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January 2000

Flight Dynamics Analysis Branch

End of Fiscal Year 1999 Report

T. Stengle, Goddard Space Flight Center, Greenbelt, MD

F. Flores-Amaya, Goddard Space Flight Center, Greenbelt, MD

National Aeronautics and

Space Administration

Goddard Space Flight Center

Greenbelt, Maryland 20771

NASA/TM—2000–209485
Acknowledgments

Available from:

NASA Center for AeroSpace Information National Technical Information Service

7121 Standard Drive 5285 Port Royal Road

Hanover, MD 21076-1320 Springfield, VA 22161

Price Code: A17 Price Code: A10

The contents of this report are based on inputs generously supplied by members of the Flight Dynamics

Analysis Branch (FDAB), a subdivision of the Guidance, Navigation and Control Center (GNCC), at

NASA/Goddard Space Flight Center (GSFC).

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

management:

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

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

Published copies of this report will be available from Catherine A. Waltersdorff, Flight Dynamics Analysis Branch, Code 572, NASA/Goddard Space Flight Center, Greenbelt, MD 20771, as well as the NASA

Center for AeroSpace Information and National Technical Information Service, both listed below.

A Hypertext Markup Language (HTML) version of this document will be available on the Internet World

Wide Web. Contact Felipe Flores-Amaya, Code 572, NASA/Goddard Space Flight Center, Greenbelt,

MD 20771, for the Uniform Resource Locator (URL).
iii

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) 1999. 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 analysis results and mission experience derived from the

various mission support roles. The primary focus of the FDAB is to provide expertise in the discipline of

flight dynamics, which involves spacecraft trajectory (orbit) and attitude analysis, as well as orbit and

attitude determination and control. The FDAB currently provides support for missions involving NASA,

government, university, and commercial space missions, at various stages in the mission life cycle.
v

TABLE OF CONTENTS

1.0 INTRODUCTION ........................................................................................................................1

2.0 FLIGHT PROJECT SUPPORT .....................................................................................................1

2.1 DEVELOPMENT MISSIONS .........................................................................................1

2.1.1 EO-1 ......................................................................................................................1

2.1.2 EOS AM-1 .............................................................................................................2

2.1.3 EOS PM-1 ..............................................................................................................2

2.1.4 GOES-L..................................................................................................................4

2.1.5 InFocus (ULDB) ....................................................................................................4

2.1.6 Landsat-7 ...............................................................................................................5

2.1.7 MAP ......................................................................................................................5

2.1.8 Spartan ...................................................................................................................6

2.1.9 Triana .....................................................................................................................7

2.1.10 WIRE ...................................................................................................................9

2.2 OPERATIONAL MISSIONS .........................................................................................11

2.2.1 ERBS ...................................................................................................................11

2.2.2 Lunar Prospector ..................................................................................................11

2.2.3 TRMM .................................................................................................................13

3.0 STUDY MISSIONS SUPPORT ..................................................................................................13

3.1 IMDC ............................................................................................................................13

3.2 Mission Concept Support ...............................................................................................15

4.0 TECHNOLOGY DEVELOPMENTS .........................................................................................16

4.1 Advanced Mission Design ..............................................................................................16

4.2 Autonomous Relative Navigation & Formation Flying ....................................................16

4.3 Advances in Navigation Technology ..............................................................................18

4.3.1 Onboard Navigation Systems.................................................................................18

4.3.2 Global Positioning System (GPS) Orbit Determination .........................................19

4.3.3 “Compound Eye” GPS Attitude and Navigation Sensor .........................................20

4.3.4 Autonomous Navigation .......................................................................................21

4.3.5 Navigation Studies................................................................................................21

4.4 Attitude Determination ...................................................................................................23

4.4.1 Advanced Attitude Analysis ..................................................................................23

4.4.2 Attitude Models Task ...........................................................................................24

4.4.3 Magnetometer Navigation ....................................................................................25

5.0 GNCC FLIGHT DYNAMICS LABORATORY ..........................................................................26

6.0 INTERAGENCY ACTIVITIES ..................................................................................................27

6.1 CCSDS ..........................................................................................................................27

6.2 Flight Mechanics Symposium ........................................................................................27

7.0 ISO9000 .....................................................................................................................................28
vi

8.0 OUTREACH ACTIVITIES ........................................................................................................28

8.1 University Satellite Operations— SAMPEX.....................................................................28

8.2 Graduate Student Researchers Program— GSRP..............................................................28

8.3 Visiting Student Enrichment Program— VSEP ................................................................28

8.4 Public Education/Community Outreach...........................................................................29

APPENDIX A ...................................................................................................................................A1

APPENDIX B ...................................................................................................................................B1
1

1.0 Introduction

This document summarizes the major activities and accomplishments carried out by the Goddard Space

Flight Center (GSFC)’s Flight Dynamics Analysis Branch (FDAB), Code 572, in support of flight projects

and technology development initiatives in Fiscal Year (FY) 1999. 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 major accomplishments in the FDAB in FY99 were:

• Provided flight dynamics support to the Lunar Prospector, MAP, and Triana missions, among a variety

of spacecraft missions

• Sponsored the Flight Mechanics Symposium

• Supported the Consultative Committee for Space Data Systems (CCSDS) workshops

• Performed numerous analyses and studies for future missions

• Started the Flight Dynamics Laboratory for in-house mission analysis and support

• Complied with all requirements in support of GSFC ISO9000 certification

• Conducted acceptance testing of MAP ACS and C&DH flight software as part of the flight software

testing team

• Provided launch support and anomaly resolution support for Landsat-7, TRMM, and ERBS.

2.0 Flight Project Support

2.1 Development Missions

2.1.1 EO-1 Earth Observing-1

Earth Observing–1, the first spacecraft in the New Millennium Program, is scheduled for launch in midApril, 2000. The Guidance, Navigation and Control Center (GNCC) has multiple interests in the successful

launch and operations of EO-1. Initially, EO-1 was solely a satellite to showcase more than 10 new and

largely separate technologies. The foremost new technology was the Advanced Land Imager (ALI), which

was intended as the next generation Landsat-type imager. The primary mission was designed to fly EO-1

one minute along track behind Landsat-7 and take the same image that Landsat-7 had just taken for

comparison. The ALI detectors are located such that they cover the easternmost 1/5 of the Landsat-7 swath.

There is also a cross track requirement of +/- 3 km which keeps the ALI detectors in the desired portion of

the Landsat-7 swath. A second instrument, the GSFC Atmospheric Corrector (AC), was designed to operate

with the ALI. About 1 year before launch, a companion instrument to the ALI and the AC, TRW’s

Hyperion instrument, was added to the spacecraft. This additional imager was to have flown on the failed

Lewis spacecraft and will image a narrow swath that includes the westernmost 1/5 of the ALI swath.

Another GNCC-based technology is the Pulsed Plasma Thruster designed for attitude pitch control.

Currently, the validation of this technology will occur after the prime imaging mission is completed.

The GNCC has played a major role in designing and implementing the EO-1 Flight Dynamics Support

System (FDSS) which will provide flight dynamics support entirely within the EO-1 Mission Operations

Center (MOC). After the first 60 days, Flight Operations Team (FOT) personnel will perform the entire

flight dynamics function within the MOC. During the first 60 days, GNCC personnel will support the

ascent and checkout phases of the mission. GNCC personnel must design the FDSS operational procedures

and train the FOT to achieve operational expertise with the FDSS. There are four major areas of flight

dynamics support for EO-1. They are: attitude validation and calibration, orbit determination, orbit

maneuver control and orbit and attitude products generation. Most of these functions are supported with

COTS-based software that has been adapted to meet EO-1 FDSS requirements. Approximately 60 percent

of the FDSS EO-1 procedures will be executed via automation. This will reduce the execution time of

certain FDSS tasks by a factor of 10.
2

The orbit maneuver control function showcases a major GNCC accomplishment. One of the EO-1

technologies is Enhanced Formation Flying (EFF), which has the goal of autonomous onboard orbit

maneuver control. The FDSS will provide a ground-based EFF tool, AUTOCON-G, to validate the onboard

EFF algorithms. After the first few months of ground-based EFF operations, designed to keep EO-1 in the

strict formation with Landsat-7, the onboard EFF software, AUTOCON-F, will be designated as prime with

the ground as a backup. The orbit and attitude product generation capabilities in the FDSS provide about

three dozen products that aid in spacecraft health and safety maintenance, command generation in the MOC

and image planning for the groups that built the ALI, AC and Hyperion instruments.

The FDSS has had all planned software releases delivered as of 7/31/99 and FDSS mission procedures are

being written with 10/1/99 targeted as the completion date. The last 4 months before launch will include

extensive simulations of the entire EO-1 ground system including the FDSS. To date, spacecraft integration

and testing problems have delayed most ground system proficiency testing and simulations. The last

quarter before launch will be one of the busiest in recent memory.

Finally, the GNCC has provided extensive mission analysis support to the EO-1 Project in the areas of

trajectory and launch window computations for launch vehicle support, ascent maneuver planning to aid in

Launch & Early Orbit timeline planning and product generation to support spacecraft and ground system

Integration & Test activities.

[Technical contacts: Robert L. DeFazio/572; Richard J. Luquette/572]

2.1.2 EOS AM-1 Earth Observing System Terra

The EOS Terra spacecraft, formerly known as EOS AM-1, is currently in the final stages of preparation for

launch from Vandenburg Air Force Base. Launch delays due to spacecraft hardware and mission operation

center software modifications, as well as a recent investigation of the Atlas IIAS upper-stage rocket, has

postponed the scheduled launch to the first quarter of FY 2000. Guidance, Navigation and Control (GN&C)

activities during the prelaunch phase included a successful completion of the comprehensive performance

test (CPT), a delivery of injection targets to the Lockheed Martin/Atlas team, as well as participation in

various launch and early orbit and routine operations simulations.

GN&C testing in August 1998 resulted in the discovery of a voltage leakage condition in the Attitude

Control Electronics (ACE) hardware. Specifically, it was found that an input signal that was outside the

voltage range of the ACE Safe Hold Digital Processor analog MUX resulted in leakage into the other MUX

channels. Input to the MUX consists of position, rate and magnetometer signals. The redundant ACE

hardware was removed from the spacecraft to install zener diodes in order to limit the MUX input signals.

The ACE was successfully reintegrated onto the spacecraft in January 1999.

