllers. 

The receiver has returned a continuous stream of code phase, Doppler, and carrier phase 

measurements useful for studying GPS signal characteristics and performing postprocessed orbit determination studies in HEO. On several occasions when the receiver 

was below the GPS constellation (below 20,000 km altitude), observations were reported 

for GPS satellites tracked through side lobe transmissions. Although the receiver has not 

returned any point solutions, there has been at least one occasion when four satellites 

were tracked simultaneously, and this short arc of data was used on the ground to 

compute point solutions. 
Three technical papers have been presented covering various aspects of the AMSAT GPS 

experiment. The first, presented at the AAS Guidance and Control Conference in 

Breckenridge, CO, described the design of the GPS experiment, and provided some 

examples of the GPS data returned from the AO-40 spacecraft. The second paper, 

presented at the AIAA Guidance, Navigation and Control Conference in Monterey, CA, 

described the initial efforts to generate AO-40 navigation solutions from pseudorange 

data reconstructed from the GPS receiver’s code phase, as well as to generate a precise 

orbit solution for the AO-40 spacecraft using a batch filter. The third paper, which was 

presented at the Institute of Navigation’s annual GPS meeting in Portland, OR, describes 

tracking performance, measured power levels from the GPS satellites, and orbit 

determination analysis based on GEONS processing of the data. 

[Technical contact: Mike Moreau] 

4.2.4 Magnetometer-based Navigation 

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 GN&CD 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. 

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 

data from the TRACE spacecraft. The magnetometer-GPS configuration has undergone 

analytical testing with the goal of developing a ‘black-box’ spacecraft navigation system. 

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. 

Based on the success with TRACE data, an inflight experiment of MAGNAV is planned 

for the WIRE spacecraft. During FY02, the MAGNAV algorithm was converted from 

MATLAB code into flight code, as a patch to the WIRE flight software (FSW). 

Numerous tests have been conducted to verify the conversion, and to ensure the code will 

run within the FSW attitude control system (ACS) in the onboard computer. The SMEX 

dynamics simulator was utilized in the testing. Once MAGNAV is successfully patched 
into the FSW, the flight experiment will be run for two weeks to test various scenarios of 

operation, and to verify the onboard performance of MAGNAV. 

[Technical Contacts: Julie Thienel, Rick Harman] 

4.3 Formation Flying Technologies 

Spacecraft formations are a subset of the global collection of multiple spacecraft 

missions, classified as Distributed Space Systems (DSS). In general a DSS is a collection 

of 2 or more space vehicles designed to accomplish similar or shared objectives; an endto-end (information) system consisting of two or more space vehicles, coordinated flight 

management, and an integrated infrastructure for data acquisition, storage, analysis, and 

distribution. In contrast a formation is comprised of multiple spacecraft with the ability 

to cooperatively detect, maintain, and agree on the appropriate maneuver to maintain a 

desired position and orientation. Formation flying is enabling technology required to 

maintain the relative separation, orientation, or position between or among the formation 

spacecraft. The FDAB is pursuing several initiatives, described below, that focus on 

formation flying technology development. 

4.3.1 Autonomous Formation Flying Control of EO-1 

NASA’s first-ever autonomous formation flying mission is a resounding success. With 

the launch of NASA’s Earth Observer-1 satellite, called EO-1, NASA’s Goddard Space 

Flight Center has demonstrated the capability of satellites to continuously fly in 

formation, to react to each other, and to maintain close proximity without human 

intervention. This unique advancement has been highlighted in Aviation Week and Space 

Technology and on space related web sites. The new capability allows satellites to 

autonomously react to each other’s orbit changes quickly and 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. 

On EO-1, formation flying was required to calibrate and compare technological advances 

made in ground observing instruments that are smaller, cheaper, and more powerful. 

Until an onboard algorithm and system were developed at GSFC, the requirements could 

not be easily met. Previously, satellites did not plan nor 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 called AutoCon 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 

maneuver algorithm is designed as a universal 3-dimensional method for controlling the 

relative motion of multiple satellites in any orbit. The AutoCon architecture and 

mathematics were first developed by Aerospace Engineers at the Goddard Space Flight 
Center (GSFC). Their idea was then combined with a new flight software that is the 

commercial predecessor of a GSFC sponsored commercial software called FreeFlyer, 

produced by Lanham, MD based a.i.-solutions, Inc. 

This unique onboard demonstration establishes the following NASA capabilities: 

• A flight demonstrated, validated, fully non-linear autonomous formation flying 

system. (A NASA first for continuous formation flying) 

• A precision universal control (Folta-Quinn) algorithm with user defined control 

accuracy that can be used for any orbit. 

• A formation flying and orbit maintenance system that allows: 

- Cartesian Control: Alongtrack, radial, and crosstrack 

- Keplerian Control: Semi-major axis, inclination, node, eccentricity 

- Any combination of the above 

- Single or multiple maneuver computations. 

• Proven executive flight code that incorporates fuzzy logic for multiple constraint 

checking for maneuver planning and control. • Use of text scripts for onboard command control, software changes not required. 

• Utilization of various Navigation inputs (GPS, TONS, Ephemeris, etc.). 

• Attitude (quaternion) required of the spacecraft to meet the V components. 

• Generation of maneuver commands onboard. 

• Calibration of the maneuver onboard. 

There are many benefits of this onboard formation flying system. 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. Onboard formation 

control enables 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. The formation flying technology flown onboard EO-1 will make 

distributing scientific instruments over many separate satellites routine and cost effective. 

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. 

Application to future NASA GSFC missions is beginning with the decision of the Global 

Precipitation Measurement mission to fly it for autonomous maneuver control to improve 

the science predictions of the orbit. 
Is it 

Planning

time?

No

Yes Get S/C 

States

Is EO-1

 Near Control Box 

Limits?

Yes

No

Propagate

States

Is The

 planning cycle 

finished?

No

Yes

Design and 

Model

Maneuver

Generate

Maneuver 

Commands

Select

Control 

Algorithm

Design and 

Model

Maneuver

Generate

Maneuver 

Commands

JPL GSFC

GUI or EFF Interface

Design and 

Model

Maneuver

Generate

Maneuver 

Commands

Industry

Algorithms

(Extended

Mission)

Maneuver Time

& Delta-v

Figure 4-6. AutoCon Control Flow

Code 572’s Autonomous Maneuver Control (AutoCon) was selected as a runner-up in the 

2002 NASA Software of the Year competition. This marks the 5th year in a row that a 

Goddard software product has been selected as runner-up or honorable mention for the 

award. Formation Flying is one of the most exciting concepts permeating NASA today and 

was enabled by AutoCon for the first autonomous demonstration. 

[Technical Contact: David Folta] 

4.3.2 Tethered Formation Flying 

The Sub-millimeter Probe of the Evolution of Cosmic Structure (SPECS) is a bold new 

mission concept designed to address fundamental questions about the Universe, including 

how the first stars formed from primordial material, and the first galaxies from pregalactic structures, how the galaxies evolve over time, and what the cosmic history of 

energy release, heavy element synthesis, and dust formation is. Ideally, a very large 

telescope with an effective aperture approaching one kilometer in diameter would be 

needed to obtain high quality angular resolution at these long wavelengths, however this 

approach proves to be too expensive and therefore impractical. Instead, a spin-stabilized, 

tethered formation is one possible configuration being considered requiring a more 

advanced form of formation flying controller, where dynamics are coupled due to the 

existence of the tethers between nodes in the formation network. To this end an 

investigation into the dynamics and control of multiple tethered spacecraft systems was 

launched. 
Figure 4-7. Tethered Formation Flying Satellite Dynamics Tool (SDT) – Satellite 

Tool Kit (STK) Connection 

The effort is divided into three separate tasks. Task-1 involves working with mission

scientists in an effort to understand and document the science requirements of a SPECS 

class facility. GSFC science and optics specialists are working with engineers from 

Payload Systems Inc. (PSI) and the Naval Research Laboratory (NRL) to capture the 

science requirements and flow those down to understand the engineering requirements 

imposed by the desired science. Using the Generalized Information Network Analysis 

(GINA) framework developed at the Massachusetts Institute of Technology, PSI will then

conduct detailed configuration trade studies. Many possibilities will be examined in an 

effort to conduct a "broad-brush" analysis so as to get an understanding of the favorable 

regions in the trade space and suggest candidate configurations capable of meeting the 

science requirements. 

Task-2 involved cooperation of GSFC with engineers at NRL in the development of the 

equations of motion for a rotating multi-tethered system applicable to the study of 

fundamental dynamic characteristics of a deep space interferometer concept. The system

is assumed to be comprised of n-particles inter-connected by any number of tethers in a

user defined configuration. The concept system is intended to execute planar rotation, to 

deploy and/or retract the tethered masses, and be capable of re-orienting the spin axis.

Accordingly, the mathematical model allows full three dimensional motion of all the
constituent elements. At this writing all masses are all assumed to be 3-DOF particles and 

a model involving 6-DOF finite masses is under development. The NRL model can be 

used in either of two modes. In one mode, the user specifies the forces to be applied to 

the particles and the resulting dynamics are computed. This is the standard means of 

implementing controls. To aid in the development of control laws for such a complicated 

system, as second mode was designed into the dynamics model. In this mode, the desired 

dynamics are prescribed by the user and model outputs the forces necessary to give rise 

to that motion. It is believed that this capability will aid controls designers in their efforts. 

NRL’s n-particle dynamics model has been implemented into Star Technologies Satellite 

Dynamics Tool (SDT) application which in turn interfaces to Analytical Graphics 

Satellite Tool Kit (STK) application. The result is a system which allows the user to 

construct n-particles inter-connected by m-tethers in any user defined configuration. SDT 

will then use the imbedded NRL model to compute the dynamics of the particles, 

generating ephemeris data that is handed over to STK for 3D graphical output. 

Working with the Virginia Tech, Task-3 is examining key linear and non-linear control 

methodologies which may prove applicable to the problem of tethered formation flying, 

specifically gain-scheduled controllers, Lyapunov based non-linear controllers and/or 

robust adaptive controllers. The object was to build upon the dynamic development from 

Task-2 in order to create a core dynamics and control model which permits iteration and 

expansion while maintaining the primary thrust of the tethered formation. The ultimate 

goal for this task is to develop a set of control laws centered on the core model from 

Task-2. This will serve as a first order tool for examining the dynamics and control of a 

variety of design configurations. 

Finally, as the project comes together the configuration trades will narrow the trade space 

identifying at least one candidate configuration which can be examined in greater detail. 

A model of the candidate configuration will be constructed in the SDT application using 

NRLs dynamics model. Virginia Tech’s controls will be imposed on the dynamics and a 

comparison made to the original requirements for scientific success. In this way we will 

have constructed an important capability and taken a crucial step towards making a 

SPECS type mission a real possibility. 

4.3.3 6DOF Nonlinear Control for Formation Flying 

Virtual platforms, based on spacecraft formations, form the strategy for improving the 

spatial and angular resolution achievable for space-based observatories. Stellar Imager, 

and the Micro-Arcsecond X-ray Imaging Mission (MAXIM) are typical missions based 

on this design concept. Precision formation flying falls among the enabling technologies 

required for mission feasibility and success. The section describes current research 

related to control algorithms for precision formation flying. The specific area of research 

considers the problem of simultaneous control of a spacecraft’s orbit and attitude, in 

order to meet the strict mission requirements (micro-arcsec pointing, and sub-millimeter 

position control). 
Spacecraft orbit and attitude trajectories are governed by nonlinear dynamics, and 

combined allow for six degrees of freedom (6DOF) in motion. Research targets the 

development of a 6DOF, nonlinear control algorithm for achieving the design 

specifications for missions requiring precision formation flying. Algorithm development 

is based on the assumption that these missions will be stationed near the L2 point in the 

Earth/Moon – Sun system. Hence, the orbital dynamics are characterized by the 

restricted-three body problem. Further external disturbances are limited to solar pressure, 

and other gravitational sources. To date the approach has been proven to work for 

relative position control. Accomplishments are published in several papers, noted in the 

references below. Current work is focused on incorporating attitude control into the 

position control algorithm, generating a full 6DOF solution. 

For further information, refer to the following: 

1. Luquette, R. J. and Sanner, R.M. “A Non-Linear Approach to Spacecraft Trajectory 

Control in the Vicinity of a Libration Point”, Flight Mechanics Symposium, Goddard 

Space Flight Center June 19-21, 2001. 

2. Luquette, R.J. and Sanner, R.M. “A Nonlinear Approach to Spacecraft Formation 

Control in the Vicinity of a Collinear Point”, AAS/AIAA Astrodynamics Specialist 

Conference, 30 July – 2 August 2001, Paper No. AAS 01-330. 

3. Luquette, R.J., Leitner J., Gendreau, K., and Sanner, R.M., “Formation Control for 

the MAXIM and MAXIM Pathfinder Missions”, Proceedings of SPIE Astronomical 

Telescopes and Instrumentation Conference, 22-28 August 2002. 

[Technical contact: Rich Luquette] 

4.3.4 Maxim Pathfinder Testbed 

The FDAB provided engineering support of the Maxim Pathfinder X-ray Interferometer 

test bed. Activities included the design and fabrication of a precision double-slit and 

mirror control system used to focus an eighty meter long x-ray interferometer beam onto 

a five millimeter CCD detector with sub-arcsec, sub-micrometer accuracy. The computer 

control system provides a real time, hardware-in-the-loop foundation with general 

applicability to the integrated Formation Flying test bed. Specifically, the new system 

addresses the FFTB's need for a real time simulator with the capability to interface with 

numerous GN&C and instrument subsystems. The current incarnation of the system 

utilizes a dual processor computer control and data handling architecture, but was design 

to be N-processor scalable in order to accommodate future scenarios with large number 

of spacecraft and/or complex interface requirements. 

[Technical contact: Steve Queen] 
4.3.5 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. 

This research has been ongoing for several years, and is highlighted in the FDAB, End of 

Fiscal Year 2001 Report. Research has focused on investigating implementation issues 

and testing in a relevant environment. Testing activities have augmented existing 

resources at Goddard with those at the University of California at Los Angeles (UCLA), 

the Naval Postgraduate School (NPS), and the Massachusetts Institute of Technology 

(MIT) through appropriate arrangements. 

 [Technical contact: J. Russell Carpenter] 

4.4 Attitude Determination 

4.4.1 PiVoT Attitude Determination 

The PiVoT GPS receiver provides position, velocity, and time using GPS. The objective 

of this work is to make attitude available from PiVoT. Attitude is determined as a 

solution via a minimum mean-squared estimate (MMSE) of the nine parameters in the 

3x3 attitude matrix. Thus as many satellites as are currently being tracked can be used to 

improve the attitude estimate. Two methods are being made available for making the 

resulting 3x3 matrix a proper orthonormal attitude matrix: cross products of columns 

(speed) vs singular value decomposition (accurate, but slower). The MMSE approach 

allows each attitude solution to be independent of previous ones (as opposed to using 

small angle approximations as has been done for SIGI on the International Space 

Station). The attitude solution can be expected to handle higher rotation rates than does 

SIGI. 

The attitude solution runs both double-differencing and single-differencing algorithms in 

tandem. Single differencing requires that the "line bias" parameters (effectively time 

delays between antenna centers and the receiver) be known, whereas double-differencing 

uses one additional satellite to cancel out the need for knowing the line bias. There are 

two costs to using double-differencing: one more satellite needs to be tracked and a 40% 

noise increase (due to the subtraction of two Gaussian noise variates). The difference in 

the single and double difference solutions is used to estimate the line bias; once the line 

bias settles down, the single difference solution is presented as the better solution. 
The Landis algorithm for obtaining attitude from two satellites has been incorporated into 

the PiVoT attitude determination software. Thus double-differencing attitude estimates 

can be obtained from tracking only three satellites and, if line biases have been 

determined, single-differencing attitude estimates obtained from only two satellites. The 

use of this algorithm is completely transparent. As a caveat, note that the primary 

position-velocity-time estimates requires four satellites. When the number of tracked 

satellites drops to three satellites, the onboard time can be used to determine position and 

velocity. When the number of tracked satellites drops below three, GPS receivers on 

satellites use an orbit model to estimate its position. Since the attitude in earth-centered 

earth-fixed coordinates depends on where the satellite is relative to earth, such attitude 

estimates will exhibit a growing error. 

The attitude determination software depends on the use of phase differences of the 

incoming signals on the different antennae (interferometric approach). Since the antennae 

are likely to be farther apart than a wavelength of about 18cm, there remains an integral 

ambiguity in the number of wavelengths between the antenna along the signal path. 

Integer ambiguity resolution is done using Iz, Ge, and Chen's grid point search algorithm. 

The approach taken allows some constant rotation during the resolution process, and 

involves singular value decomposition and considerable matrix manipulation and so the 

process is numerically intensive. Once the integer ambiguities are determined, the 

current attitude solution is used to solve for the integer ambiguities for newly tracked 

satellites. 

[Technical Contact: Charles E. Campbell, Jr.] 

4.4.2 GPS Attitude Sensor 

A novel Global Positioning System (GPS) sensor concept is under development that 

would provide data for both navigation and attitude. The Compound Eye GPS Attitude 

and Navigation Sensor (CEGANS) would be equipped with multiple directional antennas 

mounted on the convex hemispherical surface. Each antenna would be aimed to receive 

GPS signals from a restricted, but known visualization cone. By noting which GPS 

satellites are visible in the field of view of each antenna in the hemispherical array, the 

attitude of the sensor (and therefore the body to which it attached) can be estimated to 

within 3 degrees without resorting to the use of carrier-phase measurements. It is 

believed that optimization and signal-to-noise techniques can be applied to refine raw 

attitude estimates from this compact sensor to the sub-degree range. 

A simulation study is underway which is beginning to prove the CEGANS concept can 

work. To date, the sensor is given perfect measurement data and so yields perfect 

solutions. As this idealized simulation is degraded to more closely replicate the true 

environment (addition of noise models etc.), a more realistic performance expectation can 

be formulated. Here again, Star technologies Satellite Dynamics Tool (SDT) is proving 

invaluable in providing the vehicle in which the entire GPS constellation can be modeled 
as well as the RF interfaces to the satellite employing the CEGANS sensor. A patent on 

the CEGANS is expected to be issued by the end of this calendar year. 

[Technical Contact: Dave Quinn] 

4.4.3 Advanced Attitude Determination Methods 

A new star identification algorithm was developed which uses the Expectation 

Maximization (EM) method. EM consists of two steps. First, the log of the probability 

density function is formed and the expectation computed. Second, the argument is found 

which maximizes the expectation. A star field was simulated in which the observed stars 

are rotated 20 degrees away from the reference stars. The EM algorithm was used to

identify the observed stars and to calculate the rotation angle. 

This algorithm was applied to the following cases of star distributions: the number of 

observed stars is identical to the number of reference stars, the number of observed stars 

is less than the number of reference stars, and the initial uncertainty in the rotation angle 

was as large as 40 degrees. The next step is to apply this algorithm to observed mission 

star patterns as processed by our operational star identification system. 

Nutation can be a significant problem for any spinning spacecraft. Most recently, the 

IMAGE spacecraft had a nutation problem that was the result of an undersized damper. 

The problem was solved by using the torquer bar to fight the nutation during each 

perigee. A technique has been developed to estimate nutation angle. Data was simulated 

for a spacecraft rotating at 7 RPM with a nutation angle of 5 degrees and a nutation rate 

of 9.5 degrees/second. Figure 4-8 shows the true sun angle relative the spacecraft spin 

axis as well as the observed sun angles. The final result of the algorithm compared 

perfectly with the nutation angle. 

Figure 4-8. True and Measured Sun Angles for a Nutating Spacecraft
A pseudo linear Kalman Filter was developed to estimate gyro model misalignments, 

scale factors, and bias for a triad gyro configuration normally used in spacecraft. A 

simulation was set up to model the gyro errors and a maneuver about the spacecraft x, y, 

and z axes. Applications for this algorithm include real time gyro calibration on the 

ground as well as onboard the spacecraft. Further work will include adding the gyros as a 

measurement, testing with flight data, and incorporating the system into the Mission 

Three Axis Stabilized Software (MTASS) system. 

[Technical contact: Rick Harman] 

4.5 Flight Dynamics Automation Studies 

4.5.1 Flight Dynamics Automation with The University of Maryland 

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, as well as the initial release of a Web based Graphical User Interface (GUI). A 

report titled “SAMPEX Orbit Determination Automation Plan” was released. 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 modify the products produced to include 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 add flexibility to address future products. (University of Maryland 

SAMPEX Link: http://kepler.umd.edu/sampex/) 

[Technical Contact: Joseph Toth] 

4.5.2 TRMM Reengineering 

The Flight Dynamics Analysis Branch has also partnered with the Mission Applications 

Branch (NASA GSFC Code 583) and a.i. solutions in order to support TRMM 

Reengineering activities. This work is being undertaken using the development of the 

Autonomous Flight Dynamics System (AutoFDS). Work completed to date includes the 

additional development and enhancement of the AutoFDS, as well as participating in 

various meetings, presentations, and demonstrations of the AutoFDS. Work continues on 

the following: further enhancements of the AutoFDS in order to support TRMM 

Reengineering activities which includes product definition, user interface improvements, 

and the addition of capabilities such as file event based execution to the current system. 

Planned work to be carried out under the GSFC Mission Services Evolution Center 
(GMSEC) includes further TRMM Reengineering activities as well as the integration of 

the AutoFDS with the AutoProducts System, and the definition of XML based messages. 

[Technical Contact: Joseph Toth] 

4.6 Nanosat Technology

The Guidance, Navigation, and Control Division (GNCD) has assumed management and 

systems engineering oversight of two existing formation flying nanosatellite missions 

being developed at MIT, Cornell University, Utah State University, and Virginia Tech. 

Emphasis has been placed on usage of proper flight-certified materials to comply with 

shuttle flight safety standards and detailed analysis efforts to determine the proper 

dispensation system to use for ejection from the shuttle bay. The GNCD is also 

overseeing the systems integration and shuttle safety preparation process with prime 

support from Orbital Sciences Corporation, in preparing these missions for a potential 

2003 or 2004 shuttle launch. Further, GNCD is leading an effort to help Goddard Space 

Flight Center and the Air Force Research Laboratory (AFRL) extend their partnership 

from a personnel exchange, facilitating the collaboration on mutual technology and 

research interests, to developing an education and outreach “hands-on” program designed 

to teach and direct the aerospace technology systems engineering process at the 

undergraduate and graduate level engineering curriculum. The intent is to reach a greater 

volume of universities and their students possessing varied expertise in experiment and 

flight hardware development. With the challenges of space access ever present, both 

agencies will share the task of formulating a pathway to increase flight opportunity and 

encourage flight demonstrations targeting the goals and needs of future missions. 

[Technical Contact: Lucien Cox] 

4.7 GMSEC Architecture Development

The Goddard Mission Services Evolution Center (GMSEC) is envisioned as a ground 

system architecture of the future that will provide mission services such as mission 

planning, scheduling, data processing and command management to flight projects. The 

rationale behind establishment of the GMSEC architecture and supporting functions is to 

help retain and extend expertise at Goddard in ground systems and operations, and 

provide a focal point for the development of mission services related technology and 

ground system upgrades. Mission services functions include several flight dynamics 

activities usually required in ground support of flight projects. These include attitude 

determination, ground attitude control, attitude sensor calibration, orbit determination, 

and orbit maneuver planning. As part of its support to current flight projects under 

development, the FDAB works to define and implement software systems required for 

flight dynamics support. The FDAB will also be an active participant in GMSEC 

architecture development activities. During the past year, the branch participated in 
architecture planning activities and proposed mission services technology projects for 

FY03. 

[Technical contact: Thomas Stengle] 
5.0 Branch Infrastructure 

5.1 Flight Dynamics Tool Maintenance 

An active software tool maintenance and development program is required by the FDAB 

to perform future mission studies, flight project mission analysis and operations support. 

In doing so, this effort supports the maintenance, use, and administrative activities 

associated with the FDAB institutional flight dynamics tool suite in response to the needs 

of FDAB engineers. 

FDAB tools consist of a combination of homegrown systems, commercially available 

products, and extensions to commercially available products developed and/or procured 

in support of FDAB flight dynamics engineers. The flight dynamics tools support the 

following activities within the FDAB: attitude error analysis, prediction and 

determination; navigation, orbit error analysis, orbit prediction and determination; and 

mission analysis, trajectory design and maneuver planning. This tools maintenance effort 

also provides the navigation technology engineering support necessary to maintain the 

official releases of the GPS Enhanced Orbit Determination (GEODE) and GPS Enhanced 

Onboard Navigation System (GEONS) flight software, and support integration of 

GEODE and/or GEONS with one or more space receivers or satellite flight computers. 

Specific tasks performed under the Flight Dynamics Tools Maintenance task included: 

• Sustaining engineering support for FDAB institutional flight dynamics tools 

− maintained core expertise associated with FDAB institutional flight dynamics 

tools 

− provide analysis concerning the implementation and use of FDAB 

institutional flight dynamics tools 

− define and conduct maintenance activities associated with FDAB institutional 

flight dynamics tools 

− configuration management and system administration support associated with 

GNCD institutional flight dynamics tools

• Maintaining an inventory of institutional tools 

• Generating and maintaining the FDAB flight dynamics tools Maintenance Plan 

• Delivering upgrades to institutional flight dynamics tools in response to the 

Maintenance Plan 

During the past year the REPEAT utility was converted from Fortran to MATLAB, using 

a language translator that was developed during FY01. Also, a paper was prepared that 

provides the details for strategies to maximize the probability of star identification using 

the Attitude Determination Subsystem (ADS) contained within the Mission Three-Axis 

Stabilized Spacecraft (MTASS) with minimum initial attitude information. Under this 

effort, we also documented all the .m (MATLAB) files in a searchable database. The 

major area of concentration during this past year has been to develop a graphical user 

interface (GUI) for the General Maneuver (GMAN) Program, and to integrate new 
propulsion system modeling into the GMAN Program, thus greatly enhancing 

the versatility of the software tool. Additionally, the GUI could be adapted for use by 

other flight dynamics software tools (e.g. ACQSCAN) fairly easily. On-going support 

for GEODE/GEONS software engineering was also provided. 

In-house FDAB engineers worked during the past fiscal year to maintain and enhance 

major ground-based navigation systems such as the Goddard Trajectory Determination 

System (GTDS). Major work includes: 

• Place identified systems (and versions) under software configuration control by 

the GNCD Lab 

• Build optimal systems from different versions in the UNIX platform (these 

optimal systems will become official GNCD navigation systems) 

• Port all GNCD ground-based navigation systems to PC platforms 

• Develop a user friendly interface for PC versions of GNCD ground-based 

navigation systems 

• Add technical and graphical capabilities to PC versions of GNCD ground-based 

navigation systems 

• Perform research and analysis to improve performance and optimize operational 

use of ground-based navigation systems to support future missions (e.g., as 

backup systems in formation flying in case of failure of onboard navigation 

systems). 

The delivery of GNCD Release 2.1 and a partial delivery of Release 2.2 have been 

completed. Release 2.1 consists of 24 orbit related systems including the following: 

1. GTDS PC Version - This is equivalent to the FDF HP-UNIX Release 2000.02 of 

GTDS. 

2. INITTVHF - This utility allows the user to create and initialize a Trajectory 

Computation and Orbital Products System (TCOPS) Vector Hold File (TVHF) for 

use with GTDS. 

The partial delivery of Release 2.2 includes the following systems: 

1. TLE2ELE - NORAD 2-Line Element to ELCONV format Conversion Program 

2. TLE2NEP - NORAD 2-Line Element to NORADEP format Conversion Program

Work continues on GNCD Release 2.2 of GTDS, which will consist of a port of the HPUNIX Operational Release 2001.01 of GTDS to the Windows environment. GNCD 

Release 3.1 was also started, and planned for delivery in January 2003. It includes the 

following modifications: capability to use geopotential models of up to 360x360 in size, 

including the complete EGM96 geopotential model; capability to use EGM96 enhanced 

earth tide model; and the capability to use central bodies other than the Earth, Moon, or 

Sun. 