In addition, the Flight Dynamics System was tested and enhanced throughout the year to meet changing

mission requirements. An automation system, autoProducts, was implemented as a front end to the FDS.

autoProducts allows a nonexpert user to access all of the Commercial-Off-the-Shelf (COTS) and

Government-off-the-Shelf (GOTS) software that comprises the FDS through a common user interface,

changing only those parameters necessary to generate planning products. The system significantly

simplifies the process of generating the 85 required products from several different pieces of software, as

well as minimizing chances for user errors.

[Technical contact: Lauri Newman/572]

2.1.3 EOS PM-1 Earth Observing System PM-1

EOS PM-1 Mission— Completed Flight Dynamics System (FDS) support requirements definition through

meetings with the Project representatives, FOT members, and the Principal Investigators (PI’s) for each

instrument. Provided FDS inputs to the Mission Support Requirements Document (MSRD) based on the

meetings and drafted Interface Control Documents (ICD’s) for each entity. Performed orbit determination

error analysis, contact requirements analysis, launch window/maneuver analysis, and provided preliminary

inputs for attitude sensor calibration plan. Negotiated attitude sensor telemetry requirements with the
3

spacecraft vendor. Supported project-level meetings with the spacecraft vendor, PI’s, and other GSFC

mission interfaces.

Several of the FDS Support Products for EOS PM had accuracy requirements of 1 second after 7 days. This

requirement could not be met to 3-sigma certainty until approximately 1/2002 (using Schatten +2 sigma

predicted solar flux values).

The accuracy of the FDS support products is limited by the accuracy of the predicted ephemeris used in the

product generation. Knowledge of atmospheric density is necessary for the production of the predicted

ephemeris, but the uncertainty in the atmospheric density model is proportional to the uncertainty in the

solar flux. Daily solar flux variations are larger during high flux periods than when average solar flux

values are low. Periods of high average solar flux (which peak near the planned PM launch date) result in

greater uncertainty in the atmospheric model and consequently reduce the potential accuracy of the

predicted ephemeris.

Flight dynamics (FD) orbit error analysis has shown that the 3-sigma, along-track uncertainty near the

planned PM launch date is approximately 31 km at 7 days @ flux = 200 solar flux units (sfu), the +2 sigma

flux prediction for this date. An along-track uncertainty of 31 km corresponds to ~4 seconds of uncertainty

in the FD products. The FD products uncertainty is ~1 second @ 4 days when sfu = 200. The along-track

and timing uncertainties at two other +2 sigma flux values (150 and 100) were also analyzed (see table 2.1

below) and the corresponding mean flux values at these times are included.

Table 2.1

Date Solar Flux

 Mean Value

(sfu)

Solar Flux

 +2 sigma

(sfu)

Along-Track Error

7-day, +2 sigma

 Km (sec)

Along-Track Error

4-day, +2 sigma

Km (sec)

12/2000 177 201 30.4 (4) 7.5 (1)

11/2001 139 150 9.3 (1.3)

 5/2003 97 100 1.4 (0.2)

 

FDS had also received a requirement to provide a definitive ephemeris accurate to 15 meters root-sum-square

(rss) for science data processing. FD personnel used the Orbit Determination Error Analysis System

(ODEAS) to analyze this requirement using a variety of contact scenarios specified by the Project. Although

the requirement theoretically was met according to ODEAS, the result (~14.7 meters rss) was too close to the

requirement to guarantee 3 sigma accuracy since all sources of error are not modeled in ODEAS. The FD

personnel negotiated with the project to relax this requirement to 20 meters rss based on these results.

Maximum 3-sigma Total Definitive Position and Velocity Errors Using Tracking Data From Wallops and

Svalbard With Varying Definitive Lengths and Minimum Ground Elevation Angle Cutoffs. (See table 2.2 below.)

Table 2.2

Total Definitive Position and Velocity Orbit Errors (3-sigma)

5-Degree Minimum

Elevation Angle Cutoff

7-Degree Minimum

Elevation Angle Cutoff

15-Degree Minimum

Elevation Angle Cutoff

Definitive

Span (days)

Position

(m)

Velocity

(m/s)

Position

(m)

Velocity

(m/s)

Position

(m)

Velocity

(m/s)

1 30.96 0.0275 27.65 0.0257 26.53 0.0242

2 18.62 0.0155 17.00 0.0144 14.65 0.0127

3 21.40 0.0183 19.80 0.0170 17.44 0.0164

[Technical contact: David Tracewell/572]
4

2.1.4 GOES-L Geostationary Operational Environmental Satellite

The GOES-L spacecraft is the fourth of a five spacecraft series built by Space Systems/Loral and placed on

station by a GSFC FOT supported by GNCC and Computer Sciences Corporation Flight Dynamics

personnel. The first three spacecraft in the series were launched on an Atlas I launch vehicle (L/V). After

that L/V series was phased out, the remaining spacecraft were manifested on an Atlas IIA. This later L/V

had significantly more performance than the Atlas I, which allowed a transfer orbit redesign leading to

more mission lifetime. That was the good news. On the reverse side, getting a firm manifest slot on the

Atlas IIA has been a disaster. Our Flight Dynamics team planned no less than 7 mission profiles for GOESL ranging from January 1999 to June 1999. Finally, we had a firm launch slot on May 15, 1999, had our

Flight Dynamics team well trained and got to within 1 month of launch. Then, the Atlas IIA, along with

several other launch vehicles, experienced a very serious RL-10 engine problem on its Centaur stage. The

launch of GOES-L was postponed indefinitely while the FOT and the Flight Dynamics team try to remain

intact and ready to support. The GOES Project attempts to sustain operational readiness by holding

monthly simulations and proficiency testing until a firm launch date is selected. The best estimate for a

launch of GOES-L, as of mid-August 1999, is sometime in the first quarter of calendar year 2000.

[Technical contact: Robert L. DeFazio/572]

2.1.5 InFocus Mission

Ultra Long Duration Balloon (ULDB) program

The InFocus telescope is an 8-meter-long astronomical x-ray telescope being developed at GSFC for flight

on a high-altitude (40 km) stratospheric balloon. Whereas most pointed instruments on balloons use an

azimuth/elevation system, InFocus is aiming to achieve arc second-level pointing performance using a

single cup/ball oil-filled bearing between the pointed and nonpointed sections of the gondola. With this

rotational isolation the pointed section, consisting of the telescope and support subsystems, is controlled

similarly to an orbiting pointed payload, using reaction wheels, magnetic torquers, star trackers and gyros.

Studies are underway to specify critical control elements in the pointing control system via simulation and

analysis. The complicated pendulous dynamics of the load train (connecting the balloon and gondola) due

to aerodynamic loads are being studied to insure that disturbances on the pointed section due to mass

unbalances about the ball center can be adequately controlled.

To this end, a demonstration balloon flight within the next year to measure disturbances is planned. No

telescope will be on this flight. A second flight will carry the InFocus telescope using the standard

azimuth/elevation pointing system. The third and final flight for InFocus will incorporate the ball/cup

design and is anticipated to occur in 2 years.

Additionally, work is now underway to model and recommend changes to the standard 255-foot-long

balloon train (consisting of all elements between the balloon and gondola, such as the parachute and

separation systems) for the purpose of minimizing disturbance oscillations. This activity is intended to

improve the performance for all pointed balloon payloads, particularly those that use azimuth/elevations

designs. Depending on study results there could be a fourth flight within the next 2 years to qualify a newer

train design.

The analysis required for all flights is being provided by the FDAB.

[Technical contact: David Olney/572]
5

2.1.6 Landsat-7

Landsat-7 Mission Support From GNCC

The Landsat-7 spacecraft was launched on a Delta 7920 expendable launch vehicle (ELV) from the

Western Range (WR) at VAFB at 18:32:00.288 GMT on April 15, 1999. Post-deployment tracking,

commanding and telemetry were provided by the Tracking and Data Relay Satellite (TDRS) and the ground

sites located at Svalbard, Norway, Fairbanks, Alaska, Wallops Island, Virginia, and Sioux Falls, South

Dakota.

Several orbit-raising maneuvers were performed during the first 90 days of the mission to take the Landsat7 satellite from the Delta parking orbit to its mission orbit. As part of the orbit-raising sequence, the

Landsat-7 satellite did a cross-calibration with the Landsat-5 satellite. During this cross-calibration period

the Landsat-7 ground track was on the same World Reference System (WRS) as the Landsat-5 groundtrack.

The cross-calibration period lasted approximately 2 days after which the Landsat-7 completed its ascent to

mission orbit. The mission orbit for Landsat-7 is a 705 km, Sun-synchronous, near-polar orbit with a Mean

Solar Local Time (MSLT) at the descending node of 10:00AM +/- 15 minutes. The orbit ground track

repeats every 16 days. Periodic orbit adjust maneuvers are done throughout the life of the mission to keep

Landsat-7 within 5 km of the WRS. Additionally, at least one inclination maneuver will be required during

the life of the mission to keep the MSLT at the descending node within the mission limits.

Landsat-7 flight dynamics support was performed from the Landsat-7 MOC on designated flight dynamics

workstations. The Landsat-7 orbit determination, acquisition data generation, product generation, maneuver

planning and calibration, and attitude support were supported from the MOC. The attitude telemetry was

received directly from MOC system and used for various attitude displays. Tracking data were received

from the Space Network (SN) and several ground stations and used in the orbit determination process.

Attitude telemetry was received and displayed at the flight dynamics workstations and processed for gyro

bias determination. Flight dynamics also processed the Celestial Sensor Assembly (CSA) to assist in the

verification of the onboard attitude determination system. Acquisition data and product generation are

performed routinely throughout the mission.

The software system used for support of Landsat-7 GNC support was based on COTS software.

At launch +90 days, the Landsat-7 flight dynamics support was assumed by the FOT and will continue to

be supported by the FOT throughout the life of the mission. All orbit determination, attitude support,

acquisition data generation, product generation, and maneuver planning and calibration is done by the FOT.

Flight dynamics personnel from the GNCC at GSFC will be available to the FOT for consultation during

the lifetime of the mission.

In addition to flight dynamics support, the FDAB and GNCC provided real-time ACS flight support for

Landsat-7 from launch simulations through launch, orbit maneuvers, and final instrument checkout. The

ACS support provided by civil servant and contractor personnel helped to make the launch and early

operations phase of the Landsat-7 mission go extremely smoothly.

[Technical contact: Sue Hoge/572]

2.1.7 MAP Microwave Anisotropy Probe

FY 1999 was a busy one for the MAP team as they ready for a launch in the second half of the year 2000.

FDAB members of the ACS, flight dynamics, flight software, and flight software test teams began and

completed a number of important milestones during the year.

A few remaining analysis issues were studied for the MAP mission during the year. These included further

study of solar radiation pressure at L2 and its effect on spacecraft system momentum buildup. Additionally,

a decision was made to mount a second Digital Sun Sensor (DSS) head to the spacecraft to provide further
6

spacecraft rate backup during perigee passes, since it is possible that the radiation environment at that orbit

altitude may cause the star trackers to be unreliable. Analysis of the orientation for the second DSS head

was done to make sure that it would provide backup rates over an acceptable range of orientation during the

perigee pass. A final ACS analysis activity that occurred during the fiscal year was support for the

spacecraft Delta critical design review (CDR).

The major activity for FDAB personnel on the MAP team was support for flight software and flight

software testing. Flight software walk-throughs of both hand-coded and “AutoCoded” (automatically

generated flight software from the analysis and simulation tool XMath) software were supported.

Additionally, spacecraft ACS FDC was further developed and began to be implemented and tested within

the flight software. A number of critical flight software testing milestones were passed, as well. ACS flight

software build testing was completed near the beginning of the fiscal year. At the time of this writing,

acceptance testing of ACS and C&DH (Command & Data Handling) flight software is slated to be

completed by the end of September 1999. An important goal of MAP’s acceptance testing effort— the

inclusion of flight dynamics and ground ops personnel early in the process, in order to clarify as many of

the necessary interfaces as possible— was achieved.

The final MAP activity supported by FDAB personnel was spacecraft I&T. The TARA’s, RWA’s, and

DSS’s have been integrated to the spacecraft for initial checkouts. The MAC (main spacecraft electronics

and processor) box has also been integrated. Various functional tests have been conducted on integrated

hardware, including sensor and actuator phasing in all control modes. Additionally, the electronics interface

to the propulsion subsystem has been tested. As this report is being written, I&T activities continue, and the

FDAB members of the MAP team are gearing up to provide support for spacecraft CPT’s.

The public MAP home page is at http://map.gsfc.nasa.gov. You can see real-time pictures of the spacecraft,

the instrument, or the software test area at http://mapweb.gsfc.nasa.gov/wbforms/webpics/.

[Technical contacts: Steve Andrews/572; Karen Richon/572]

2.1.8 SPARTAN

Spartan 201-5

The Spartan 201-5 solar physics mission was launched on STS-95 October 29, 1998. It was deployed

October 31 and successfully completed its 9-day mission for the Solar Astronomical Observatory (SAO)

before being retrieved and returned to Earth. The FDAB supported the mission by determining the desired

deploy orientation and computing the remote adjust values to be entered by the shuttle crew immediately

before deploy.

Spartan 250

The Spartan 250 series of spacecraft is planned to be an upgrade of the existing Spartan 200 series. It is

designed for better performance and increased mission lifetime. It will have a Power PC CPU-based digital

attitude determination and control system and an RF communication capability.

The Spartan 251 missions will carry an experimental microsatellite mission for the Air Force. Preparations

for the Spartan 251/XSS-10 mission were stopped in May when it was determined that the mission would

not be manifested within a timeframe acceptable to the Air Force. The XSS-10 mission will now be flown

on an ELV, without the Spartan carrier. Before work was halted the digital ACS, onboard environmental

models, and dynamic simulator algorithms had been fully developed and tested, and were being transferred

to flight software.

We are now working on the Spartan 251/XSS-11 mission that is scheduled for launch in the 2002–2003

timeframe. Only slight modifications to the existing Spartan 251/XSS-10 ACS algorithms described above

will be required to support this mission. The design lifetime for this mission will be greater than that of the

XSS-10 mission. This will result in a need for reaction wheel actuators and further modifications to the

ACS algorithms.
7

Spartan Lite

The Spartan Project has discontinued proposal work on the Spartan Lite spacecraft. It was determined that

the need for small spacecraft buses of this sort is being adequately met by industry so the focus is now on

Spartan 250 and Spartan 400 spacecraft.

Spartan 400

Spartan 400 is Spartan project’s newest proposal. This carrier will extend the mission life from 12 days of

Spartan 200 series to 12 months. It is shuttle-launched and will be able to house large instruments (1 meter

plus and 2000 pounds). Since the carrier and the instrument are retrievable, it provides the space industry

the ability to obtain science quickly, inexpensively and at a lower risk.

Spartan 401 is an Earth pointer with half a dozen instruments on board, including LIDAR, STW/AR,

SHIMMER, GPSOS, OOAM, StOLSS. LIDAR is nadir pointing; STW/AR includes six antennas pointed

towards nadir; SHIMMER points towards limb and includes limited lunar track; GPSOS includes three

large antennas with specific orientations; OOAM is a limb pointer that tracks sunrise and sunset with

movable optics; StOLSS is a limb pointer that requires unobstructed FOV. The required orbit is 450 km x

450 km; the desired orbit is 600 km x 600 km. Inclination is 51.6 degrees. The attitude and pointing control

accuracy and knowledge varies. The first task for the ACS analysis team was to develop an error budget

bounding the performance capability based on the selected sensors and actuators for Spartan 400 series.

The analytic results were verified with simulations. Based on the analysis results we provided, ACS

imposed requirements on the other subsystem, such as optical bench thermal warping, solar array thermal

snap, etc. Ground tracking over +/-30 degrees was also studied. The ACS analysis team was also heavily

involved in the solar array design to ensure its orientation is attitude-control favorable, by investigating the

aerodynamic torque and its effect on wheels and torquers sizing for each configuration. In addition to the

analysis, ACS analysts also participated in the development of the product development plan in the area of

coordination systems definition, control modes and performance specification, ACS-derived requirement

definitions and hardware specifications.

Spartan 402 is an advanced solar coronal mission with two coaligned instruments: spectroscopic and

polarimetric coronagraph (SPC) mounted to a central mounting plate (CMP). A 10-meter-long deployable

mast also mounts to the CMP, positioning an external occulter assembly 10 meters forward of the SPC

aperture. Spartan 402 will be deployed at a 300 km attitude with 28-57 degree inclination and raised to 580

km. Because the 10-meter-long boon has a significant mass at the tip, the ACS analysis team performed

stability analysis and the tip mass deflection effect on the spacecraft attitude performance. ACS analysts

also assisted the structure and instrument team in defining the design requirement of the boon in terms of

structure frequency, damping ratio and modal gain. In addition to the conventional attitude sensors for the

Spartan 400 series, the fine attitude is from the instrument Guide Telescope (GT) with fine steering

mechanism that has an 8.5 x 8.5 arc-minute field-of-view. Since this GT is flown on TRACE, the GT FOV

searching algorithm can be adopted. However, detailed error budget was needed for the fine Sun

acquisition to ensure the GNCC’s capability to locate the Sun within the FOV of the TRACE guide

telescope/ISS. The analytical error budget result was crossed checked with a simulation. Additionally,

analysts participated in the development of the Product Development Plan in the area of coordination

systems definition, control modes and performance specifications, ACS-derived requirement definitions

and hardware specifications. A Spartan 402 proposal was submitted in July and an oral review was held on

August 19.

[Technical contacts: Josephine San/572; Jim Morrissey/572]

2.1.9 Triana

The past year has been an incredible challenge for the Triana ACS analysis team. This is a project with a

very tight schedule. The baseline design, SMEX-lite, a protoflight spacecraft designed in-house, was

conceived for low-Earth orbiting spacecraft. The ACS system required extensive modifications to

accommodate a mission to the L1 libration point. Working in a concurrent engineering environment, the
8

team is continually challenged to respond to constantly changing requirements. To mitigate the

requirements creep, the team created a flexible design capable of accommodating the observatory's

evolving requirements. We have successfully completed control system analysis for the Triana ACS, to the

satisfaction of the Triana project.

All of the onboard controllers have been designed and simulated, proving our performance meets our

requirements. Work is currently underway to support software implementation of our controller design. At

the same time, the team is preparing to support in-house integration and testing of the Triana spacecraft,

beginning in the fall of 1999.

Triana is the latest mission planned to visit the Earth-Sun libration point. From L1, Triana will have a fulldisk, sunlit view of the Earth at all times. Triana will provide pictures of the full sunlit Earth on the Internet

every 15 minutes. FDAB took Triana from a mission concept to a mature mission design in just a few

months. FDAB worked with Triana PI’s to alter the baseline mission trajectory to better meet science

objectives. The current mission includes a Lissajous orbit about L1, a shorter transfer time to L1 and

improved Sun-Earth-vehicle angles. The baseline trajectory is the first Lissajous orbit to be designed using

Dynamical Systems Theory (DST) from Purdue University. DST is an analytic approximation of classes of

orbits that intersect L1 and the Earth. DST greatly improves the efficiency of the mission design process

and the fuel-efficiency of the trajectory design. FDAB worked closely with ACS and propulsion engineers

to design an ACS and propulsion system that would meet mission requirements. This included number and

placement of spacecraft thrusters, fuel tank sizing and attitude modes. FDAB worked closely with the

Johnson Space Center on space shuttle launch and deployment scheduling. Triana requires four deployment

opportunities for any launch time, with a crew sleep cycle in between, and within ground station coverage.

Figures 2-1/-2/-3 represent views of the Triana spacecraft and the orbital trajectory.

Figure 2-1 Figure 2-2
9

Figure 2-3

For complete details about the Triana mission refer to http://triana.gsfc.nasa.gov/home/

[Technical contacts: MarkBeckman/572;Greg Marr/572; Wendy Morgenstern/572]

2.1.10 WIRE Flight Dynamics Attitude Support

The FDAB successfully supported the Wide-Field Infrared Explorer (WIRE) launched on March 5, 1999.

Shortly after launch, the spacecraft started spinning out of control due to the science instrument cover being

blown early. The pyrotechnics controlling the hatch opening malfunctioned and the heat from the Sun in

conjunction with onboard cryogen conspired to create an unanticipated thruster which was beyond the

capabilities of the attitude actuators to control. FDAB personnel provided attitude determination and rate

estimation support as well as providing various unplanned planning products until the spacecraft was

successfully transitioned to pseudo-normal operations though with a severely damaged science instrument.