In addition, the development of a sophisticated user interface and graphics capabilities for 

GTDS is in progress with the support of Code 583 software engineers. 
[Technical Contacts: Son Truong, Joseph Toth, John Lynch] 

5.2 Guidance, Navigation and Control Lab 

The Guidance, Navigation and Control Lab (the Lab formerly known as the Flight 

Dynamics Analysis Lab) under the oversight and management of the Flight Dynamics 

Analysis Branch continued to provide a state of the art computing environment for both 

the FDAB engineers and the GNCD engineers. Consolidation of the computing domains 

was complete in FY02 and new domain servers were put online. These servers increased 

the online storage capacity to several terabytes. In addition, a new tape backup system 

was put online. This new tape backup system increases backup capability to nine 

terabytes and gives improved performance and expandability. New user workstations 

were added to the lab that provide the engineers with increased computational capability 

and allow for major analysis projects to be completed in a more timely manner. 

The GNC Lab hosted several classes and demos in FY02. Hosting classes for the COTS 

engineering tools used by FDAB and division engineers allows them to receive training 

on the systems they will use to run the software. 

The GNC Lab also provided operations support during several International Space 

Station (ISS) EVAs. 

[Technical Contact: Sue Hoge] 

5.3 Flight Dynamics Models 

5.3.1 SKYMAP Maintenance 

The SKY2000 Master Catalog (MC) stellar database is the starting point for SKYMAP 

mission star catalog generation, and is also used in many non-SKYMAP systems 

developed by the astronomy/space science community. The SKY2000 MC contains 

~300,000 entries, is complete to ~8 visual magnitude (Mv) with some entries fainter than 

Mv = 10. The current MC was developed and refined using the best data from multiple 

star catalogs. Improvement in the quality and quantity of MC data is a continuous 

process as new catalogs are released, more star tracker data are collected and existing 

data are incorporated. 

An important capability of the SKY2000/SKYMAP System is the prediction of 

instrumental passband magnitudes (Mi), principally star tracker magnitudes for space 

missions using the Multi-Mission Star Catalog (MMSCAT) generation software. 

Missions currently using SKYMAP star catalogs include: RXTE, SWAS, Terra, Aqua, 

Landsat-7, GOES I-M, UARS, SOHO, and ACE. 
The SKY2000 Task operates under a limited budget, so yearly improvements to the MC 

and MMSCAT software are moderate and must be prioritized. Two activities completed 

this year include an update to the SKYMAP Web page and some improvements to the 

MMSCAT executable to enhance near-neighbor/magnitude blending subsystems. See the 

link above to the SKYMAP Web page for general information on the SKY2000 MC, 

MMSCAT, and to download the entire SKY2000 MC or selected mission catalogs and 

delivery memoranda. 

Task personnel attended the Star Catalog Working Group meeting at USNO 

(Washington, DC) in January 2002. They led a discussion group on star catalogs and 

software tools. The task lead was nominated as co-chair of a working group examining 

issues involving a single, centralized star catalog for the various user communities. 

 

Task personnel also provided consultative support in response to questions regarding the 

SKYMAP Master Catalog and related issues, and in response to questions regarding star 

catalog/star tracker issues faced by current and upcoming missions (including RXTE, 

Aqua, ACE, Landsat-7, and Swift). 

Visit the website: http://mmfd.gsfc.nasa.gov/prod_center/pc_frame_page.htm 

[Technical contact: David Tracewell] 

5.3.2 Solar Flux Predictions 

The FDAB continued to support periodic updates of the Schatten solar flux predictions. 

These predictions are used for future mission analysis and for decay studies of 

operational missions. 

Solar activity during 2002 has been slightly higher overall than predicted by the Schatten 

nominal solar flux activity predictions (Figure 5-1). This plot shows the actual solar 

activity (the jagged line) along with the –2 sigma, nominal and +2 sigma prediction 

curves (bottom, middle and top smooth curves) for 2001 & 2002. While the measured 

solar activity was lower than predicted (centered on the –2 sigma curve) for the first 9 

months of 2001, there was a sharp increase in activity in the Fall of 2001, from ~160 to 

225 F10 Radio Flux. The solar activity level continued to be more than the +2 sigma 

prediction for approximately 6 months, but has decreased and is currently centered on the 

+2 sigma prediction curve, and is trending towards the nominal prediction. It should be 

noted that the Schatten solar flux predictions are not meant to capture these short-term 

variations, and are used as a more general trending tool for long-term predictions. 

However, the increase in predicted solar activity did cause low-Earth satellite orbits to 

decay more quickly than the preceding year. 
F10.7 Prediction for 2001,2002

0

50

100

150

200

250

2000.5 2001 2001.5 2002 2002.5 2003 2003.5

time, Year

F10.7 , Radio Flux

Figure 5-1. Predicted and Actual Solar Activity for 2002, 2002 

[Technical Contact: Karen Richon 

5.4 FDAB Website 

During this FY, branch members worked to refurbish the FDAB website. Information 

was added concerning all of the missions currently being supported by the branch. 

Information on typical analysis provided to customers as well as on our tool suite and 

major research areas is now documented on the site. Pictures of branch members were 

updated, and news items were added. Current and past End-of-Year reports are available, 

as well as a partial reference list of the papers published by branch members. Future 

capability for a searchable database of published papers is planned for next fiscal year.

Visit the FDAB website at: http://fdab1.gsfc.nasa.gov/ 

[Technical Contact: Sue Hoge] 

 
6.0 Interagency Activities 

6.1 X43-A Failure Investigation 

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 wingspan of about five feet 

(see figure). The first X43 and its modified Pegasus-XL booster rocket was 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. At that point, a 

decision was made to terminate the flight. In support of the failure investigation board, 

FDAB personnel provided support in: (a) multi-rate analysis of the autopilot, and (b) 

assessing Monte Carlo and perturbation analyses. 

Figure 6-1. X-43 Air-Breathing Hypersonic Aircraft 

The OSC’s Autopilot lateral and longitudinal design and linear analysis models were 

examined. Specifically, the continuous-time implementation of a 100-Hz digital bending 

rate filter in the analysis was flagged for further investigation. The autopilot’s main loop 

ran at 25 Hz, hence a need for multi-rate analysis. An analytical multi-rate 

implementation of the filter within the autopilot loop was formulated and verified via an 

independent MATLAB simulation. The multi-rate formulation was also validated via 

independent SIMULINK simulations. The analytical multi-rate formulation was provided 

to the Hyper-X team for future implementations. 
An assessment of the adequacy of the pre-flight Monte Carlo analyses, stress tests, and 

sensitivity analysis were made. Parameters, parameter variations, distributions, and 

correlations in several key areas (aerodynamics, Fin actuation System, and GN&C) were 

investigated for adequacy. Additional parameters to be included in the return to flight 

analyses were identified and reported. 

[Technical Contact: Peiman Maghami] 

6.2 SCISAT 

SCISAT is Canada’s first science satellite in over 30 years. The mission is to study the 

Earth’s upper atmosphere ozone layer, predominantly over Canada. As the spacecraft 

orbits the Earth, there is a Sun rise and Sun set every 90 minutes. During those events, the 

spacecraft uses the Sun light cutting through the upper atmosphere to measure the ozone 

levels. 

Due to the very long time between missions, the Canadian Space Agency (CSA) invited 

the FDAB attitude control experts to review its design. With concurrence from NASA 

Headquarters, the review was held in Winnipeg Canada. Bristol Aerospace presented its 

design as well as its development flow. The FDAB experts commented on the design and 

concentrated on the integration and testing process. Both Bristol Aerospace and CSA 

engineers found the meeting to be very helpful. The senior ACS engineers from Bristol 

Aerospace and FDAB continue to stay in contact, receiving updates and giving advice. 

SCISAT is scheduled to launch in January 2003. 

[Technical contact: Dave Mangus] 

6.3 NASA Technical Standards Program 

http://www.ccsds.org/

http://standards.gsfc.nasa.gov/

The FDAB supports the NASA Technical Standards Program by contributing to the work 

of the GSFC standards program, the NASA Data Standards Steering Council (DSSC), 

and the Consultative Committee for Space Data Systems (CCSDS). The GSFC standard 

program aims to expand the scope of best practices, and to develop an agency-endorsed 

database of preferred technical standards for NASA. The Data Standards Steering 

Council (DSSC) is the hub of the NASA Data Systems Standards Program and is 

sponsored by NASA Headquarters. The Consultative Committee for Space Data Systems 

(CCSDS) is an international organization of space agencies interested in mutually 

developing standard data handling techniques, to promote the interchange of space 

mission support information. 
The CCSDS Sub-Panel P1J is specifically chartered to investigate and recommend 

navigation data standards. P1J has a membership representing several international 

agencies. The main task of P1J is to develop preferred standards for the exchange of 

navigation data. The work of P1J is accomplished primarily at workshops, conducted at 

least twice a year, at facilities coordinated by the hosting member agency. The fiscal 

year 2002 workshops were conducted at the European Space Research and Technology 

Center (ESTEC), Holland, and the European Space Operations Center (ESOC), Germany, 

facilities of the European Space Agency (ESA), in October 2001 and April 2002, 

respectively. 

Accomplishments: The P1J navigation team completed a status review of the green book 

(technical report), titled "Navigation Definitions and Conventions", officially released for 

distribution in July 2001, but did not identify any updates; and a second CCSDS-wide 

review of the red book, titled "Orbit Data Messages" (ODM), which proposes a 

recommendation for space data systems standards for the exchange of spacecraft orbit 

information. This red book is expected to be approved by all CCSDS member agencies 

for promotion to blue book status (accepted preferred standard) by the end of calendar 

year 2002. P1J completed a set of questions pertaining to spacecraft identification 

specifications and tracking data interface requirements for CCSDS-wide responses, for 

work on future recommendations. ESOC and JPL conducted a successful test to verify 

proper exchange of the ephemeris (EPM) format, using the Ulysses spacecraft, in 

preparation for their support of the ROSETTA and Mars Express missions in 2003. P1J 

also prepared a poster to advertise the ODM development at the SpaceOps and 

International Telemetering Conferences. Both events are scheduled for October 2002. 

Future work of P1J will involve developing new technical reports and recommendations 

for navigation data exchange in the areas of tracking, attitude, time services, 

environmental models and astrodynamic constants, and proximity operations. 

[Technical contact: Felipe Flores-Amaya/572] 
7.0 Employee Development Activities 

7.1 New Employee Profiles 

During FY2002, the Flight Dynamics Analysis Branch welcomed three new employees. 

John Van Eepoel started in the Flight Dynamics Analysis Branch at Goddard Space 

Flight Center on September 30, 2002. He received a B.S. in Aerospace Engineering from 

the University of Maryland in 2000, going on to earn his S.M. from M.I.T. in the fall of 

2002. His work at M.I.T. focused on fault detection and repair for spacecraft, 

culminating in his thesis titled "Achieving Real-time Mode Estimation through Offline 

Compilation". His work at Goddard has taken him in a new direction. His current work 

includes supporting the JWST mission with analysis of the attitude control system, and 

molding the branch's attitude estimation system, MTASS, to the TRMM mission. In the 

future, John hopes to advance the attitude estimation systems used on the ground and onboard spacecraft, and in the process gain mission experience by implementing these 

systems for the control centers. 

Oscar Hsu has been working with the Flight Dynamics Analysis Branch since September 

9, 2002. He received his B.S. and M.S. in Aerospace Engineering from the University of 

Maryland, College Park and the title of his M.S. thesis was "Effect of Inlet Enthalpy on 

Liquid-Fueled Active Combustion Control." He is currently working on developing an 

18 degree-of-freedom nonlinear SIMULINK model for the ST-7 mission, which will be 

used as the validation model for the controller used in the disturbance rejection system. 

Bo Naasz joined the Flight Dynamics Analysis Branch on July 1, 2002. Bo received his 

Bachelor's and Master's Degrees in aerospace engineering from Virginia Tech, where he 

specialized in vehicle dynamics and control. Bo's M.S. thesis title was "Classical 

Element Feedback Control for Spacecraft Orbital Maneuvers." He is currently working 

on autonomous guidance, navigation, and control for spacecraft and spacecraft 

formations, including simulated and GPS hardware-in-the-loop studies. 

7.2 Professional Intern Program 

The Professional Intern Program (PIP) is a Goddard developmental program for entrylevel scientists, engineers, and administrative professionals. Within the FDAB it is an 

important development program for new engineers, designed to acquaint them with 

NASA and GSFC missions and operations, integrate them into the workforce as quickly 

as possible, and prepare them for more complex and responsible duties that they can 

perform with increasing independence. There are two levels of participation within the 

program. Employees entering with a BS degree begin in Level I and graduate into Level 

II following completion of Level I requirements and their first promotion. New 

employees entering Goddard with an MS degree begin in Level II. Required program 

activities include the establishment of a mentor relationship with an experienced staff 

member, various orientation activities, formal and on-the-job training, and completion of 
a PIP project, which the intern describes in a written report and oral presentations given 

in Levels I and II to a panel of evaluators. During the past year, four PIP projects were 

completed and presented. A description of each of these PIP projects (prepared by each 

intern) is given below. 

PIP Level I Project: The Effects of Deployment and Sun Acquisition on the ST-5 

Constellation (Melissa Fleck)

My PIP Level I presentation, entitled “The Effects of Deployment and Sun Acquisition 

on the ST5 Constellation,” was successfully completed on December 7, 2001. The Space 

Technology 5 (ST5) mission is composed of three 25 kg class spacecraft, each controlled 

by a single, offset maneuvering thruster. One of the early requirements of the mission 

was that at any point during the mission, over a two-hour span centered at apogee, the 

three spacecraft were to be separated by no less than 100 km and no more than 1000 km. 

Because the spacecraft have a small amount of fuel and no closed loop constellation 

maintenance, the major concern was how to get the spacecraft into a constellation that 

would limit spacecraft drift. My project was to determine the best method of spacecraft 

deployment and sun acquisition so that the constellation would not drift apart during the 

three-month mission lifetime. 

In order to determine the best deployment and sun acquisition method, I created a Matlab 

simulation that modeled the spacecraft deployment and sun acquisition maneuvers. I 

then ran a Monte Carlo simulation that varied the three unknown parameters: spacecraft 

launch date, right ascension of the ascending node, and true anomaly. In all of the 

simulations run, sun acquisition was performed immediately after spacecraft deployment. 

The original deployment plan for the ST5 spacecraft called for the launch vehicle to be 

despun, and the spacecraft deployed simultaneously from the launch vehicle 

approximately 120 degrees apart. From the results of the Monte Carlo simulation, it was 

clear that a constellation deployed in such a fashion would drift apart very quickly. An 

alternative deployment plan that was proposed had the launch vehicle rotate 120 degrees 

between spacecraft deployments such that each spacecraft was deployed along 

approximately the same vector. According to the results from my Monte Carlo 

simulation, this method of deployment yielded the best initial constellation formation. 