Once in normal operations, FDAB personnel performed a calibration of the gyros. An example of the precalibration results can be seen in figure 2-4, and the resulting post-calibration results can be seen in figure 2-5.
10

Figure 2-4 WIRE Gyro Pre-Calibration Result
11

Figure 2-5 WIRE Gyro Post-Calibration Result

[Technical contact: Rick Harman/572]

2.2 Operational Missions

2.2.1 ERBS Earth Radiation Budget Satellite

The ERBS spacecraft, which was launched in 1984, experienced an under-voltage condition on battery #2

due to cell failures. This failure, which occurred on January 16, 1999, caused the ACS to autonomously

enter the B-Dot mode. Note that this was the first time that ERBS ACS had ever autonomously entered BDot mode. Following the successful commanding of the spacecraft back to normal ACS pointing mode, a

full-scale effort was made in order to recondition battery #1, which was disconnected in 1993 due its own

cell failures. Resolution of this anomaly has returned ERBS to a safe power configuration, thereby enabling

the ACS to provide the required attitude pointing as well as the periodic 180-degree yaw maneuvers. This

has allowed for the resumption of science-gathering activities.

[Technical contact: Samuel J. Placanica/572]

2.2.2 Lunar Prospector

Lunar Prospector (LP) completed its primary mission on December 31, 1998. The LP Extended Mission

was authorized for 6 months and included a lower altitude above the lunar surface. On January 19, 1999,

GNCC planned the maneuvers to lower the altitude to a mean of 40 km. After 1 month, the altitude was

again lowered to a mean of 30 km. GNCC planned all spacecraft maneuvers, including monthly
12

maintenance maneuvers required to keep LP in its orbit. GNCC also performed all orbit determination

including maintaining a web site with all LP definitive ephemerides (fdd.gsfc.nasa.gov/lp) for instant

access by mission controllers or PI’s. On July 30, 1999, the first of two final maneuvers was performed to

raise apoapsis to target the final impact of LP into a permanently shadowed crater on the lunar southern

pole. On July 31, 1999, LP’s final maneuver was performed on the dark side of the Moon and LP was never

heard from again. GNCC provided complete flight dynamics support to the operations center at

Ames Research Center. As part of targeting the final impact, improved lunar topography was needed.

GNCC worked with ARC, University of Texas, GSFC Code 900, and the Smithsonian to obtain the best

estimates of lunar topography around the impact site. All indications are that the impact was on target.

Analysis of the final impact and the determination of water content are still ongoing.

Figures 2-6/-7 represent the final trajectory.

Figure 2-6

Figure 2-7

[Technical contacts: David Folta/571;Mark Beckman/572]
13

2.2.3 TRMM Tropical Rainfall Measuring Mission

The TRMM ACS has been successfully providing the required performance with respect to attitude

pointing and maneuver capabilities throughout FY 1999. This includes the periodic 180-degree yaw

maneuvers to maintain thermal constraints, as well as orbit-raising maneuvers using thrusters in order to

keep the spacecraft at its 350-km orbit altitude.

During FY 1999 a number of spacecraft issues required attention. The first of these issues was that the

drive for the minus Y solar array had been experiencing temperatures above the qualified test levels since

early mission operations. These high temperatures could result in a failure of the array drive, thereby

stopping the array. This situation would create a larger-than-designed-for aerodynamic torque environment

which would result in higher momentum unloading requirements for the ACS. The failure would also effect

the capability of the power subsystem. A long-term study was undertaken by various spacecraft

subsystems, including ACS, to see what could be done. The first action of the study was to reduce the range

of motion for the arrays in order to decrease the integrated angular motion for the minus Y array drive. The

study then investigated the possibility of “parking” the minus Y array at an angle that would satisfy power

requirements with the least amount of impact to the ACS with regards to attitude pointing and momentum

unloading. Parking the minus Y array at an angle of 30 degrees was deemed acceptable by the investigation

group; however, the group’s recommended course of action was to not park the array. This

recommendation was based upon the benefits as well as the risks to spacecraft operations of a parked array

configuration. Results of the study were presented to a review panel at which the panel agreed with the

investigation’s findings and recommendations. Procedures and actions have been designed for future

operations in case it is deemed necessary to park the array, or if the array drive does malfunction.

The second spacecraft issue for TRMM during FY 1999 was that a thermally induced flexible mode of the

solar array boom was magnified to unexpected levels at high solar beta angles (54 to 58 deg). This resulted

in a noisy gyro signal which is used to drive the magnetic torquer bars (MTB) for momentum unloading

purposes. A software test within the ACS fault, detection and correction (FDC) logic switched from the

nominal 50/50 usage of each of the two coils for each bar to 100% usage of the backup coil. Once it was

realized that the MTB’s were not at fault, the FDC test threshold was increased and nominal 50/50

commanding of the MTB was re-enabled. The noisy gyro is also used to compute solar array drive

commands, resulting in noisy commands to the array drive. This “closed-loop” situation may be reduced

under a proposed ACS code patch which is currently under investigation.

[Technical contact: Samuel J. Placanica/572]

3.0 Study Missions Support

In FY99, the GNCC continued its participation in supporting a wide variety of future mission concepts.

This section describes the analyses performed which range from science instrument feasibility studies to

full mission concept definitions.

3.1 IMDC Integrated Mission Design Center

The Integrated Mission Design Center (IMDC) is dedicated to innovation in space mission design &

advanced concepts development to increase value for NASA and its customers. In the IMDC, subsystem

engineers (e.g., mission design, attitude control, propulsion, mechanical, communications, thermal,

command and data handling, power, and ground systems) gather to conceptualize and perform mission

trade studies in the presence of the customer (i.e., Project Scientist, Principal Investigator, etc.). The end

result is a spacecraft design and an operations concept that the customer can use in proposal development

for new mission opportunities (i.e., Announcements of Opportunity).

The IMDC became operational in spring 1997 and as of this writing has supported over 50 mission studies

and focus sessions for both Earth science and space science applications. A typical IMDC study session

begins with a mission design briefing by the customer. The briefing is followed by design sessions to
14

perform the system and subsystem level trades agreed upon with the customer. Typical design sessions last

from 1–2 weeks. On many occasions, multiple visits to the IMDC are needed as a mission concept matures.

GNCC analysts support the IMDC using their expertise in trajectory design, orbit analysis, and mission

planning as well as ACS design, ACS hardware selection and performance evaluation. A sample of the

IMDC missions supported by the GNCC in FY99 included:

ÿ Coriolis: Coriolis is a Naval Center for Space Technology (NCST) mission designed to measure

surface level wind direction and speed over the ocean using a microwave polarimeter. The NCST

used the IMDC on two occasions as a simulated Rapid Spacecraft Development Office (RSDO)

vendor to evaluate their requirements specifications document. GNCC analysts analyzed the ACS

requirements, orbit selection, and maneuver requirements, and assisted in the operations concept

analysis.

ÿ NPOESS Preparatory Project (NPP): NPP intends to demonstrate and validate pre-operational

instruments and algorithms prior to first flight of NPOESS (NOAA Polar Orbiting Environmental

Satellite System). NPP visited the IMDC twice in FY99. The first visit was to evaluate the

requirements for creating a baseline design. The final visit was to evaluate the design against

potential spacecraft found in the RSDO catalog.

ÿ EO-3 Nexus/RedEye: The IMDC supported 3 weeks of analysis on the Nexus/RedEye

collaboration for the EO-3 competition. GNCC analysts supported the operations design concept

of moving Nexus/RedEye from a low-Earth, shuttle orbit to geostationary orbit. Trade studies

were performed on different propulsion methods (solid/liquid, all liquid, solid/solid) and in the

end, a dual-solid motor architecture was adopted. ACS work involved analyzing the two different

observing modes of this proposal— Earth pointing for RedEye and stellar pointing for Nexus, and

working with instrument control engineers and PI’s to establish necessary ACS accuracy and jitter

requirements. ACS analysts also devised necessary control modes for the different phases of the

mission and recommended the necessary components for each mode. Trade studies were

performed on one star tracker versus two; gyro selections to accommodate extreme high rate and

low rate; four reaction wheels versus three; and Global Positioning System (GPS) options for

future technology research.

ÿ MAXIM: The Micro-Arc-second X-ray Imaging Mission (MAXIM) involves the formation flying

of separate optics and detector spacecraft to collect x rays in a heliocentric drift-away orbit. The

mission concept relies on an entirely new concept of formation flying. GNCC analysts worked to

try to validate the MAXIM concept. In order to achieve their x-ray science, the MAXIM PI

requires a 450 km focal length between his optics and detectors. It is desired to obtain this focal

length by formation flying the optics and detectors as two independent spacecraft. The challenge is

to formation fly the detectors around the optics at a specific range and maintain this range and

attitude to point at celestial targets for observations on the order of 7 days. When a new target is

selected, the optics will slew to the target but the detectors must correct its orbit to move in space

relative to the optics and maintain the 450-km focal length. Work will continue outside of the

IMDC to simulate the formation-flying concept. Another challenge of the MAXIM mission is the

0.5 milli-arc-seconds pointing accuracy requirement for the optics spacecraft and the 3-mm lateral

pointing accuracy requirement for the detector spacecraft. ACS analysts worked with the PI to

derive ACS pointing accuracy and jitter requirements, performed simulations to assess feasibility

to adopt NGST design for coarse pointing, and analyzed fine pointing using the instrument sensor

as an attitude sensor to close the loop. ACS modes for both the optics and detector spacecraft were

derived and sensors for each mode were selected. Trades were performed on actuator selection

(reactions wheels, pulsed plasma thrusters, or both) for rate null/Sun acquisition, science, and slew

control.

ÿ DS-5 Collaboration: GNCC analysts supported the IMDC effort to evaluate three potential DS-5

missions as part of NASA’s New Millenium Program (NMP). Work included mission concept

brainstorming of the three options— a solar sail demonstration, a force-free spacecraft pair for
15

detecting gravity waves, and a nanosat concept to evaluate miniaturization of instrument and

hardware technologies.

ÿ MFV: The Maxwell, Faraday, Vlasov (MFV) is a 1-month mission designed to directly measure

electric currents using the three independent techniques proposed by physicists’ whose names

comprise the project name. One of the three techniques involves the release of 12 miniature freeflying magnetometers (FFM’s) prior to entry into high auroral activity at three times during the

mission. The main spacecraft would receive data from the FFM’s over one-half to three orbits.

GNCC analysts worked with the PI to examine FFM release velocities and subsequent relative

motion. Minimizing the relative range between the FFM’s and the main spacecraft maximized the

science return. Attitude control work involved defining spin axis orientation, spin rates, control

modes and functions of each control mode. ACS analysts specified the locations and alignment

requirements for magnetometers and magnetic coils. Sizing of spin and precession coils was based

on the FAST mission. Since MFV is a slow spinner, managing the inertia ratio is critical. ACS

analysts worked closely with the structure designer in the arrangement of the hardware. This

mission was made more challenging by the requirement to fly as a secondary payload.

ÿ Extremely Heavy Galactic Cosmic-Ray Composition Observer (ECCO) & Cosmic Ray

Energetics and Mass (CREAM): In addition to Earth-orbiting free-flyers, IMDC work has

included the support of two balloon payloads designed to measure cosmic rays. GNCC support has

primarily been in the area of identifying ACS components to meet the instrument pointing

requirements. Further analysis was performed to understand the implications of the balloon

disturbance environment. Analysts also supported communications link studies.

[Technical contacts: Jennifer Bracken/571; Michael Mesarch/572; Dave Olney/572; Josephine San/572;

Frank Vaughn/572]

3.2 Mission Concept Support

In addition to IMDC support, GNCC analysts supported other varied mission concept studies:

Sun-Earth Connections (SEC) Roadmap: GNC mission analysts supported mission design activities for

the SEC Roadmap. In all, 12 mission concepts were examined for methods of collecting the desired science

data. The 12 missions included conventional, low-Earth orbit, geosynchronous orbit, highly elliptical Earth

orbit, non-Keplerian ‘pole-sitters’ above the Earth, and orbits at or inside Mercury’s orbit (0.39

astronomical units). The science data capture was accomplished with single satellites or constellations of up

to 40 satellites. In supporting the roadmap, it was apparent that many missions would benefit greatly from

advances in solar sail technology and miniaturization of instruments and components.

Global Electrodynamics Connections (GEC): The GEC mission involves taking in situ measurements of

the Earth’s ionosphere using four identical spacecraft. The four spacecraft will fly in formation (as close as

8 km) to examine the variability in the ionosphere. As these spacecraft fly at such low altitudes,

atmospheric drag will continually cause the orbit to decay. Mission analysts continue to examine the

amount of propellant necessary to provide for ample science return during a 2-year mission. Further

analysis on the launch initial conditions and evolution of the perigee latitude and local time history will

help the GEC science team to plan for interesting and varied ionospheric conditions (e.g., combinations of

season, day vs. night, auroral vs. mid-latitude, etc.). GEC has also visited the IMDC on three separate

occasions for iterations on the spacecraft design. During these visits, ACS engineers have continued to

refine hardware selection. Included in the analysis is the feasibility of designing an Integrated Power and

ACS (IPACS) flywheel system— a new technology development.

Magnetospheric Multi-Scale (MMS): The MMS mission is designed to investigate the Earth’s

magnetosphere in the regions of the magnetopause and plasma sheet. Five spacecraft will fly in a loose

formation that will form a double tetrahedron at orbit apogee. The mission consists of four phases in which

the orbit ranges between 1.2 x 12 Earth radii (at 10-degree inclination) and 10 x 50 Earth radii (at 90-

degree inclination). Flight dynamics analysis support for the past year has consisted primarily of conferring

with the Project Manager and Project Scientist to formulate mission science requirements and spacecraft
16

operational requirements and performing analysis to identify and resolve conflicts as related to the orbit

design. Analysis to determine the effect of launch epoch and initial clock angle, with respect to the

magnetosphere, on maximum shadow length and dwell time in the plasma sheet during Phases 1 and 2 is

nearing completion. Analysis tasks for the upcoming year will include the design of the two sets of double

lunar swingbys (Phase 3) required to transition to the Phase 4 orbit and development of an efficient control

strategy to maintain the double tetrahedron at orbit apogee throughout the mission life.

Space Weather Concepts: GNCC analysts provided inputs to GSFC scientists for a Space Weather

Workshop in April 1999. Analysts assisted in assessing feasibility of several different mission concepts.

The missions included a solar-polar orbiter, simulated solar comet, and a distant retrograde orbit (DRO)

constellation. Interest in a four spacecraft DRO, to monitor solar weather with 10 times the response time

of SOHO, has prompted a future paper in The Journal of Atmospheric and Solar-Terrestrial Physics.

[Technical contacts: Greg Marr/572; Michael Mesarch/572; Karen Richon/572; Frank Vaughn/572]

4.0 Technology Development Activities

4.1 Advanced Mission Design (Dynamical Systems Theory)

Dynamical Systems Theory (DST) is being used by the GNCC to design trajectories for libration point

orbits. DST techniques improve the mission design process by allowing one to first choose the desired

mission orbit and then work backwards to find the transfer trajectory. Traditional mission design

techniques, as applied to libration point orbits, often rely on inefficient “hit or miss” shooting techniques

requiring numerous simulations.

The DST work has been performed in partnership with Purdue University. Purdue graduate students, in

conjunction with their university advisor, Dr. K. C. Howell, have made significant contributions to the

GNCC’s mission design efforts. Their work on the Triana trajectory design is especially noteworthy; it

would be extremely difficult, if not impossible, to evaluate the numerous candidate Triana trajectories,

without the use of DST techniques.

Mark Beckman of the GNCC and Jose Guzman of Purdue University presented a paper on the Triana

Mission Design, which included a discussion of the DST techniques involved, at the Alaska AAS

conference.

[Technical contact: Steven Cooley/572]

4.2 Autonomous Relative Navigation & Formation Flying

In October 1998, NASA/GSFC initiated efforts to coordinate the formation-flying and virtual platform

technology development programs. In parallel with these efforts, the Cross-Enterprise Technology

Development Program (CETDP) was restructured to create 10 technology Thrust Areas (TA). As a result of

these changes, the GSFC formation-flying program gained a larger role within the agency. GSFC was

selected as the lead center for Distributed Spacecraft Control. Working in close cooperation with the

CETDP Thrust Area Manager (TAM) for Distributed Spacecraft Control, Kate Hartman, GSFC expanded

the formation-flying program and broadened its scope.

One of the major highlights of FY99 was the creation of a partnership between NASA and the Air Force

Research Laboratory (AFRL). AFRL has instituted a program to push technology development to support

their TechSat 21 mission. This program involves 10 universities developing and flying nanosatellites in

various formations. AFRL asked the GSFC formation-flying team to help evaluate the university

submissions they received for the nanosat program. In particular, AFRL wanted GSFC to take the lead role

for formation-flying aspects of the program and to rank the proposals accordingly. The formation-flying

team, with the help of CETDP Distributed Spacecraft Control Thrust Area, decided to take this a step

further and augment the selected university funding to develop technologies of specific interest to NASA.

In May, a Request For Information (RFI) was posted seeking proposals that would develop underlying

technologies necessary for formation flying. As a result of the RFI, 10 interesting technologies were
17

selected and paired with spacecraft to provide a flight opportunity. Currently, the universities are

continuing with AFRL to plan the missions and design their spacecraft with NASA/GSFC/GNCC

technologies incorporated.

The GSFC formation-flying team wrapped up the Earth Science Technology Office (ESTO)-funded studies

in formation control. These studies included analysis by Johns Hopkins University/Applied Physics

Laboratory (JHU/APL), AI Solutions, and Stanford University. The JHU/APL work looked at defining the

general requirements of formation-flying missions. The AI Solutions work included incorporating

formation-control techniques and modeling in FreeFlyer. The Stanford study examined using differential

GPS to determine high-precision spacecraft relative positioning. ESTO was not able to fund subsequent

work in this fiscal year. However, a follow-on effort was instituted thanks to funding from CETDP. This

work at the Massachusetts Institute of Technology (MIT) includes analysis and algorithm development to

specify the number of spacecraft needed for any particular interferometry missions.

In December 1998, the formation-flying team supported the first CETDP Distributed Spacecraft Control

Workshop. The workshop was an invitation-only affair to exchange program information, progress and

ideas. The workshop was a great success. It served as a kickoff for the CETDP and the materials from the

workshop are still in high use today. It was the first opportunity to get most of the NASA-distributed

spacecraft players together in one place to exchange ideas.

During FY99, the Formation-Flying Team completed the design, development and testing of the EO-1

Enhanced Formation Flying (EFF) flight code. With this code and the incorporation of the flight version of

AutoCon, also completed during this fiscal year, the EO-1 spacecraft will be able to autonomously maintain

formation with Landsat-7 using multiple algorithms for control. Currently, GSFC and the Jet Propulsion

Laboratory have algorithms that will be used during the mission. However, the flight system has been

designed such that other algorithms could be uploaded after launch to further test and enhance the

formation control of EO-1.

As FY99 went on, the FDAB began developing the Formation Flying Testbed (FFTB). The FFTB is a joint

project between GSFC, the Hammers Co., JHU/APL and AI Solutions. Funding for this effort includes

SBIR’s, CETDP and SOMO money. The testbed will include GNCC’s GPS testbed to create an

environment that allows algorithms, hardware and other technologies to be developed and tested for

formation control. The testbed uses the Hammers Co.’s VirtualSAT Pro to provide a dynamic simulator for

a spacecraft or series of spacecraft as well as a hardware interface for any test hardware. AI Solutions’

FreeFlyer provides a graphical display for the testbed.

Also during FY99, the GSFC formation-flying team supported the Technology Development for Explorers

Missions NASA Research Announcement (NRA). The team supplemented the NRA with additional

funding targeted specifically for formation-flying efforts. The team also submitted some proposals to the

NRA. A submission by Russell Carpenter for Decentralized Estimation and Control of Distributed

Spacecraft was accepted and funded. This effort includes using the FFTB to test Dr. Carpenter’s control

algorithms and provides a small amount of additional funding for the FFTB. Also, a submission by

JHU/APL to develop a crosslink-transceiver was selected. This transceiver will be flown on board some of

the university nanosats. Also, a submission by Jim Garrison for development of a GPS receiver suitable for

use by formations of satellites in highly elliptical and geosynchronous orbits was accepted and funded.

This effort will make use of the FFTB to test GPS relative navigation algorithms.

The GSFC formation

from March through May. This effort included reviewing the entire set of EOS, EX and OP missions and

suggesting technologies that could be used on certain missions and the benefits of those technologies.

These technologies were included in the final submission to 

EOS missions.

The formation Submillimeter Probe and the Evolution of Cosmic

Structure (SPECS) mission development. Team member Dave Quinn provided the SPECS team with

spacecraft design suggestions to meet their complicated mission requirements. Dave’s design allowed the
18

spacecraft to accommodate the complex dynamics of multiple collectors moving in tight and drifting

further out. The SPECS mission continues to evolve and has very changing requirements.

The formation flying team continues to support the Earth Science Vision. Team members have participated

in numerous workshops and presentations and provided input to the vision. In particular, the team has been

asked to support the web sensor portion of the vision.

The formation-flying team supported the ST-5 proposal development. The team provided input into the

technologies and hardware necessary to support formation flying or constellation control among the

nanosats. The proposal was accepted and the team anticipates the need to meet the challenge of the mission

design, and incorporate formation-control technologies in a very tight budget and limited spacecraft

capability.