Upon the recommendation of the GNC team, the ST5 project adopted this method of 

spacecraft deployment. 

(Missie Fleck has been a Goddard employee since July 2001. She received her BS 

degree in Mechanical Engineering from the University of Maryland. She is currently 

pursuing an MS degree in Aerospace Engineering from the University of Maryland) 

PIP Level I Project: ST5 Orbit Definition and Constellation Control (Anne DeLion)

This project accomplished two goals for the ST5 project: it defined the ST5 mission orbit 

by determining the lowest initial perigee that would allow the spacecraft to maintain a 

minimum altitude over the mission lifetime, and it studied the possibility of maintaining a 
mission-required separation between the ST5 satellites by taking advantage of differences 

in the satellites’ drag profiles. 

Thermal concerns required ST5 to maintain at least a minimum altitude of 200-km 

throughout the entire mission. The initial perigee altitude had to be high enough to 

accommodate altitude fluctuations due to orbital perturbations. Satellite Took Kit (STK) 

analysis showed that the lowest allowable initial perigee was 270-km to maintain the 

minimum lifetime perigee height. 

ST5 originally required a separation that was greater than 100-km but less than 1000-km 

between any two satellites near apogee. Using the idea that a difference in drag force 

between two ST5 satellites could be used like a ∆V maneuver, this project preliminarily 

showed that separation between two ST5 spacecraft could be controlled using small 

attitude maneuvers to change the cross-sectional area of one spacecraft with respect to 

another. 

(Anne DeLion has been a full time Goddard employee since July 2001. Prior to that 

time, she was a coop student within the branch. She received her BS degree in Aerospace 

Engineering from Purdue University) 

PIP Level II Project: An Application for Sizing Momentum Wheels and Magnetic 

Torquer Bars on Spacecraft (Kristin Makovec)

During this year, I completed my PIP Level 2 project, which was titled “An Application 

for Sizing Momentum Wheels and Magnetic Torquer Bars on Spacecraft.” The project 

resulted in the development of a Matlab application which performs preliminary sizing 

for momentum wheels and magnetic torquer bars by examining a worst case condition of 

three-axis stabilization. The developed application is general for all non-spinning 

spacecraft and is adaptable for various missions. A GUI (Graphical User Interface) is 

available with the application to allow for a user-friendly method of data entry and is 

shown in Figure 7-1. 

The GUI contains options which vary parameters to allow for multiple runs or Monte 

Carlo simulations, as well as giving the user the option to choose which environmental 

torques are included. In addition, there are choices in the method of data entry for certain 

environmental torques, and the option to perform slew calculations. Default, save, and 

load capabilities are available on the GUI. The application runs through Matlab and calls 

upon a Simulink model to perform the calculations. The results are displayed graphically 

and in a table of maximum required values. A users manual is available with the 

application. 
Figure 7-1. Momentum Wheel Sizing Utility

The application was verified by comparing generated results with TRMM (Tropical 

Rainforest Measurement Mission) data. After verification, the application was used for 

the preliminary wheel sizing of SDO (Solar Dynamics Observatory) and wheel and 

torquer bar sizing of GPM (Global Precipitation Measurement). The results from the 

simulation runs were provided to the analysis teams for each of the missions. 

(Kristin Makovec has been a Goddard employee since August, 2001. She received her BS 

and MS degrees in Aerospace Engineering from Virginia Tech)

PIP Level II Project: A Two-Wheel Observing Mode for the Microwave Anisotropy 

Probe (Scott Starin)

The Microwave Anisotropy Probe (MAP) launched successfully in June, 2001, with a 

mission to map the Cosmic Microwave Background. A complete first mapping of the 

sky was completed after the first six months (the minimal mission length), and continuing 

observations improve upon those measurements. 
MAP was designed under a philosophy of “selective redundancy” to keep mass and 

hardware costs relatively low. As a result, MAP has only three reaction wheels—the 

minimum required to complete its nominal mission. In the event of a wheel failure, 

however, the mission should be completed in a degraded mission mode. My PIP project 

was to explore options that would allow the mission to continue with as little impact to 

science observations as possible. 

Before I started work, Dr. Tom Flatley had suggested that MAP, if it suffered a wheel 

failure, could be given an angular momentum bias using thrusters and that the two active 

wheels could be used to stabilize and control the natural motion. The resulting 

combination of nutation and spin would then provide a useful, though degraded, 

observing scan pattern for the fixed microwave instruments. 

My main contribution was to design a stable algorithm to control nutation angle by 

calculating wheel commands based on sensor inputs already available to the Attitude 

Control software; the design used existing software table structures to allow for on-orbit 

adaptability. I also aided in the design and development of algorithms to be used during 

the critical thruster maneuvers of the early operations period and during the establishment 

of a momentum bias for two-wheel observations. Software development and testing was 

done in two phases so that mission-critical elements would be available before launch; 

the final software products are ready to be uploaded in the event of a MAP wheel failure. 

(Scott Starin has been a Goddard employee since October, 2000. He received his BS 

degrees in Aerospace Engineering, Physics, and Multidisciplinary Studies from N.C. 

State and his MS degree in Aerospace Engineering from Ohio State.) 

7.3 Cooperative Education Program 

The Cooperative Education Program is an important link in the educational process that 

integrates college level academic study with full-time meaningful work experience. This 

is achieved through a working agreement between GSFC and a number of educational 

institutions. This agreement allows the students, through study and work experience, to 

enhance their academic knowledge, personal development, and professional preparation. 

Additionally, Co-op employees earn income that is based on the level of education and 

work experience they have attained. The FDAB fully supports the Goddard Co-op 

program and many of its full time employees were former Co-ops. In FY2002, three Coops worked in the branch. Given below are descriptions of their work experiences. 

Aaron Cooper (University of Minnesota)

During this past summer I worked on a variety of different projects. The first was 

deriving the equations for the gravitational force of center bodies of different geometries 

on a point mass. Dave Mangus and John Lynch oversaw the work I did in this area. The 

second project I worked on was a study to determine whether it was possible to obtain the 

mass of the fuel remaining on the Tropical Rainforest Measurement Mission (TRMM) 
spacecraft. The idea here was to use flight data and the dynamics of the spacecraft to 

obtain an equation that could solve for a single variable describing the geometry of the 

fuel. Using this variable we would subsequently determine the mass of the fuel on-board. 

This method turned out to be too sensitive to changes in the value used for the force 

being exerted by each thruster. My mentor for this project was Stephen Andrews. The 

third thing I worked on was assisting Paul Mason in some analysis of the induced torque 

caused by the migration of the Center of Mass on the Solar Dynamics Observatory (SDO) 

spacecraft, and how different thruster configurations could reduce the magnitude of the 

induced torques. Finally, I spent a good chunk of time this summer working on the 

testing and verification of a PC version of the Interactive Controls Analysis (INCA) 

software developed by John Downing. I worked on comparing the results of INCA with 

those produced by MATLAB for a number of simple transfer functions. I also built up 

the TRMM delta-V control system that was previously developed by Stephen Andrews, 

and analyzed it in INCA and MATLAB. These results were also compared with the 

results of Steve’s previous analysis, which was done on the VAX version of INCA. 

Stephen Andrews was my mentor on this project and I also worked with John Downing 

on the software issues surrounding INCA. 

Rivers Lamb (Virginia Polytechnic Institute and State University)

The summer of 2002 was my fourth co-op tour with the Flight Dynamics Analysis 

Branch. My primary assignment during this time was to learn about geosynchronous 

mission design, specifically in support of the Solar Dynamics Observatory (SDO) 

spacecraft. This project was conducted under the guidance of Bob DeFazio, and included 

learning both ascent optimization and engine modeling programs. The ultimate result of 

this work was a possible mission profile for the ascent maneuvers to the mission orbit 

based on a launch in August 2007. This work included such considerations as ground 

station coverage and stationkeeping requirements. I also supported a trade study 

investigating the options for the ascending node of the orbit based on obscuration of the 

high gain antennas. Finally, this work led to support of the geostationary GOES-R 

mission during the time it was in the Integrated Mission Design Center (IMDC). 

Working alongside Bob DeFazio and Charlie Petruzzo, this included the selection of the 

geosynchronous transfer orbit as well as an analysis of the requirements for station 

changes. 

Leigh Janes (Purdue University)

Over my ten week tour this past summer, I worked on two different projects. The first 

was to begin constructing a toolbox in Matlab for trajectory design and mission analysis 

tools. Steven Hughes was my mentor for this project. For the toolbox I took existing 

Fortran code and interfaced it with Matlab using mex functions established in Matlab. I 

worked on Fortran mexing for converting Keplerian elements to circular elements and for 

a two-body propagator. My second task was to learn the process used for designing 

geosynchronous missions. For this project I worked with Robert DeFazio. In order to 

learn this process I looked at the orbit planning for Solar Dynamics Observatory (SDO). 
I modeled maneuvers for the spacecraft, from launch vehicle separation to orbit 

placement. I also planned station keeping maneuvers to allow the spacecraft to remain 

within the drift box. 

7.4 Professional Development Program 

http://www.hq.nasa.gov/office/codef/codeft/pdp/

The Professional Development Program (PDP) is designed to broaden the participants' 

knowledge and understanding of the Agency and encourage the development of their 

leadership skills through a combination of expanded work experiences and formal 

training. Participants in this program, who are competitively selected at each Center and 

then across the Agency, identify developmental work assignments away from their home 

Center. Benefits include learning new job skills, being exposed to new areas of NASA 

and senior NASA officials, and participating in a variety of developmental activities. 

Dr. James O’Donnell, a senior Aerospace Engineer in the Flight Dynamics Analysis 

Branch, is participating in the PDP from August, 2002, through July, 2003. His primary 

work assignment is in the Office of the Chief Engineer at NASA Headquarters, and he is 

hoping to arrange a collateral work assignment at either the European Space Research 

and Technology Centre (ESTEC) or the Jet Propulsion Laboratory (JPL). 

[Contact: James O’Donnell] 

7.5 In-house Employee Classes 

Technology transfer is a key element in sustaining a well-trained work force. The FDAB 

encourages all engineers to take advantage of the many training opportunities that 

Goddard has to offer. In augmentation with the Goddard training, the FDAB is 

continuing to offer courses that are developed in house and taught by the senior 

engineers. These courses are targeted for the new engineer, yet other senior engineers 

who wish to expand their field are encouraged to attend. Three of the more prominent 

courses are listed below: 

The “Attitude Control Systems for Non-ACS Engineers” course was designed as an 

introduction for the new engineer who has only a scholarly background in control 

systems. The course expanded on how to apply their knowledge to the real world design 

of attitude control systems. All classes, with guest lecturers and hardware demonstrations, 

were video taped for future use. 

Dr. Bar-Itzhack, who was under a NASA grant from the Technion-Israel Institute of 

Technology, presented the “Attitude Estimation and Kalman Filtering” course. The 

course covered the basics of attitude determination, Kalman filter design and then the 

combination of the two for attitude estimation. 
The “Kane’s Method for Solving Multi-body Dynamics” course presented a method of 

solving multi-body dynamic problems in a cleaner way than the more conventional 

methods. The course covered Kane’s basic theory as well as how to use the AUTOLEV 

software. The participants were assigned homework that involved setting up the problems 

on paper and using AUTOLEV to produce dynamic equations of motion. 

[Technical contact: Dave Mangus] 

7.6 Attitude Control System Handbook 

The FDAB has drafted a document from which an ACS subsystem may be 

conceptualized, designed, validated, and supported throughout all mission phases. It 

includes both programmatic and technical guidelines to be used throughout the life cycle 

of such a development, and is based upon a combination of in-house experience, NASA 

input, recent red-team guidelines, and outside resources. 

This document is not intended to be a control systems design and development class. It is 

assumed the subsystem engineer using this document is well versed in control theory and 

in the use of simulation. For example, the use of Bode or Nichols plots is not discussed, 

however, design guidelines pertaining to criteria to apply to these tools is covered. The 

Swales Corporation was contracted to work with the senior ACS engineers within the 

FDAB, and capture their philosophies within the document. 

[Technical contact: Dave Mangus] 

7.7 TableSat 

TableSat is an interactive, single axis hardware simulator that physically demonstrates the 

reaction of attitude control systems. Using a simple radio communications link, the table 

is controlled by a laptop computer. A gyro package and set of fans are mounted on a 1.5 

foot diameter table that is suspended on a centered pin. Also on the table are coarse sun 

sensors, a receiver, transmitter and batteries. The laptop, containing SIMULINK, is 

outfitted with a receiver and transmitter set. This system allows the user to ‘fly’ the table. 

Any control system can be modeled and modified in SIMULINK resulting in a real-time 

reaction of the table. The table was developed as a demonstration tool for the “Attitude 

Control Systems for Non-ACS Engineers” Course. Due to the very positive feedback 

from the class participants, the table is to be expanded to demonstrate areas including 

flexible body dynamics. A storyboard will also be developed to create interest in attitude 

control systems at conferences and universities. 
Figure 7-2. TableSat Side View

Figure 7-3. TableSat Top View

[Technical contact: Dave Mangus] 
8.0 Outreach Activities 

8.1 SAMPEX University Operations

The University of Maryland Aerospace Engineering Department completed its third 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. As an additional benefit, this 

program helps give students valuable experience for future employment. One of the new 

hires in the branch this year (Oscar Hsu) was a member of the student team at Maryland 

that supported SAMPEX flight dynamics operations. 

[Technical Contact: Tom Stengle] 

8.2 PREST Program 

During FY02, 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 completed his work in residence at the GSFC, which focused on 

formation flying control approaches. 

[Technical Contact: Tom Stengle] 

8.3 Graduate Student Research Program (GSRP) 

The Graduate Student Researchers program (GSRP) is a program that brings quality 

graduate students into collaborative research arrangements with NASA scientists and 

engineers. Entrance into the program is highly competitive and results typically in a 

three-year commitment by Goddard and the selected students to perform research of 

mutual interest. In FY02, a total of seven students were sponsored by the FDAB. These 

students and their research abstracts are listed below: 

“Feasibility of Atmospheric Penetration for Satellite Formation Flying Experiment.” 

Researcher: Joseph Schultz, University of Maryland (third year).

This research develops optimal control and guidance methods needed to maneuver a 

formation of satellites through penetrations into the upper atmosphere. The research has 

direct application to NASA’s Geospace Electrodynamic Connections (GEC) Mission 
where its formation of satellites will be dipping into the upper atmosphere for electric and 

magnetic field measurements. Of large concern is that the flexibility of large booms on 

the satellite may cause it to become unstable during the atmospheric portion of flight. 

This research will determine guidance and control laws that optimize the fuel savings of 

formation maneuvers during an aeroassist while also preventing satellite instabilities. 

“ Adaptive Satellite Attitude Control.” Researcher: Kevin Walchco, University of 

Florida (second year).

Attitude control of small spacecraft is a particularly important element in any of 

Goddard’s existing and future missions. Typically designers use classical control due to 

its simplicity and ease of implementation. However, classical control cannot provide 

robustness in the presence of disturbances and uncertainties such as solar snap, fuel slosh, 

and vibrations. One of the most promising robust methods is sliding mode control. In 

this work, an adaptive sliding mode controller, which utilizes Kalman filtering to estimate 

dynamic and uncertain parameters in a satellite, is used to improve performance and 

robustness of the attitude control. 

“Investigation of Libration Orbits in the Earth-Moon System.” Researcher: Raquel 

Jarabek, University of Maryland (second year).

Additional uses for libration point orbits are currently being discovered and more 

satellites will make use of libration points in the future. This creates a need for more 

automated software to plan trajectories to libration points. To meet this need, an 

automated method will be developed to find the optimal path for spacecraft trajectories to 

Sun-Earth libration points. Three-body motion will be simulated through computer 

programming and the applicable optimal control problem will be formulated using the 

Hamilton-Jacobi-Bellman equations. The solution to the optimal control problem will be 

compared to manifold theory to ensure that the optimal path is found. 

“Decentralized Control of Distributed Satellite Networks.” Researcher: Belanger, 

UCLA (second year).

Decentralized control of satellite clusters in the context of limited information exchange 

is studied. Optimal control laws are derived for various communications paradigms, 

using a combination of static team theory and dynamics programming techniques. These 

control laws are linear in the local measurement history and the a priori state estimate. 

One communication paradigm discussed appears to represent the minimal information 

exchange necessary to sustain an affine control law. These communication paradigms 

and associated control laws are applied and studied in the context of a high fidelity 

satellite cluster simulation. 

“Trajectory Planning for Coordinating Satellites Using Command Generation.” 

Researcher: Erika Biediger, Georgia Institute of Technology (first year).

The proposed work seeks to enhance the current state-of-the-art formation flying 

algorithms by utilizing Command Generation, specifically input shaping. Input Shaping 

is a relatively simple technique that will allow the flexible spacecraft to move without 

inducing residual vibration, limit transient deflection, and move in a fuel-efficient 
manner. The proposed work will develop and employ a real-time pulse generator to 

address the vibration issues onboard each satellite. The real-time pulse generator will be 

integrated with the current formation flying algorithms for enhanced performance. 

“Synthesis of Attitude Determination Algorithms on a PiVoT GPS Receiver.” 

Researcher: Jared Madsen, University of Texas at Austin (first year)

This research will combine the carrier wave and signal to noise ratio methods of attitude 

determination on a PiVoT GPS receiver. To accomplish this synthesis, the carrier wave 

method of attitude determination must be altered to allow for solutions with canted 

antennas. A filter will be designed to integrate information from both methods to achieve 

the best possible solutions. This system will provide solutions even when the integer 

ambiguity resolution process is unfinished or has failed. It is also anticipated that this 

new system will have improved integer ambiguity resolution speed. This new system 

will be accurate and robust. 

[Technical Contact: Tom Stengle] 

8.4 Mission Design & Navigation Workshop 

http://FDAB1.GSFC.NASA.GOV/

The FDAB held a Mission Design and Navigation Workshop on January 9, 2002. It was 

a very successful event, drawing over 90 registered participants. Branch head Tom 

Stengle and senior engineer David Folta gave the presentation which covered the FDAB 

history and business plan, information about and recent accomplishments in each flight 

regime (LEO, MEO/HEO, GEO, lunar, libration, and interplanetary), and branch 

technology development efforts. In addition to the presentation, branch members gave 

hands-on demonstrations of FDAB tools and capabilities. FDAB hopes to sponsor a 

similar event each year, both to update the mission design and navigation information and 

also to highlight the branch's work in the attitude determination and control areas. 

For a copy of the presentation and to view the posters that were on display, visit the 

FDAB website listed above and click on the Workshop link. 

[Technical contact: Lauri Newman] 

 
Appendix A – Goddard and NASA Awards

Team Awards 

NASA Software of the Year Award (SOYA); Code 572/583 and ai-solutions inc. won 

runner up of the prestigious software of the year competition for "AutoCon - 

Autonomous Maneuver Control Flight Software". The SOYA is held yearly and involved 

8 competitors from each NASA center. 

NASA Group Achievement Awards for the EO-1 Project Team and for the EO-1 

Formation Flying Team 

NASA Group Achievement Award – Microwave Anisotropy Probe (MAP) Guidance, 

Navigation and Control (GN&C) Team 

NASA Group Achievement Award -- Center of Excellence MAP team 

Goddard Center of Excellence Award- Triana Project 

Special Act Award for the Sun-Earth Connection Roadmap Missions Definition Team 

Special Act Award for the NASA GSFC X-43 Mishap Investigation Team 

Individual Goddard/NASA level Awards

NASA Exceptional Achievement Award- Aprille Ericsson 
Appendix B - University Grants 

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

FY02: 

1. GRANT NAG5-11839 with Purdue University titled “Application of Dynamical 

System Theory, Control Methods, and Optimization Strategies to Trajectory 

Design and Mission Analysis Involving Formation Flying at Libration Points for 

GFSC Missions” This grant was established to investigate the application of 

dynamical systems to formation flying missions at the Sun-Earth libration points. 

Investigation involved study of DS applications to traditional non-linear and 

linear control methods and the use of the natural dynamics regions. The results 

are being used to design the trajectories of Constellation-X, FKSI, and MAXIM. 

[Technical Contact: David.C.Folta] 

2. GRANT NAG5-3587 with Purdue University titled “Dynamical System Theory, 

Numerical Methods, Optimization Strategies, and their Applications to Trajectory 

Design and Mission Analysis Involving Lissajous and Halo Orbits at the GFSC 

GNCC” This grant was established to continue Dynamical Systems (DS) 

applications. Investigation involved further study of DS applications of manifolds, 

improvements to the Generator utility to include additional information on lunar 

gravity assist and targeting schemes, investigation of orbit bifurcation of 

manifolds, and investigation of the use of combinatorics for trajectory design and 

optimization. The result of this grant was used to design the trajectories of JWST, 

Triana, Constellation-X, FKSI, and others 

[Technical Contact: David.C.Folta] 

3. GRANT NAG5-12228 with University of Illinois at Urbana Champaign titled 

“Advanced Methods For Optimal Trajectory Design” Investigation and 

development of genetic algorithms and primer vector theory applications to 

optimization of orbit deign. The UIUC Computational Astrodynamic Research 

Laboratory (CARL) investigated orbital state transition matrix methods. Several 

established techniques have been identified. The varying strengths and 

weaknesses of the methods were established for Gim and Alfriend, Goodyear, 

Danby, and Battin, Matrix methods for the approximate solution of differential 

equations are applied to the development of general perturbations in rectangular 

coordinates. A tutorial session to demonstrate how to use the trajectory analysis 

software that was co-developed with Spectrum Astro was organized. Sample 

missions are being sought out to develop into example scenarios in the software. 

Hopefully this will facilitate the use of the software. The result of this grant was 

used to minimize the DV and Fuel costs on SDO, GPM, and Leonardo. 

[Technical Contact: David.C.Folta] 
4. GRANT NAG5-12119 with Virginia Tech titled “Time-Optimal Control for 

Formation Establishment and Maneuvering” Investigation of time optimal transfer 

for Low Earth orbits. Investigation focused on time-constrained problems. The 

results of this grant were used on SDO and LISA. 

[Technical Contact: David.C.Folta] 

5. GRANT NAG5-10384 with University of Cincinnati titled “Dynamics and 

Control of Distributed Spacecraft” Investigation of the use of differential solar 

radiation pressure effects for formation control and establishment. It was applied 

to the transfer and formation maintenance of Low Earth circular and elliptical 

orbits. Investigation focused on MMS mission like problems. The results of this 

grant were used on MMS and Leonardo. 

[Technical Contact: David.C.Folta] 

6. 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] 

7. 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] 

8. 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] 
9. 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] 

10. GRANT NAG5-12179 with State University of New York Buffalo and Dr. John 

Crassidis of the Department of Mechanical and Aerospace Engineering titled 

“Development of Automated Alignment and Calibration Algorithms”. This grant 

is studying sensor calibration algorithms which would be suitable for a spacecraft 

onboard computer. 

[Technical Contact: Richard Harman] 

11. GRANT NAG5-11919 with Cornell University and Dr. Mark Psiaki of the 

Sibley School of Mechanical & Aerospace Engineering titled “Euler DynamicsBased Estimation Algorithms for Spacecraft Attitude and Rate Determination”. 

This grant looks at accurately estimating spacecraft attitude and rate when no 

gyros are available. 

[Technical Contact: Richard Harman] 

12. 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] 

13. 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] 
14. GRANT NAG5-12228 with the University of Illinois Urbana Champaign 

Aeronautical and Astronautical Engineering Dept. Entitled “Multi-Function 

Stochastic Optimization for Space Mission Design”. Under this grant, UIUC is 

developing optimization algorithms for use in matlab. These capabilities will be 

included into the TrajOpt architecture to be used with existing mission design ools 

such as FreeFlyer, Astrogator, or GMAT. 

[Technical Contact: Steven Hughes] 

15. GRANT NAG5-12312 with the University of California San Diego Department 

of Mathematics Entitled “ Software for Constrained Optimization”. Under this 

grant, the UCSC is developing constrained optimization approaches to be 

included in the TrajOpt software. The TrajOpt software will enable optimal 

mission design in FreeFlyer, Astrogator, and GMAT. 

[Technical Conact: Steven Hughes] 
Appendix C – Conferences and Papers 

Given below are abstracts from professional papers and technical presentations that were 

prepared and delivered in FY02 by branch members. 

JOURNAL ARTICLES 

IEEE Transactions on Automatic Control, submitted May 2002

“ A Coupled Nonlinear Spacecraft Attitude Controller and Observer with an Unknown 

Constant Gyro Bias and Gyro Noise”, Thienel (GSFC), Sanner (University of Maryland) 

Abstract: A nonlinear control scheme for attitude control of a spacecraft is combined with 

a nonlinear gyro bias observer for the case of constant gyro bias. A persistency of 

excitation analysis shows the observer bias estimates converge to the true bias values 

exponentially fast. The resulting coupled, closed loop dynamics are proven by a 

Lyapunov analysis to be globally stable, with asymptotically perfect tracking. The 

analysis is extended to consider the effects of noise in addition to the gyro bias. A 

simulation of the proposed observer-controller design is given for a rigid spacecraft 

tracking a specified, time-varying attitude sequence to illustrate the theoretical claims. 

Journal of Guidance, Control, and Dynamics (Published January-February 2002)

 “Optimal Fusion of a Given Quaternion with Vector Measurements”, Bar-Itzhack 

(Technion), Harman (GSFC). 

ABSTRACT: Several satellites that use devices called autonomous star trackers (ASTs) 

are presently operating. These devices put out the satellites’ attitude in the form of a 

quaternion. Usually the satellites also carry other attitude measuring devices, such as sun 

sensors and magnetometers that measure vectors in body coordinates. Although 

the accuracy of the AST surpasses that of the other sensors, due to the synergistic effect 

of sensor fusion, it is still desirable to incorporate the measurements of the less accurate 

sensors in the attitude determination process. The question is, then, how to blend 

optimally the AST-generated quaternion with vector measurements. This problem rose, 

for example, in the design of the attitude determination algorithm of the Microwave 

Anisotropy Probe (MAP) satellite, which was launched on 30 June 2001. MAP has two 

ASTs and two sun sensors, one of which is more accurate than the other. Although we 

have two quaternion-generating devices and two vector-measuring sensors, we consider 

here only one of each. The extension of the solution, proposed here, to multiple devices 

and multiple sensors, is immediate. 

We note that the quaternion is a four-element vector that yields the whole attitude, 

whereas vector-measuring sensors yield three dimensional vectors each containing only 

partial information on the attitude. Therefore, we cannot cast the problem of optimal 
attitude determination in the form of Wahba’s problem.1 That is, we cannot blend 

quaternions with vector measurements using known algorithms. 

If we have more than one simultaneous vector measurement, we can use the vector 

measurements to, first, find attitude expressed in quaternion form and, then, blend this 

quaternion with the given one. However, when we have only one vector measurement, 

this is not possible. Therefore, we need an algorithm that can blend the given quaternion 

even with one vector measurement. 

The algorithm presented here consists of two steps. In the first step, the quaternion is 

converted into a pair of pseudo vector measurements that express the attitude, and then, 

in the second step, these pseudo vector measurements, together with the given vector 

measurement (or measurements), are used as inputs to the q-method algorithm,2 which 

generates the optimal quaternion. The resultant quaternion is optimal in the sense that it is 

the best . t, in the least squares sense, to all of the vectors. 

Journal of Guidance, Control, and Dynamics (Published September-October 2002)

“In-Space Calibration of a Skewed Gyro Quadruplet”, Bar-Itzhack (Technion), Harman 

(GSFC). 

ABSTRACT: A new approach to gyro calibration, where the spacecraft dynamics 

equation, attitude measurements, and the gyro outputs are used in a pseudo linear Kalman 

filter that estimates the calibration parameters. Also an algorithm is presented for 

calibrating a skewed quadruplet rather than the customary triad gyro-set aligned along the 

body coordinate axes. In particular, a new misalignment error model is derived for this 

case. The new calibration algorithm is applied to the EOS-AQUA satellite gyros. The 

effectiveness of the new algorithm is demonstrated through simulations. 

CONFERENCES 

16th International Symposium on Space Flight Dynamics Pasadena, CA, December 3-

7, 2001

"Application Of Monte Carlo Analyses For The Microwave Anisotropy Probe (Map) 

Mission", Mesarch(GSFC), Rohrbaugh, Schiff (aisolutions). 