Finally, the formation-flying team in cooperation with other NASA groups submitted their own proposal

for the EO-3 OA. This proposal defined a formation-flying mission of up to 11 nanosats to do soil moisture

and precipitation monitoring. Unfortunately, the proposal was rejected, but the team did gain valuable

experience in the process and generated research that has been used to support other missions planning to

use interferometry.

[Technical contact: John Bristow/572]

4.3 Advances in Navigation Technology

Increasing interest in maximizing autonomy, operations “beyond LEO,” and distributed spacecraft has

brought new challenges for navigation systems. Addressing and anticipating new requirements has led to

technology initiatives in onboard navigation systems (ONS) using communications links, GPS orbit

determination, and autonomous navigation for high-Earth, libration, gravity-assist, and deep-space orbits.

The continuing needs of flight projects and future mission studies for navigation consulting have also been

supported.

Recognizing the benefits of technology commercialization, in-house ONS and GPS technology has been

transferred to Motorola and Orbital Sciences Corporation, and a provisional patent has been filed for a new

type of GPS attitude determination sensor1. A web site2 has been launched, currently focusing on GPS and

ONS, that is intended to raise awareness among the GNCC community of navigation activities at GSFC,

and stimulate further commercial utilization of NASA-developed technologies. Helping to ensure

continued support of in-house development, funding from the Explorers and Cross-Enterprise Technology

Development programs have been secured for GPS and ONS initiatives in the areas of formation flying,

and highly elliptical and geostationary orbit applications of GPS.

4.3.1 Onboard Navigation Systems Using Communications Links

The Enhanced ONS (EONS) flight software package, schematically illustrated in figure 4-1, has been

delivered to Motorola for integration with a navigation processor board. Motorola is currently developing

the flight software interfaces and looking at board improvements for improving Doppler data accuracy and

frequency independence. A mathematical specification for EONS that describes the algorithms associated

with the integration of ONS capabilities with a spacecraft’s ground systems communications receiver has

been delivered, and a draft System Description and User’s Guide (SD&UG) has been prepared.

 

1

 Compound Eye GPS Attitude & Navigation Sensor (CEGANS), patent number GSC 13966-1.

2

 http://geons.gsfc.nasa.gov
19

ANTENNA

SPACECRAFT 

CLOCK

GROUND 

STATIONS

FREQUENCY 

SPACECRAFT OPERATIONS 

SUPPORT FACILITIES

TELEMETRY

USER

DOPPLER 

MEASUREMENT 

CAPABILITY

• ONS PERFORMANCE

MONITORING AND 

CALIBRATION

• ONS QUALITY

ASSURANCE

RECEIVER

CONTROL

ATTITUDE 

CONTROL

SPACECRAFT COMPUTER

 GS RECEIVER/TRANSPONDER

TELEMETRY

UPLINK SIGNALS

ONS

FSW

TIME

COMMAND AND 

TELEMETRY 

INTERFACE UNIT

SERIAL

INTERFACE

 UNIT

RECEIVER COMMANDS, 

ONS COMMANDS 

TELEMETRY

DATA 

RECEIVER 

TELEMETRY, 

USER STATE,

FREF, FCWs 

TIME

STABLE 

REFERENCE

(FREF)

COMMAND

 DATA

 Figure 4-1

[Technical contact: Cheryl Gramling/572]

4.3.2 Global Positioning System (GPS) Orbit Determination

Studies supporting formation flying and highly elliptical and geosynchronous orbit missions have been

accomplished. A major new release of the GPS Enhanced Orbit Determination (GEODE) flight software

has been delivered. These accomplishments are described below. In addition, a paper describing

relationships among semi-major axis, position, and velocity which are often not maintained in space

applications of GPS was presented to the American Astronautical Society3. Figure 4-2 illustrates these

results.

Figure 4-2

GPS State Vector Differencing. The accuracy of relative navigation solutions that can be achieved by

differencing independently computed absolute state vectors, for a formation of eccentric, medium-altitude

Earth-orbiting satellites using GPS measurements has been assessed. This scenario is shown in figure 4-3.

Assessments show that relative state accuracy is five to six times worse when the differences are performed

on deterministic point solutions rather than filtered solutions, and degrades rapidly as the number of

measurements from uncommon GPS satellite vehicles (SV’s) that are used in the solutions increases. The

best accuracy (2 meters, 0.001 meter per second, root-mean-square (rms), and 10 meters, 0.005 meter per

second, maximum) is achieved when four or more GPS SV’s are mutually visible, and if the point solutions

are computed using only measurements from GPS SV’s common to both user satellites.

 

3

 Carpenter, J. Russell and Emil Schiesser, “The Importance of Semi-major Axis Knowledge in the Determination of Near-Circular

Orbits,” Space Flight Mechanics 1999, Univelt Publishing, San Diego.
20

Satellite 1

• GPS Tracking

• Absolute State Vector

 Estimation

• Absolute State Vector

 Differencing for

 Relative Navigation

• Formation Maintenance

 State Vector N

Via Comm Crosslink

GPS

GPS SV Navigation

Messages,

Pseudoranges

 State Vector 1

Via Comm Crosslink

Satellite N

• GPS Tracking

• Absolute State Vector

 Estimation

• Absolute State Vector

 Differencing for

 Relative Navigation

• Formation Maintenance

Figure 4-3

GPS in Highly Elliptical Orbits. The accuracy of navigation solutions computed for an eccentric mediumaltitude Earth orbit (MEO) satellite as a function of GPS SV visibility has been investigated in a

cooperative study with the University of Colorado at Boulder. The orbit studied is a geosynchronous

transfer orbit, which has an apogee altitude that is well above the altitude of the GPS constellation. The

analysis reported indicates that, with the increased visibility provided if lower strength signals from the

GPS SV antenna side lobes can be acquired, navigation accuracy for a geosynchronous transfer orbit

improves significantly, as shown in Figure 4-4. The rms total position errors reduce from 100 meters to 40

meters and the peak errors near apogee reduce from 250 meters to 100 meters. These results were reported

to the Institute of Navigation’s Annual Meeting4.

0

50

100

150

200

250

1.75 2.00 2.25 2.50 2.75

Elapsed Days from Start

Total Position Error (meters)

25 dB-Hz

35 dB-Hz

Figure 4-4

GEODE. GEODE provides the capability to use the GPS Standard Positioning System (SPS) to provide

high-accuracy orbit determination and time autonomously onboard NASA spacecraft. For flight

applications, the GEODE flight software can be integrated either within a GPS receiver or hosted on the

spacecraft’s onboard computer (OBC). For ground-processing applications, the GEODE flight software can

be executed using GEODE ground-processing software. The current version includes revisions to

incorporate algorithm enhancements that have been implemented in GEODE versions 4.01, 4.02, and 4.03.

These enhancements include the following capabilities: estimation of both a correction to the atmospheric

drag coefficient and the solar radiation pressure coefficient, extension of the gravity process noise model to

handle eccentric and geostationary orbits, reformulation of the receiver clock bias model, and processing of

measurements from all GPS space vehicles in view. Mathematical specifications and SD&UG documents

reflecting these enhancements have been delivered.

[Technical contacts: Russell Carpenter/572; Jim Garrison/572]

4.3.3 “Compound Eye” GPS Attitude & Navigation Se nsor (CEGANS)

A novel GPS is under development which would provide data for both navigation and attitude. The sensor

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 is attached) can be estimated to within 3 degrees without

 

4

 Moreau, M., et al. “GPS Receiver Architecture and Expected Performance for Autonomous GPS Navigation in Highly Eccentric

Orbits,” Proceedings of the Annual Meeting of the Institute of Navigation, June 1999.
21

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 subdegree range.

A paper was presented on this invention at the AIAA Boston conference last year5 and was elected best paper of

the session. The CEGANS was featured in a NASA Tech Briefs article in December 1998, and a provisional

patent has since been issued by the patent office while a formal patent application awaits submission.

[Technical contact: David Quinn/572]

4.3.4 Autonomous Navigation for High-Earth, Libration, Gravity-assist, and Deepspace Orbits

An effort has been initiated to characterize the accuracy of autonomous navigation systems for satellites in

regimes in which use of the GPS or Tracking and Data Relay Satellite System (TDRSS) is not feasible.

This work assesses the feasibility of using standard spacecraft attitude sensors and communication

components to provide autonomous navigation for high-Earth orbit (HEO) and libration-point orbit (LPO)

missions, as illustrated in figure 4-5. Performance results are presented as a function of sensor measurement

accuracy, measurement types, measurement frequency, initial state errors, and dynamic modeling errors.

This analysis indicates that real-time autonomous navigation accuracies ranging from 100 meters rms for a

3-by-20-Earth-radii HEO satellite to 10 kilometers rms for an LPO satellite are achievable using highaccuracy attitude sensor and one-way Doppler measurements. This work was presented to the American

Astronautical Society’s Astrodynamics Specialists Conference6.

Autonomous Navigation Scenario

S/C to Sun

directional

measurement

S/C to Earth

directional

measurement

Ground Station

to S/C Doppler

measurement

Figure 4-5

[Technical contacts: David Folta/571; Cheryl Gramling/572]

4.3.5 Navigation Studies Supporting Flight Projects and Future Missions

Flight projects and future missions supported included ICESAT, various magnetospheric survey missions,

and the International Space Station version of the Honeywell Space Integrated GPS/INS (SIGI).

 

5 Crassidis, J.L, & Quinn, D.A., et al. “A Novel Sensor for Attitude Determination Using Global Positioning System Signals,”

Proceedings of the American Institute of Aeronautics and Astronautics, September 1998.

6 Gramling, C., et al. “Autonomous Navigation Using Celestial Objects,” AIAA Astrodynamics Specialists Conference, August 1999.
22

ICESAT. Expected orbit determination accuracies were estimated for candidate orbit determination

scenarios with Selective Availability (SA) on and off. Raw geometric point solutions clearly will not

provide the desired accuracy levels. Smoothing of point or filtered solutions significantly improves radial

position and in-track velocity accuracy. If SA is on (current GPS operational configuration), the smoothed

solution and post-processed differential solution can provide radial position accuracies of <10 meters and

in-track velocity accuracies of <1 centimeter per second after 24 hours. Of these two methods, the

smoothed solution method is considerably simpler to implement. If SA is off (expected GPS operational

configuration in the 2000–2002+ timeframe), the filtered, smoothed, and post-processed differential

solutions can provide radial position accuracies of <5 meters and in-track velocity accuracies of <1

centimeter per second after 24 hours. Of these three methods, the filtered and smoothed solution methods

are considerably simpler to implement than the post-processed differential approach.