ABSTRACT: The Microwave Anisotropy Probe (MAP) is the third launch in the 

National Aeronautics and Space Administration’s (NASA’s) a Medium Class Explorers 

(MIDEX) program. MAP will measure, in greater detail, the cosmic microwave 

background radiation from an orbit about the Sun-Earth/Moon Lagrangian point. 

Maneuvers will be required to transition MAP from its initial highly elliptical orbit to a 

lunar encounter which will provide the remaining energy to send MAP out to a lissajous 

orbit about L2. Monte Carlo analysis methods were used to evaluate the potential 

maneuver error sources and determine their effect of the fixed MAP propellant budget. 
This paper will discuss the results of the analyses on three separate phases of the MAP 

mission recovering from launch vehicle errors, responding to phasing loop maneuver 

errors, and evaluating the effect of maneuver execution errors and orbit determination 

errors on stationkeeping maneuvers at L2. 

"Trajectory Design for the Microwave Anisotropy Probe (MAP)", Newman (GSFC), 

Rohrbaugh (a.i. Solutions) 

ABSTRACT: The Microwave Anisotropy Probe (MAP) mission orbit is a Lissajous 

orbit about the L2 Sun-Earth Lagrange point. The trajectory design for MAP is complex, 

having many requirements that must be met including shadow avoidance, sun angle 

constraints, lissajous size and shape characteristics, and limited Delta-V budget. The 

results of the trajectory design analysis are presented. The results show relationships 

between the requirements and the resulting trajectory in an effort to establish patterns and 

repeatability in defining satisfactory orbits. The paper discusses the preliminary tradeoffs to establish a baseline trajectory, analysis to establish the nominal daily trajectory, 

and the launch window determination. 

“Designing Phase 2 for the Double-Lunar Swingby of the Magnetospheric Multiscale 

Mission”, Edery, Schiff (a.i. solutions) 

ABSTRACT The double-lunar swingby (DLS) of the Magnetospheric MultiScale 

(MMS) mission is required, within a tight delta-V budget of 90 m/s, to change 

significantly the orbital elements of an initial orbit: to increase dramatically its inclination 

(by 570), decrease significantly its eccentricity (from 0.96 to 0.66) and keep its 

semimajor axis approximately constant. We obtain double-lunar swingbys that 

accomplish this task with a remarkably low delta-V of 50.3 m/s. Our approach is semianalytical: a derived analytical expression is used to determine that a non-symmetrical 

targeting scheme is best suited for this problem 

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

 “A Coupled Nonlinear Spacecraft Attitude Controller/Observer with an Unknown 

Constant Gyro Bias”, Thienel (GSFC), Sanner (University of Maryland) 

ABSTRACT: A nonlinear control scheme for attitude control of a spacecraft is combined 

with a nonlinear gyro bias observer for the case of constant gyro bias. A persistency of 

excitation analysis shows the observer bias estimates converge to the true bias values 

exponentially fast. The resulting coupled, closed loop dynamics are proven by a 

Lyapunov analysis to be globally stable, with asymptotically perfect tracking. The 

analysis is extended to consider the effects of noise in addition to the gyro bias. A 

simulation of the proposed observer-controller design is given for a rigid spacecraft 

tracking a specified, time-varying attitude sequence to illustrate the theoretical claims. 

25th Annual AAS Guidance and Control Conference, Breckenridge, CO, February, 

2002
“Preliminary Results of the GPS Flight Experiment on the High Earth Orbit AMSATOSCAR 40 Spacecraft”, Moreau, Bauer, Carpenter, E. Davis (GSFC), G. Davis 

(Emergent Space Technologies), Jackson (Orbital Sciences Corporation)

ABSTRACT: The GPS flight experiment on the High Earth Orbit (HEO) AMSATOSCAR 40 (AO-40) spacecraft was activated for a period of approximately six weeks 

between 25 September and 2 November, 2001, and the initial results have exciting 

implications for using GPS as a low-cost orbit determination sensor for future HEO 

missions. AO-40, an amateur radio satellite launched November 16, 2000, is currently in 

a low inclination, 1000 by 58,800 km altitude orbit. Although the GPS receiver was not 

initialized in any way, it regularly returned GPS observations from points all around the 

orbit. Raw signal to noise levels as high as 12 AMUs (Trimble Amplitude Measurement 

Units) or approximately 48 dB Hz have been recorded at apogee, when the spacecraft 

was close to 60,000 km in altitude. On several occasions when the receiver was below 

the GPS constellation (below 20,000 km altitude), observations were reported for GPS 

satellites tracked through side lobe transmissions. Although the receiver has not returned 

any point solutions, there has been at least one occasion when four satellites were tracked 

simultaneously, and this short arc of data was used to compute point solutions after the 

fact. These results are encouraging, especially considering the spacecraft is currently in a 

spin-stabilized attitude mode that narrows the effective field of view of the receiving 

antennas and adversely affects GPS tracking. Already AO-40 has demonstrated the 

feasibility of recording GPS observations in HEO using an unaided receiver. 

Furthermore, it is providing important information about the characteristics of GPS 

signals received by a spacecraft in a HEO, which has long been of interest to many in the 

GPS community. Based on the data returned so far, the tracking performance is expected 

to improve when the spacecraft is transitioned to a three axis stabilized, nadir pointing 

attitude in Summer, 2002. 

“Microwave Anisotropy Probe (MAP) Launch and Early Operations”, O'Donnell, 

Andrews, Starin, Ward (GSFC) 

ABSTRACT: The Microwave Anisotropy Probe (MAP), a follow-on to the Differential 

Microwave Radiometer (DMR) instrument on the Cosmic Background Explorer (COBE), 

was launched from the Kennedy Space Center at 19:46:46 UTC on June 30, 2001. The 

powered flight and separation from the Delta II appeared to go as designed, with the 

launch placing MAP well within 1σ launch dispersion and with less than 7 Nms of tip-off 

momentum. Because of this relatively low momentum, MAP was able to acquire the sun 

within only 15 minutes with a battery state of charge of 94%. After MAP’s successful 

launch, a six week period of in-orbit checkout and orbit maneuvers followed. The dual 

purpose of the in-orbit checkout period was to validate the correct performance of all of 

MAP’s systems and, from the attitude control system (ACS) point of view, to calibrate 

the performance of the spacecraft ACS sensors and actuators to maximize system 

performance. In addition to the checkout activities performed by the MAP team, the other 

critical activity taking place during the first six weeks after launch were a series of orbit 
maneuvers necessary to get the spacecraft from its launch orbit out to its desired orbit 

about L2, the second Earth-Sun Lagrange point. 

As MAP continues its standard operations, its ACS design is meeting all of its 

requirements to successfully complete the mission. This paper will describe the launch 

and early operations summarized above in greater detail, and show the performance of 

the attitude control and attitude determination system versus its requirements. 

Additionally, some of the unexpected events that occurred during this period will be 

discussed, including two events which dropped the spacecraft into its Safehold Mode and 

the presence of an “anomalous force” observed during each of the perigee orbit 

maneuvers that had the potential to cause these critical maneuvers to be prematurely 

aborted. 

42nd Israel Annual Conference on Aerospace Sciences, February 20-21, 2002, Tel-Aviv 

-Haifa, Israel

“State-Dependent Pseudo-Linear Filter for Spacecraft Attitude and Rate Estimation”, 

Bar-Itzhack (Technion) ,Harman (GSFC) ,Choukroun (Technion). 

ABSTRACT: This paper presents the development and performance of a special 

algorithm for estimating the attitude and angular rate of a spacecraft. The algorithm is a 

pseudo-linear Kalman filter, which is an ordinary linear Kalman filter that operates on a 

linear model whose matrices are current state estimate dependent. The nonlinear 

rotational dynamics equation of the spacecraft is presented in the state space as a statedependent linear system. Two types of measurements are considered. One type is a 

measurement of the quaternion of rotation, which is obtained from a newly introduced 

star tracker based apparatus. The other type of measurement is that of vectors, which 

permits the use of a variety of vector measuring sensors like sun sensors and 

magnetometers. Several measurement models are presented, a linear model for the case 

where the measured quantity is the quaternion, and two pseudo-linear measurement 

models when the measured quantities are vectors. In the latter situation, one model is for 

the case when the attitude is represented by a quaternion and the other is for the case 

where it is represented by a direction cosine matrix. The state-dependent pseudo linear 

filter is tested using simulated spacecraft rotations. The efficiency of each case is 

demonstrated through simulation. 

“Optimal Fusion of a Given Quaternion with Vector-Measurements”, Bar-Itzhack 

(Technion), Harman (GSFC). 

ABSTRACT: Several satellites that use devices called Autonomous Star Trackers 

(ASTs) are presently operating. These devices put out the satellites’ attitude in the form 

of a quaternion. Usually the satellites also carry other attitude measuring devices, like sun 

sensors and magnetometers that measure vectors in body coordinates. Although the 

accuracy of the AST surpasses that of the other sensors, due to the synergistic effect of 

sensor fusion, it is still desirable to incorporate the measurements of the less accurate 

sensors in the attitude determination process. The question is then, how to optimally 
blend the AST generated quaternion with vector-measurements. This problem rose, for 

example, in the design of the attitude determination algorithm of the Microwave 

Anisotropy Probe (MAP) satellite, which was launched on June 30, 2001. MAP has two 

ASTs and two sun sensors, one of which is more accurate than the other. Although we 

have two quaternion-generating devices and two vector-measuring sensors, we consider 

here only one of each. The extension of the solution, proposed here, to multiple devices 

and multiple sensors, is immediate. 

We note that the quaternion is a four-element vector that yields the whole attitude 

whereas vector-measuring sensors yield three-dimensional vectors each containing only 

partial information on the attitude. Therefore we cannot cast the problem of optimal 

attitude determination in the form of Wahba’s Problem [1]. That is, we cannot blend 

quaternions with vector-measurements using known algorithms. 

If we have more than one simultaneous vector-measurement we can use the vectormeasurements to, first, find attitude expressed in quaternion form and then blend this 

quaternion with the given one. However, when we have only one vector-measurement, 

this is not possible. Therefore we need an algorithm that can blend the given quaternion 

even with one vector-measurement. 

The algorithm presented here consists of two steps. In the first step the quaternion is 

converted into a pair of pseudo vector-measurements that express the attitude, and then, 

in the second step, these pseudo vector-measurements, together with the given vectormeasurement (or measurements), are used as inputs to the q-Method algorithm [2], which 

generates the optimal quaternion. The resultant quaternion is optimal in the sense that it is 

the best fit, in the least squares sense, to all the vectors. 

STK Users' Conference, Washington, DC, June 3-4, 2002

“MAP Trajectory Design Using STK”, Woodard (GSFC) 

ABSTRACT: The NASA Goddard Space Flight Center (GSFC) developed the 

Microwave Anisotropy Probe (MAP) mission to produce an accurate full-sky map of the 

cosmic microwave background temperature fluctuations – anisotropy. The MAP mission 

orbit is a Lissajous orbit about the Sun-Earth L2 Lagrange Point. The mission duration is 

approximately 27 months with 3 to 4 months of transfer time to the final mission orbit 

about L2. The mission is exceptional from a trajectory perspective because it is the first 

mission to orbit the Sun-Earth L2 Lagrange Point. 

MAP was launched on June 30, 2001 into a highly elliptical Earth orbit with a 28.7°

inclination. In the following weeks, the GSFC trajectory design team planned and 

executed a sequence of phasing loops and performed a lunar gravity assist to reach the 

final mission orbit about the Sun-Earth L2 Lagrange point. MAP used a lunar swingby 

strategy since it reduced the fuel required to achieve the desired Lissajous orbit. 
The trajectory team used STK and its Astrogator module to assist in nearly all aspects of 

trajectory maneuver planning and execution. This included performing prelaunch 

trajectory analyses, computing delta-V budgets, verifying mission orbital constraints, 

providing Detailed Trajectory Objective (DTO) analyses to the launch vehicle team, 

defining launch windows, performing contingency analyses, and on-orbit maneuver 

planning and verification. To accomplish these tasks, the trajectory team defined several 

mission-specific Astrogator objects, and created several Matlab scripts that use the STK 

Connect interface to automate repetitive tasks. 

SPIE Astronomical Telescopes and Instrumentation Conference, 22-28 August 2002

“Formation Control for the MAXIM and MAXIM Pathfinder Missions”, Luquette, 

Leitner (GSFC), Gendreau, Sanner (University of Maryland) 

ABSTRACT: Over the next twenty years, a wave of change is occurring in the spacebased scientific remote sensing community. While the fundamental limits in the spatial 

and angular resolution achievable in spacecraft have been reached, based on today's 

technology, an expansive new technology base has appeared over the past decade in the 

area of Distributed Space Systems (DSS). A key subset of the DSS technology area is 

that which covers precision formation flying of space vehicles. Through precision 

formation flying, the baselines, previously defined by the largest monolithic structure 

which could fit in the largest launch vehicle fairing, are now virtually unlimited. Several 

missions including the Micro-Arcsecond X-ray Imaging Mission (MAXIM) [1], the 

associated MAXIM Pathfinder mission [2], and the Stellar Imager will drive the 

formation flying challenges to achieve unprecedented baselines for high resolution, 

extended-scene, interferometry in the ultraviolet and X-ray regimes. This paper focuses 

on establishing the feasibility for the formation control of the MAXIM Pathfinder 

mission (which will involve seven spacecraft flying in formation). The MAXIM mission 

(which involves over thirty spacecraft flying in formation) is also discussed. The Stellar 

Imager mission requirements are on the same order of those for MAXIM. This paper 

specifically addresses: (1) high-level science requirements for these missions and how 

they evolve into engineering requirements; (2) the formation control architecture devised 

for such missions; (3) the design of the formation control laws to maintain very high 

precision relative positions; and (4) the levels of fuel usage required in the duration of 

these missions. Specific preliminary results are presented for one formulation of the 

MAXIM pathfinder mission. 

Libration Point Orbits and Applications, Parador d'Aiguablava, Girona, Spain10 - 14 

June, 2002

 

“Orbit Determination Issues for Libration Point Orbits”, Beckman (GSFC) 

ABSTRACT: Libration point mission designers require knowledge of orbital accuracy 

for a variety of analyses including station keeping control strategies, transfer trajectory 
design, and formation and constellation control. Past publications have detailed orbit 

determination (OD) results from individual libration point missions. This paper collects 

both published and unpublished results from four previous libration point missions 

(ISEE-3, SOHO, ACE and MAP) supported by Goddard Space Flight Center’s Guidance, 

Navigation & Control Center. The results of those missions are presented along with OD 

issues specific to each mission. All past missions have been limited to ground based 

tracking through NASA ground sites using standard range and Doppler measurement 

types. Advanced technology is enabling other OD options including onboard navigation 

using onboard attitude sensors and the use of the Very Long Baseline Interferometry 

(VLBI) measurement Delta Differenced One-Way Range (DDOR). Both options 

potentially enable missions to reduce coherent dedicated tracking passes while 

maintaining orbital accuracy. With the increased projected loading of the DSN, missions 

must find alternatives to the standard OD scenario. 

“Libration Orbit Mission Design: Applications Of Numerical And Dynamical Methods”, 

Folta, Beckman (GSFC) 

ABSTRACT: Sun-Earth libration point orbits serve as excellent locations for scientific 

investigations. These orbits are often selected to minimize environmental disturbances 

and maximize observing efficiency. Trajectory design in support of libration orbits is 

ever more challenging as more complex missions are envisioned in the next decade. 

Trajectory design software must be further enabled to incorporate better understanding of 

the libration orbit solution space and thus improve the efficiency and expand the 

capabilities of current approaches. 

The Goddard Space Flight Center (GSFC) is currently supporting multiple libration 

missions. This end-to-end support consists of mission operations, trajectory design, and 

control. It also includes algorithm and software development of applications. The 

recently launched Microwave Anisotropy Probe (MAP) and upcoming Next Generation 

Space Telescope (NGST) and Constellation-X missions are examples of the use of 

improved numerical methods for attaining constrained orbits parameters and controlling 

their dynamical evolution at the collinear libration points. This paper presents a history of 

libration point missions, a brief description of the numerical and dynamical design 

techniques including software used, and a sample of GSFC future mission designs. 

AIAA Guidance, Navigation, and Controls Conf., Monterrey, CA, August 5-8, 2002

“MAP Attitude Control System Design and Flight Performance”, Andrews, O'Donnell 

(GSFC) 

ABSTRACT: The Microwave Anisotropy Probe (MAP) is a follow-on to the Differential 

Microwave Radiometer (DMR) instrument on the Cosmic Background Explorer (COBE) 

spacecraft. To make a full-sky map of cosmic microwave background fluctuations, a 

combination fast spin and slow precession motion is used that will cover the entire 

celestial sphere in six months. These rates and the sunline angle are tightly controlled to 

generate the full sky map. Sufficient attitude knowledge is provided to yield instrument 
pointing to a standard deviation (1-sigma) of 1.3 arc-minutes per axis. In addition, the 

spacecraft acquires and holds the sunline at initial acquisition, and in the event of a 

failure. Finally, the spacecraft slews to the proper orbit adjust orientations and to the 

proper off-sunline attitude to start the compound spin. The design and flight performance 

of the reaction wheel based control modes and attitude determination system will be 

discussed. Flight results will be shown for comparison to system requirements. 

“Recent Flight Results of the Trmm Kalman Filter”, Andrews (GSFC), Bilanow (SAIC) 

ABSTRACT: The Tropical Rainfall Measuring Mission (TRMM) spacecraft is a nadir 

pointing spacecraft that nominally controlled the roll and pitch attitude based on the Earth 

Sensor Assembly (ESA) output. TRMM's nominal orbit altitude was 350 km, until raised 

to 402 km to prolong mission life. During the boost, the ESA experienced a decreasing 

signal to noise ratio, until sun interference at 383 km altitude made the ESA data 

unreliable for attitude determination. At that point, the backup attitude determination 

algorithm, an extended Kalman filter, was enabled. After the boost finished, TRMM 

reacquired its nadir-pointing attitude, and continued its mission. This paper will briefly 

discuss the boost and the decision to turn on the backup attitude determination algorithm. 

A description of the extended Kalman filter algorithm will be given. In addition, flight 

results from analyzing attitude data and the results of software changes made onboard 

TRMM will be discussed. Some lessons learned are presented. 

“Partially Decentralized Control Architectures For Satellite Formations”, Carpenter 

(GSFC) 

ABSTRACT: In a partially decentralized control architecture, more than one but less than 

all nodes have supervisory capability. This paper describes an approach to choosing the 

number of supervisors in such an architecture, based on a reliability vs. cost trade. It also 

considers the implications of these results for the design of navigation systems for 

satellite formations that could be controlled with a partially decentralized architecture. 

Using an assumed cost model, analytic and simulation-based results indicate that it may 

be cheaper to achieve a given overall system reliability with a partially decentralized 

architecture containing only a few supervisors, than with either fully decentralized or 

purely centralized architectures. Nominally, the subset of supervisors may act as 

centralized estimation and control nodes for corresponding subsets of the remaining 

subordinate nodes, processing all the measurement data from all their subordinates, or 

fusing the subordinates' local estimates with a scheme that compensates for correlations 

among the local states estimates. The supervisors may then act as decentralized 

estimation and control peers with respect to each other. However, if the state estimates of 

each spacecraft are uncorrelated, the supervisors may command globally optimal 

maneuvers derived from simple differencing of the subordinates' local estimates. Since 

the absolute positions and velocities of each spacecraft are unique, correlations may only 

occur through common biases, common process noise, or indirectly through common 

model errors. A simple example illustrates the discussion on the role of correlation. 
“A Convex Approach to Fault Tolerant Control”, Maghami (GSFC), Cox (NASA LaRC) 

ABSTRACT: The design of control laws for dynamic systems with the potential for 

actuator failures is considered in this work. The use of Linear Matrix Inequalities allows 

more freedom in controller design criteria than typically available with robust control. 

This work proposes an extension of fault-scheduled control design techniques that can 

find a fixed controller with provable performance over a set of plants. Through 

convexity of the objective function, performance bounds on this set of plants implies 

performance bounds on a range of systems defined by a convex hull. This is used to 

incorporate performance bounds for a variety of soft and hard failures into the control 

design problem. 

“GPS-Based Navigation and Orbit Determination for the AMSAT AO-40 Satellite”, G. 

Davis (Emergent Space Technologies), Moreau, Carpenter, Bauer (GSFC) 

ABSTRACT: The AMSAT OSCAR-40 (AO-40) spacecraft occupies a highly elliptical 

orbit (HEO) to support amateur radio experiments. An interesting aspect of the mission 

is the attempted use of GPS for navigation and attitude determination in HEO. Previous 

experiences with GPS tracking in such orbits have demonstrated the ability to acquire 

GPS signals, but very little data were produced for navigation and orbit determination 

studies. The AO-40 spacecraft, flying two Trimble Advanced Navigation Sensor 

(TANS) Vector GPS receivers for signal reception at apogee and at perigee, is the first to 

demonstrate autonomous tracking of GPS signals from within a HEO with no interaction 

from ground controllers. Moreover, over 11 weeks of total operations as of June 2002, 

the receiver has returned a continuous stream of code phase, Doppler, and carrier phase 

measurements useful for studying GPS signal characteristics and performing postprocessed orbit determination studies in HEO. This paper presents the initial efforts to 

generate AO-40 navigation solutions from pseudorange data reconstructed from the 

TANS Vector code phase, as well as to generate a precise orbit solution for the AO 40 

spacecraft using a batch filter. 

“Restoring Redundancy to the MAP Propulsion System”, O'Donnell, Davis, Ward 

(GSFC) 

ABSTRACT: The Microwave Anisotropy Probe is a follow-on to the Differential 

Microwave Radiometer instrument on the Cosmic Background Explorer. Sixteen months 

before launch, it was discovered that from the time of the critical design review, 

configuration changes had resulted in a significant migration of the spacecraft’s center of 

mass. As a result, the spacecraft no longer had a viable backup control mode in the event 

of a failure of the negative pitch axis thruster. Potential solutions to this problem were 

identified, such as adding thruster plume shields to redirect thruster torque, adding mass 

to, or removing it from, the spacecraft, adding an additional thruster, moving thrusters, 

bending thrusters (either nozzles or propellant tubing), or accepting the loss of 

redundancy for the thruster. The impacts of each solution—including effects on the mass, 

cost, and fuel budgets, as well as schedule—were considered, and it was decided to bend 
the thruster propellant tubing of the two roll control thrusters, allowing that pair to be 

used for backup control in the negative pitch axis. This paper discusses the problem and 

the potential solutions, and documents the hardware and software changes that needed to 

be made to implement the chosen solution. Flight data is presented to show the 

propulsion system on-orbit performance. 

“The Microwave Anisotropy Probe (MAP) Mission”, Markley, Andrews, O’Donnell, 

Ward (GSFC) 

ABSTRACT: The Microwave Anisotropy Probe mission is designed to produce a map of 

the cosmic microwave background radiation over the entire celestial sphere by executing 

a fast spin and a slow precession of its spin axis about the Sun line to obtain a highly 

interconnected set of measurements. The spacecraft attitude is sensed and controlled 

using an inertial reference unit, two star trackers, a digital sun sensor, twelve coarse sun 

sensors, three reaction wheel assemblies, and a propulsion system. This paper presents an 

overview of the design of the attitude control system to carry out this mission and 

presents some early flight experience. 

“The Microwave Anisotropy Probe Guidance, Navigation, and Control Hardware Suite”, 

Ward, Davis, O’Donnell (GSFC) 

ABSTRACT: In order to meet the complex attitude determination and control 

requirements of the Microwave Anisotropy Probe (MAP) mission, a diverse set of 

components was used. The set included two Lockheed Martin AST-201 star trackers, 

two Kearfott Two-Axis Rate Assemblies mounted to provide X, Y and redundant Z-axis 

rates, two Adcole Digital Sun Sensor heads sharing one set of electronics, twelve Adcole 

Coarse Sun Sensor eyes, three Ithaco E-sized Reaction Wheel Assemblies, a Propulsion 

Subsystem that employed eight Primex Rocket Engine Modules, and a pair of GSFCdesigned Attitude Control Electronics to connect all of the components to the spacecraft 

processor. The on-orbit success of the MAP Guidance, Navigation, and Control System 

can partially be attributed to the performance of this hardware suite, in addition to the 

successful algorithm and software design work of the MAP project team. The 

performance of this hardware is documented, as are some of the spacecraft 

accommodations and lessons learned that came from working with this particular set of 

hardware. 

“Development of a Two-Wheel Contingency Mode for the MAP Spacecraft”, Starin, 

O’Donnell (GSFC) 

ABSTRACT: Microwave Anisotropy Probe (MAP) is a follow-on mission to the Cosmic 

Background Explorer (COBE), and is currently collecting data from its orbit near the 

second Sun-Earth libration point. Due to limited mass, power, and financial resources, a 

traditional reliability concept including fully redundant components was not feasible for 

MAP. Instead, the MAP design employs selective hardware redundancy in tandem with 

contingency software modes and algorithms to improve the odds of mission success. One 

direction for such improvement has been the development of a two-wheel backup control 

strategy. This strategy would allow MAP to position itself for maneuvers and collect 
science data should one of its three reaction wheels fail. Along with operational 

considerations, the strategy includes three new control algorithms. These algorithms 

would use the remaining attitude control actuators—thrusters and two reaction wheels—

in ways that would achieve control goals while minimizing adverse impacts on the 

functionality of other subsystems and software. 

“An Anomalous Force on the MAP Spacecraft”, Starin, O’Donnell, Ward, Wollack 

(GSFC), Bay (Jackson & Tull), Fink (CSC) 

ABSTRACT: The Microwave Anisotropy Probe (MAP) orbits the second Earth-Sun 

libration point (L2)—about 1.5 million kilometers outside Earth’s orbit—mapping cosmic 

microwave background radiation. To achieve orbit near L2 on a small fuel budget, the 

MAP spacecraft needed to swing past the Moon for a gravity assist. Timing the lunar 

swing-by required MAP to travel in three high-eccentricity phasing loops with critical 

maneuvers at a minimum of two, but nominally all three, of the perigee passes. On the 

approach to the first perigee maneuver, MAP telemetry showed a considerable change in 

system angular momentum that threatened to cause on-board Failure Detection and 

Correction (FDC) to abort the critical maneuver. Fortunately, the system momentum did 

not reach the FDC limit; however, the MAP team did develop a contingency strategy 

should a stronger anomaly occur before or during subsequent perigee maneuvers. 

Simultaneously, members of the MAP team developed and tested various hypotheses for 

the cause of the anomalous force. The final hypothesis was that water was outgassing 

from the thermal blanketing and freezing to the cold side of the solar shield. As radiation 

from Earth warmed the cold side of the spacecraft, the uneven sublimation of frozen 

water created a torque on the spacecraft. 

“State-Dependent Pseudo-Linear Filters for Spacecraft Attitude and Rate Estimation”, 

Bar-Itzhack (Technion), Harman (GSFC), Choukroun (Technion) 

ABSTRACT: This paper presents the development and performance of a special 

algorithm for estimating the attitude and angular rate of a spacecraft. The algorithm is a 

pseudo-linear Kalman filter, which is an ordinary linear Kalman filter that operates on a 

linear model whose matrices are current state estimate dependent. The nonlinear 

rotational dynamics equation of the spacecraft is presented in the state space as a statedependent linear system. Two types of measurements are considered; one type is a 

measurement of the quaternion of rotation, which is obtained from a newly introduced 

star tracker based apparatus. The other type of measurement is that of vectors, which 

permits the use of a variety of vector measuring sensors like sun sensors and 

magnetometers. Several measurement models are presented, a linear model for the case 

where the measured quantity is the quaternion, and two measurement models, one of 

which is pseudo-linear, when the measured quantities are vectors. In the latter situation, 

one model is for the case where the attitude is represented by a quaternion, and the other 

is for the is introduced for the latter case. The state-dependent pseudo-linear filter is 

tested using simulated spacecraft rotations. case where it is represented by a direction 

cosine matrix. A special observability enhancement algorithm is introduced for the latter 
case. The state-dependent pseudo-linear filter is tested using simulated spacecraft 

rotations.

“An Algorithm for Optimal Fusion of Quaternions with Vector-Measurements”, BarItzhack (Technion), Harman (GSFC). 

ABSTRACT: This paper presents an algorithm for optimal blending of two sources of 

attitude information. One source is the attitude quaternion, which fully represents 

attitude, and the other source is a vector measurement, which contains only partial 

information on the attitude. When the two sources, which are of a different nature, are 

mixed properly, the resultant attitude information is superior to that contained in either 

one of the sources. Monte-Carlo simulations are used to demonstrate the algorithm 

presented in the paper.