Magnetospheric Survey Missions. Orbit determination error analysis has been performed to determine the

predicted accuracy for three magnetosphere-mapping missions: a mission that would use a constellation of

250 small satellites (“Magneto”), and two phases of a mission that uses a four-satellite formation (Magnetic

Multiscale mission, MMS).

For the Magneto mission, each satellite would have a perigee height of 1.4 Earth radii. The apogee height

values would range from 5 to 25 Earth radii. The inclination varies from 0 to 15 degrees. Tracking data

consists of one-way return Doppler data from 11-meter S-band ground stations at Hawaii, Puerto Rico,

Hartebeesthok, Diego Garcia, and the three DSN sites at Canberra, Goldstone, and Madrid; tracking station

angle data could also be utilized. The analysis shows that 14-day tracking arcs with Doppler-only data will

meet the mission’s 3000 km position requirement.

MMS Phase I will require knowledge of absolute position and inter-spacecraft range to 100 km in its 1.2 by

12 Earth radii, 10 degree inclination orbit. Tracking data consisting of 30 minutes per orbit of Doppler data

using Universal Space Network (USN) X-band ground stations located at Hawaii, Wallops Island, and

Poker Flats, and both coherent and noncoherent Doppler data could be utilized. When using coherent

Doppler data, an 8-day tracking span length should be used with at least one pass near perigee each day.

When using noncoherent Doppler data, the frequency bias and drift should be estimated using an 8-day

tracking span length with at least one pass near perigee each day. With respect to the inter-range accuracy,

it has been determined that maximum error between two spacecraft was 60 meters when using a minimum

tracking station elevation angle of 15 degrees, and 127 meters when using a minimum tracking station

elevation angle of 7 degrees. It should be mentioned that the latter results correspond to a 97% to 99%

correlation between the nonrandom errors common to the satellites, which is somewhat higher than the

90% to 95% correlation typically expected.

MMS Phase III differs from Phase I in having an 8x80 Earth radii orbit, and 28.5 degree inclination. Both

DSN and USN tracking stations could be utilized. Due to the long period of this orbit (approximately 17

days), an 18-day tracking span has been studied. Results show that the DSN-only cases will not meet the

100 km requirement. The cases with the additional USN tracking will meet the 100 km requirement. The

definitive accuracies using the DSN and USN tracking vary from 2 to 31 km. At least two perigee passes

per data span are recommended.

SIGI. The SIGI provides an attitude determination capability that has been developed in-house. Support to

this effort has included ground support equipment and performance analysis software development,

consultation for and development of flight software, troubleshooting support to resolve coordinate systems

and interface issues with both the GPS Test Facility’s GPS signal simulator and the ISS flight software, and

attitude performance analysis. Support has been provided for SIGI flight software releases 4.11 through

4.23. Consultation in the areas of attitude performance evaluation, SIGI software issues, and mission design

has also been provided to JSC in support of a space shuttle flight test of the SIGI to occur on STS-101.

[Technical contact: Russell Carpenter/572]
23

4.4 Attitude Determination

4.4.1 Advanced Attitude Analysis

Additional studies were conducted on a Kalman Filter designed to estimate attitude and rate using

quaternions from a quaternion output star tracker. The results of this testing indicate that this Kalman Filter

is capable of providing highly accurate rate estimates eliminating the need for an expensive gyro. This

algorithm was applied to flight data from the Rossi X-Ray Timing Explorer (RXTE). The flight star tracker

data was converted to single frame attitude quaternions using the Q-method. The results were compared to

the onboard attitudes computed from an EKF with star tracker data and a high performance gyro package.

Figure 4-6 demonstrates the rate estimation capabilities during an RXTE maneuver. The rss of the error for

this particular span of flight data was 6.1839e-4 degrees/second.

Figure 4-6 RXTE True Rate Versus Rate Estimate

An improved magnetometer calibration algorithm was developed and tested on flight data from the RXTE

mission which had particularly difficult magnetometers to calibrate. Besides improved accuracy, this

algorithm is more efficient and provides a covariance of the resultant calibration. Figure 4-7 provides a

sampling of the magnetometer calibration for RXTE.
24

Figure 4-7 Magnetometer Calibration Using RXTE

Flight Data (Solid Line is Post-Calibration Results)

[Technical contact: Rick Harman/572]

4.4.2 Attitude Models Task

The SKY2000 task has provided support in the form of new star catalogs (MARGIE–Minute of Arc

Resolution Gamma ray Imaging Experiment–balloon mission/La. Tech), updated star catalogs (Landsat-7

& SWAS), consultation with mission personnel (Landsat-4/5, EOS AM/PM, UARS, SOHO, HST, GRO,

WIRE), and for other groups in the space science field (STScI, Toshiba, CDS, TRW).

Version 2 of the SKY2000 Master Catalog (MC) was completed in September 1998 and represents a

substantial improvement over Version 1 and a unique consolidation of astrometric data from the highest

quality sources available. Position and proper motion data were improved by at least one order of

magnitude and Johnson V&B magnitudes are now available for almost all of the ~300,000 stars in the MC.

More than 4000 stars observed by the Ball CT-601 star trackers on the RXTE spacecraft now have

observed CCD ST magnitudes in Version 2. IAU designations based on J2000.0 positions have been added

to the catalog, while all previous identifiers (SKY2000, HD, SAO, DM, HR, WDS, PPM, AG, bright- and

variable-star designations) have been retained. The IAU-approved identifiers allow new objects to be

inserted without disrupting the natural order of the principal catalog identifier.

The SKY2000 MC Version 2 is available for download from the Flight Dynamics Division (FDD) web site

and is in use by both the space hardware and the astronomy communities. Analysis conducted on the

observational data from the Ball CT-601 star tracker on RXTE has resulted in a substantial improvement in

the accuracy of the instrumental magnitude (Mi) prediction algorithm. That improvement, coupled with the

actual observed CT-601 magnitudes now present in the MC, will provide end users significantly better Mi

data for both onboard and ground-based attitude determination star catalogs.

[Technical contact: David Tracewell/572]
25

4.4.3 Magnetometer Navigation

GNCC personnel, jointly with Dr. Itzhack Bar-Itzhack of the Technion Israel Institute of Technology, have

developed and tested a low-cost system for simultaneous and autonomous attitude, orbit, and rate

estimation of a low-Earth orbit satellite. The system relies mainly on magnetometer data, a reliable and

low-cost sensor. The estimation is performed using an Extended Kalman filter (EKF) algorithm. The filter

state can be chosen to estimate all three components, namely attitude, orbit, and rate or a subset. This

algorithm was successfully implemented for the Transition Region and Coronal Explorer (TRACE)

mission. The TRACE sensor complement consisted of a magnetometer and a course (DSS). The results of

estimating all 13 states can be seen in figures 4-8/-9/-10. During FY99, analysis of the attitude and orbit

estimation subset was conducted to determine the feasibility of further improving the measurement data

through the incorporation of more sophisticated sensor noise models. Dr. Wallace Larrimore of Adaptics,

Inc. developed a statistical model of the RXTE magnetometer data. This model was incorporated into the

RXTE EKF, resulting in improved attitude and orbit estimates. The results were presented at the 21st

Annual AAS Guidance and Control Conference in a paper titled “Improved Modeling in a Matlab-Based

Navigation System.”

Negotiations with the CRV X-38 project at JSC began in FY99. The X-38 project is interested in

determining if the magnetometer can provide a backup attitude solution to the prime attitude system based

on GPS data. Attitude analysis using the magnetometer-based EKF will be conducted in Code 572.

Numerous studies with the EKF will be conducted to determine the convergence potential from an

unknown initial attitude within the time constraints of the necessary CRV orbit maneuvers. A preliminary

result is given in figure 4-11. With simulated data, the EKF overcame an initial attitude error of 115

degrees and converged to within 5 degrees in 5 minutes, meeting the CRV constraints. Concurrently with

the analysis, Code 573 will build the flight and test magnetometers for JSC based on the SMEX-lite

magnetometer design. Following the test flight of the X-38 vehicle, the recorded magnetometer flight data

will be analyzed in Code 570 and the magnetometer-based attitude estimates will be provided to JSC for

further evaluation.

Figure 4-8 RSS of TRACE Orbital Position Error Figure 4-9 RSS of TRACE Attitude Estimation Errors
26

Figure 4-10 RSS of TRACE Rate Estimation Errors

Figure 4-11 Preliminary CRV Magnetometer Attitude

Test Result— Initial Error was 115 Degrees

[Technical contact: Julie Deutschmann/572]

5.0 GNCC Flight Dynamics Laboratory

GNCC began configuration of the Flight Dynamics Laboratory located in Building 11. This lab will be

used for the development, test, integration, and operation of software systems to perform flight dynamics

functions and analysis in support of space missions. When fully configured the lab will have the capability

to receive spacecraft telemetry and navigation data, and have voice and video communication links for the

performance of these functions.
27

Items completed were the room construction, installation of voice and video equipment, and installation of

several computers. This initial configuration allowed support for the Lunar Prospector end of mission to be

done from the lab.

Plans for the coming year call for the installation of computers, installation of an on line data storage

device, and installation of servers. When this is complete the lab will be able to support all planned

functions as well as provide an area for GNCC software analysis instruction.

[Technical contact: Sue Hoge/572]

6.0 Interagency Activities

6.1 CCSDS— Consultative Committee for Space Data Systems

Felipe Flores-Amaya (Code 572) represents GNCC as the chairman for the P1J Subpanel under the

Consultative Committee for Space Data Systems (CCSDS).

The CCSDS is an international organization of space agencies interested in mutually developing standard

data handling techniques to support space research conducted exclusively for peaceful purposes. Ref.

http://www.ccsds.org/

Subpanel P1J is specifically chartered to investigate and recommend Navigation Data standards. P1J

currently has 16 members, representing the following international agencies: Chinese Academy of Space

Technology (CAST), China; Communications Research Centre (CRC), Canada; Centre National D’etudes

Spatiales (CNES), France; Deutsche Zentrum fur Luft- und Raumfahrt (DLR), Germany; European Space

Agency (ESA), Europe; National Aeronautics and Space Administration (NASA), USA; and National

Space Development Agency (NASDA), Japan.

The work of P1J is accomplished primarily at workshops, conducted at least twice a year, at facilities

coordinated by the hosting member agency. The main task of P1J is to generate documents defining the

preferred standards for the exchange of navigation data. A draft of the latest document was reviewed during

the most recent workshop, May 1999 at Newport Beach, California.