AIAA/AAS Astrodynamics Specialist Conference, Monterey, CA August 5-8, 2002

“Formation Flying Satellite Control Around The L2 Sun-Earth Libration Point”, Hamilton 

(United States Air Force), Folta, Carpenter (GSFC)

ABSTRACT: This paper discusses the development of a linear control algorithm for 

formations in the vicinity of the L2 sun-Earth libration point. The development of a 

simplified extended Kalman filter is included as well. Simulations are created for the 

analysis of the stationkeeping and various formation maneuvers of the Stellar Imager 

mission. The simulations provide tracking error, estimation error, and control effort 

results. For formation maneuvering, the formation spacecraft track to within 4 meters of 

their desired position and within 1.5 millimeters per second of their desired zero velocity. 

The filter, with few exceptions, keeps the estimation errors within their three-sigma 

values. Without noise, the controller performs extremely well, with the formation 

spacecraft tracking to within several micrometers. Each spacecraft uses around 1 to 2 

grams of propellant per maneuver, depending on the circumstances. 

“Results Of NASA’s First Autonomous Formation Flying Experiment: 

Earth Observing-1 (EO-1)”, Folta (GSFC), Hawkins (ai solutions) 

ABSTRACT: NASA’s first autonomous formation flying mission completed its primary 

goal of demonstrating an advanced technology called enhanced formation flying. To 

enable this technology, the Flight Dynamics Analysis Branch at the Goddard Space 

Flight Center implemented a universal 3-axis formation flying algorithm in an 

autonomous executive flight code onboard the New Millennium Program’s (NMP) Earth 

Observing-1 (EO-1) spacecraft. This paper describes the mathematical background of the 

autonomous formation flying algorithm, the onboard flight design and the validation 

results of this unique system. Results from fully autonomous maneuver control are 

presented as comparisons between the onboard EO-1 operational autonomous control 

system called AutoConTM, its ground-based predecessor used in operations, and the 
original standalone algorithm. Maneuvers discussed encompass reactionary, routine 

formation maintenance, and inclination control. Orbital data is also examined to verify 

that all formation flying requirements were met. 

"An Overview of Trajectory Design Operations for the Microwave Anisotropy Probe 

Mission", Cuevas, Newman, Mesarch, Woodard (GSFC). 

ABSTRACT: The purpose of this paper is to document the results of the pre-launch 

trajectory design and the real-time operations for the Microwave Anisotropy Probe 

(MAP) mission, launched on June 30, 2001. Once MAP was successfully inserted into a 

highly elliptical phasing orbit, three perigee maneuvers and a final perigee correction 

maneuver were performed to tailor a lunar encounter on July 30, 2001. MAP achieved its 

final Lissajous orbit (0.5° by 10.5°) about the Sun-Earth/Moon L2 libration point via this 

lunar encounter. This paper will show the maneuvers that were designed to arrive at the 

mission orbit. A further discussion of how the MAP trajectory analysts altered the prelaunch phasing loop maneuvers as well as the lunar encounter to meet all mission 

constraints, including the constraint of zero lunar shadows is also included. 

"The Maneuver Planning Process For The Microwave Anisotropy Probe (Map) Mission", 

Mesarch, Andrews (GSFC). 

ABSTRACT: The Microwave Anisotropy Probe (MAP) mission utilized a strategy 

combining highly eccentric phasing loops with a lunar gravity assist to provide a zerocost insertion into a Lissajous orbit about the Sun-Earth/Moon L2 point. Maneuvers were 

executed at the phasing loop perigees to correct for launch vehicle errors and to target the 

lunar gravity assist so that a suitable orbit at L2 was achieved. This paper will discuss the 

maneuver planning process for designing, verifying, and executing MAP's maneuvers. 

This paper will also describe how commercial off-the-shelf (COTS) tools were used to 

execute these tasks and produce a command sequence ready for upload to the spacecraft. 

These COTS tools included Satellite Tool Kit, MATLAB, and Matrix-X. 

"Contingency Planning for the Microwave Anisotropy Probe Mission", Mesarch (GSFC), 

Rohrbaugh, Schiff (ai solutions) 

ABSTRACT: The Microwave Anisotropy Probe (MAP) utilized a phasing loop/lunar 

encounter strategy to achieve a small amplitude Lissajous orbit about the SunEarth/Moon L2 libration point. The use of phasing loops was key in minimizing MAP’s 

overall ∆V needs while also providing ample opportunities for contingency resolution. 

This paper will discuss the different contingencies and responses studied for MAP. These 

contingencies included accommodating excessive launch vehicle errors (beyond 3σ), 

splitting perigee maneuvers to achieve ground station coverage through the Deep Space 

Network (DSN), delaying the start of a perigee maneuver, aborting a perigee maneuver in 

the middle of execution, missing a perigee maneuver altogether, and missing the lunar 

encounter (crucial to achieving the final Lissajous orbit). It is determined that using a 

phasing loop approach permits many opportunities to correct for a majority of these 

contingencies. 
“Preliminary Optimal Orbit Design For The LaserInterferometer Space Antenna (LISA)” 

Hughes (GSFC) 

ABSTRACT: In this paper we present a preliminary optimal orbit analysis for the Laser 

Interferometer Space Antenna (LISA). LISA is a NASA/ESA mission to study 

gravitational waves and test predictions of general relativity. The nominal formation 

consists of three spacecraft in heliocentric orbits at 1 AU and trailing the Earth 

by twenty degrees. This configuration was chosen as a trade off to reduce the noise 

sources that will affect the instrument and to reduce the fuel to achieve the final orbit. 

We present equations for the nominal orbit design (Folkner 97) and discuss several 

different measures of performance for the LISA formation. All of the measures directly 

relate the formation dynamics to science performance. Also, constraints on the formation 

dynamics due to spacecraft and instrument limitations are discussed. Using the nominal 

solution as an initial guess, the formation is optimized using Sequential Quadratic 

Programming to maximize the performance while satisfying a set of nonlinear 

constraints. Results are presented for each of the performance measures. 

“On the Singularity in the Estimation of the Quaternion-of-Rotation”, Thienel (GSFC), 

Bar-Itzhack (Technion) 

ABSTRACT: It has been claimed in the archival literature that the covariance matrix of a 

Kalman filter, which is designed to estimate the quaternion-of-rotation, is necessarily 

rank deficient because the normality constraint of the quaternion produces dependence 

between the quaternion elements. In reality, though, this phenomenon does not occur. 

The covariance matrix is not singular, and the filter is well behaved. Several simple 

examples are presented that demonstrate the regularity of the covariance matrix. First, a 

Kalman filter is designed to estimate variables subject to a functional relationship. Then 

the particular problem of quaternion estimation is analyzed. It is shown that the 

discrepancy stems from the fact that the functional relationship exists between the 

elements of the quaternion but not between its estimated elements. 

ION GPS 2002, September 24-27, 2002, Portland, OR

“Results from the GPS Flight Experiment on the High Earth Orbit AMSAT OSCAR-40 

Spacecraft”, Moreau, E. Davis, Carpenter (GSFC), Kelbel (Computer Sciences 

Corporation), G. Davis (Emergent Space Technologies), Axelrad (University of 

Colorado) 

ABSTRACT: A GPS receiver flying on the High Earth Orbit (HEO) AMSAT-OSCAR 

40 (AO 40) spacecraft has been returning GPS observations from high above the altitude 

of the GPS constellation. AO 40, an amateur radio satellite launched November 16, 

2000, is currently in a low inclination, 1000 by 59000 km altitude orbit. This low-cost 
experiment utilizes a mid 1990’s era, 6-channel, C/A code receiver configured with high 

gain receiving antennas for tracking above the GPS constellation. The receiver has 

performed well, despite operating significantly outside of its original design environment. 

It has regularly returned GPS observations from points all around the orbit, with over ten 

weeks of GPS tracking data collected to date. Signal to noise levels as high as 48 dB Hz 

have been recorded near apogee, when the spacecraft was at an altitude of close to 60000 

km. GPS side lobe signals have been tracked on several occasions, primarily from Block 

IIR GPS satellites. Although the receiver has not computed a solution in real-time, point 

solutions have been computed on the ground using simultaneous measurements from four 

satellites. This experiment has provided important experience dealing with the many 

challenges inherent to GPS tracking at high altitudes, and the measurements returned are 

providing valuable information about the characteristics of GPS signals available for 

future HEO users. 

CONFERENCE PRESENTATIONS 

NTA National Conference, Las Vegas, NV, 

“Guidance, Navigation & Control Innovations at the NASA Goddard Space Flight 

Center”, Dr. A. Ericsson 
Appendix D- Acronyms and Abbreviations 

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

AAS American Astronautical Society 

ACE Attitude Control Electronics 

ACS Attitude Control System 

AETD Applied Engineering and Technology Directorate 

AGI Analytical Graphics, Inc. 

AIAA American Institute of Aeronautics and Astronautics 

AKM Apogee Kick Motor 

APL Applied Physics Laboratory 

ATS Absolute Time Sequence 

AU Astronomical Unit 

AUTOFDS Autonomous Flight Dynamics System 

BRDF Bi-directional Reflectance Distribution Function 

BSS Boeing Satellite Systems 

CCSDS Consultative Committee for Space Data Systems 

COTS Commercial Off-the-Shelf 

DFC Drag Free Control 

DoD Department of Defense 

DOF Degree of Freedom 

DRS Disturbance Reduction System 

DSC Deep Space Calibration 

DSN Deep Space Network 

DSS Distributed Space System 

DST Dynamical Systems Theory 

EFF Enhanced Formation Flying 

EKF Extended Kalman Filter 

EM Expectation Maximization 

EO Earth Observing 

EOS Earth Observing System 

EPGM EOS Polar Ground Network 

EPM Earth Point Mode 

ERBS Earth Radiation Budget Satellite 

ESA European Space Agency 

ESE Earth Science Enterprise 

ESTO Earth Science Technology Office 

FD Flight Dynamics 

FDAB Flight Dynamics Analysis Branch 

FDC Fault Detection Correction 

FDF Flight Dynamics Facility 

FDT Flight Dynamics Team 

FFTB Formation Flying Test Bed 

FPM Fine Point Mode 

FOT Flight Operations Team 
FSW Flight Software 

FUSE Far Ultraviolet Spectroscopic Explorer 

FY Fiscal Year 

GA Genetic Algorithm 

GALEX Galaxy Evolution Explorer 

GEC Geospace Electrodynamic Connections 

GEO Geosynchronous Earth Orbit 

GEODE GPS Enhanced Orbit Determination Experiment 

GEONS GPS-Enhanced Orbit Navigation System 

GMSEC Goddard Mission Services Evolution Center 

GN&C Guidance, Navigation and Control 

GNC Guidance, Navigation and Control 

GNCD Guidance, Navigation, and Control Division 

GOES Geostationary Operational Environmental Satellite 

GPM Global Precipitation Mission 

GPS Global Positioning Satellite 

GRS Gravity Reference Sensor 

GSFC Goddard Space Flight Center 

GSRP Graduate Student Research Program 

GTDS Goddard Trajectory Determination System 

GTO Geostationary Transfer Orbit 

GUI Graphics User Interface 

GUS Gyroscopic Upper Stage 

HEO High Earth Orbit/ Highly Elliptical Orbit 

HGA High Gain Antenna 

HRG Hemispherical Resonator Gyro 

HTML HyperText Markup Language 

I&T Integration and Test 

ICD Interface Control Document 

IMDC Integrated Mission Design Center 

IMU Inertial Measurement Unit 

INCA Interactive Controls Analysis 

ITAR International Traffic In Arms Regulation 

JPL Jet Propulsion Laboratory 

L&EO Launch and Early Orbit 

LAQ Liquid Apogee Motor 

LEO Low Earth Orbit 

LISA Laser Interferometric Space Antenna 

LPT Low Power Transceiver 

LQG Linear Quadratic Gaussian 

LSE Lunar Science Explorer 

MAGNAV Magnetometer Navigation 

MAP Microwave Anisotropy Probe 

MAXIM Micro-Arcsecond X-ray Imaging Mission 

MC Master Catalog 

MCC Mid Course Correction 
Mi Instrumental Magnitude 

MLT Mean Local Time 

MMS Magnetic Multi-scale Mission 

MMSE Minimum Mean Squared Estimate 

MMSCAT Multi-Mission Star Catalog 

MOC Mission Operations Center 

Mv Visual Magnitude 

NASA National Aeronautical and Space Administration 

NGST Next Generation Space Telescope 

NOAA National Oceanic and Atmospherics Administration 

OSC Orbital Sciences Corporation 

OD Orbit Determination 

PC Personal Computer 

PD Proportional Derivative 

PI Principal Investigator 

PM Proof Mass 

POR Power On Reset 

PR Precipitation Radar 

PREST Program of Research and Education in Space Technology 

R&D Research and Development 

RF Radio Frequency 

RFI Radio Frequency Interference 

RMS Root-Mean-Square 

RSDO Rapid Spacecraft Development Office 

RTADS Real Time Attitude Determination System 

RXTE Rossi X-Ray Timing Explorer 

SADA Solar Array Drive Assembly 

SAMPEX Solar Anomalous and Magnetospheric Particle Explorer 

SAM Scan Angle Monitor 

SDO Solar Dynamics Observatory 

SET Single Event Transient 

SIGI Space Integrated Global Position System Inertial Navigation System 

SK Stationkeeping 

SMEX Small Explorer 

SOHO Solar and Heliospheric Observatory 

SOMO Space Operations Management Office 

SPM Sun Point Mode 

SPS Standard Positioning Service 

SSE Space Science Enterprise 

ST Space Technology 

STP Solar Terrestrial Probe 

TAM Three Axis Magnetometer 

TDRSS Tracking Data Relay Satellite System 

TIMED Thermosphere, Ionosphere, Mesosphere Energetics and Dynamics 

TRL Technology Readiness Level 

TRMM Tropical Rainfall Measuring Mission 
URL Uniform Resource Locator 

USGS United States Geological Survey 

USN Universal Space Network 

USNO United States Naval Observatory 

VESPER Venus Sounder for Planetary Exploration 

WIRE Wide-Field Infrared Explorer 

WRS World Reference System 

WSB Weak Stability Boundary 

WWW World Wide Web