Future work will involve investigating

- Proximity and in situ navigation

- Interface with GPS, GLONASS, and similar systems

- Autonomous navigation

- Tracking stations on other planets

- GPS-like constellations around other planets

[Technical contact: Felipe Flores-Amaya/572]

6.2 Flight Mechanics Symposium

The FDAB conducted the Flight Mechanics Symposium, May 18–20, 1999. The purpose of this

symposium was to provide a forum for specialists from industry, academia and Government agencies to

present, discuss and exchange information on a wide variety of topics, including attitude determination,

orbit determination, attitude control system analysis and design, sensor calibration, mission analysis,

mission design, orbit control, and error analysis. A total of 40 papers were presented at the 1999

symposium. The papers presented at this symposium are available in a formal NASA publication

(NASA/CP— 1999–209235).

[Technical contact: John P. Lynch/572]
28

7.0 ISO9000

The FDAB complied with all requirements to ensure readiness for GSFC ISO9000 certification. For details

about GSFC ISO activities, please refer to http://arioch.gsfc.nasa.gov/iso9000/index.htm

[Technical contact: Felipe Flores-Amaya/572]

8.0 Outreach Activities

8.1 University Satellite Operations— SAMPEX

The GNCC awarded a new 5-year contract to the University of Maryland for SAMPEX Flight Dynamics

Operations support. SAMPEX, the first in the series of small explorer class satellites, was launched in

1992. Through a partnership with the GNCC, the University of Maryland developed the capability to

perform some spacecraft operations. With this new contract, complete responsibility has been given to the

university for performing operational orbit determination, orbit prediction, event scheduling (e.g., groundstation access times, shadow timing, ground track), attitude determination, and sensor monitoring for the

SAMPEX mission. GNCC engineers still provide consultation support and program review. Through this

relationship with the university, the GNCC also participated in classroom lectures as guest speakers and

offered tours of GSFC facilities to university students.

[Technical contact: Thomas Stengle/572]

8.2 Graduate Student Researchers Program— GSRP

The Graduate Student Researchers Program (GSRP) offers graduate students summer employment

positions through the University Programs Office. This is a competitive placement and is highly regarded.

The student worked with GSFC’s GNCC. As part of a 3-year plan, the student is matched with a designated

staff member as mentor for about 10 weeks (June 1–August 8) at GSFC in Greenbelt, Maryland. In addition

to offering a meaningful work experience, the program offers innovative ideas from outstanding students

for mission and activities at GSFC. This year we had the pleasure of working with Mr. Jose Guzman, a

Ph.D. candidate from Purdue University. Mr. Guzman worked on Dynamical System Theory Concepts in

regard to manifolds for libration orbit transfers. He held a class and presented his work in manifolds to the

mentor and other GNCC professionals. This work will be used for libration orbit design.

[Technical contact: David Folta/571]

8.3 Visiting Student Enrichment Program— VSEP

The Visiting Student Enrichment Program (VSEP) offers students summer employment positions with the

Universities Space Research Association (USRA), working with GSFC’s Earth & Space Data Computing

Division and numerous Center organizations that may include the GNCC. A selected student is matched

with a designated staff member as mentor for about 10 weeks (June 1–August 8) at GSFC in Greenbelt,

Maryland. In addition to offering a meaningful work experience, the program offers field trips and lectures

to provide broader appreciation for the mission and activities at GSFC. This year we had the pleasure of

hosting Mr. Steven Hughes, a master’s candidate from Virginia Tech University. Mr. Hughes worked on

Formation Flying Concepts in regard to attitude control and spacecraft pointing. He gave a final

presentation to the VSEP mentors and students.

[Technical contact: David Folta/571]
29

8.4 Public Education/Community Outreach

Supported the National Technical Association (NTA) annual conference, science fair & engineering

judging events, rocket building/engineering design projects, speeches (open, award, closing & graduation

ceremonies), program visits for GSFC mentor/student/fellow events, education/career conferences and

panels, online interactive session, and television/radio/magazine interviews.

[Technical contact: Aprille Ericsson- Jackson/572]
A1

Appendix A— Conferences and Papers

Bristow, J.O., G.T Dell, K.B Chapman, and D.L. Rosenberg, “Autonomous Formation Flying From the

Ground to Flight,” 1999 Flight Mechanics Symposium, GSFC, Greenbelt, MD.

Bristow, J.O., D. Weidow, “NASA/DoD University Nanosatellites for Distributed Spacecraft Control,” 13th

Annual Conference on Small Satellites, Logan, UT.

Carpenter, J.R., “Feasibility of Decentralized LQG Control of Autonomous Satellite Formations,” 1999

Flight Mechanics Symposium, GSFC, Greenbelt, MD.

Carpenter, J.R., D.C. Folta, and D.A. Quinn, “Integration of Decentralized LQG Control into GSFC’s

Universal 3-D Formation-Flying Algorithm,” 1999 AIAA GN&C Conference, Portland, OR.

Folta, D.C., and D.A. Quinn, “A Universal 3-D Method for Controlling the Relative Motion of Multiple

Spacecraft in Any Orbit,” AIAA/AAS Astrodynamics Specialist Conference, Boston, MA.

“Janus Trajectory Design,” Greg Marr, presented at the Flight Mechanics Symposium at GSFC, sponsored

by Code 572, May 1999.

Lauri Newman attended the AAS/AIAA Astrodynamics Specialist Conference in Girdwood, AK, on

August 15–19, 1999.

McComas, David C., James R. O’Donnell, Jr., Ph.D., and Stephen F. Andrews, “Using Automatic Code

Generation in the Attitude Control Flight Software Engineering Process,” 1998 Software Engineering

Laboratory Conference, GSFC, Greenbelt, MD, December 1998.

O’Donnell, James R., Jr., Ph.D., Stephen F. Andrews, and David C. McComas, “Development and Testing

of Automatically Generated ACS Flight Software for the MAP Spacecraft,” 14th International Symposium

for Space Flight Dynamics, Iguassu Falls, Brazil, February 1999.

O’Donnell, James R., Jr., Ph.D., Stephen F. Andrews, Aprille J. Ericsson-Jackson, Ph.D., Thomas W.

Flatley, Ph.D., David K. Ward, and P. Michael Bay, “Backup Attitude Control Algorithms for the MAP

Spacecraft,” 1999 Flight Mechanics Symposium, GSFC, May 1999.

“On-Orbit Performance of the TRMM Mission Mode,” Brent Robertson, Sam Placanica, Wendy

Morgenstern, Joseph Hashmall (CSC), Jonathan Glickman (CSC) and Gregory Natanson (CSC), 14th

International Symposium on Space Flight Dynamics, February 1999.

“The Importance of Semi-major Axis Knowledge in the Determination of Near-Circular Orbits,” Carpenter,

J. Russell, and Emil Schiesser, Space Flight Mechanics 1999, Univelt Publishing, San Diego.

“TRMM On-Orbit Attitude Control System Performance,” Brent Robertson, Sam Placanica and Wendy

Morgenstern, 22nd Annual AAS Guidance and Control Conference, February 1999.

Ward, David K., Stephen F. Andrews, David C. McComas, and James R. O’Donnell, Jr., Ph.D., “Use of the

MATRIXx Integrated Toolkit on the Microwave Anisotropy Probe Attitude Control System,” 1999 AAS

Guidance, Navigation and Control Conference, Breckenridge, Colorado, February 1999.
B1

Appendix B— Acronyms and Abbreviations

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

AAS American Astronautical Society

ACE Attitude Control Electronics

ACS Attitude Control System

AFRL Air Force Research Laboratory

AI Artificial Intelligence

ARC Ames Research Center

C&DH Command & Data Handling

CAST Chinese Academy of Space Technology

CCD Charged Coupled Device

CCSDS Consultative Committee for Space Data Systems

CDR Critical Design Review

CMP Central Mounting Plate

CNES Centre National D'etudes Spatiales

COTS Commercial Off-the-Shelf

CPU Central Processing Unit

CPT Comprehensive Performance Test

CRC Communications Research Centre

CREAM Cosmic Ray Energetics and Mass

DLR Deutsche Zentrum fur Luft- und Raumfahrt

DLS Double Lunar Swingby

DSN Deep Space Network

DSS Digital Sun Sensor

DST Dynamical Systems Theory

ECCO Extremely Heavy Galactic Cosmic-Ray Composition Observer

EIKF Extended Interlaced Kalman Filter

EKF Extended Kalman Filter

ELV Expendable Launch Vehicle

EO-1 Earth Orbiter-1

EOS Earth Observing System

ERBS Earth Radiation Budget Satellite

ESA European Space Agency

ESTO Earth Science Technology Office

FAST Fast Auroral Snapshot Explorer

FD Flight Dynamics

FDAB Flight Dynamics Analysis Branch

FDCL Flight Dynamics & Control Laboratory

FDS Flight Dynamics System

FDSS Flight Dynamics Support System

FM Flight Mechanics

FOT Flight Operations Team

FOV Field of View

FY Fiscal Year

GADFLY GPS Attitude Determination Flyer

GEAS GPS Error Analysis System

GEM Goddard Electronics Module

GEODE GPS Enhanced Orbit Determination Experiment

GLONASS Global Orbiting Navigation Satellite System

GN Ground Network

GNCC Guidance, Navigation and Control Center

GOES Geostationary Operational Environmental Satellite

GOTS Government-off-the-shelf

GPS Global Positioning Satellite
B2

GRO Gamma Ray Observatory

GSFC Goddard Space Flight Center

GSRP Graduate Student Researchers Program

HGA/CPU High-Gain Antenna/Contact Prediction Utility

HTML HyperText Markup Language

I&T Integration and Test

IKF Interlaced Kalman Filter

IMAGE Imager for Magnetopause-to-Aurora Global Exploration

IMDC Integrated Mission Design Center

ISS International Space Station

JHU/APL Johns Hopkins University/Applied Physics Laboratory

LP Lunar Prospector

MAP Microwave Anistropy Probe

MARGIE Minute of Arc Resolution Gamma ray Imaging Experiment

MAXIM Micro-Arc-second X-ray Imaging Mission

MIT Massachusetts Institute of Technology

MMS Magnetic Multi-scale Mission

MOC Mission Operations Center

MSRD Mission Support Requirements Document

MTASS Multimission Three-Axis-Stabilized Attitude Support System

NASA National Aeronautics and Space Administration

NASDA National Space Development Agency of Japan

NGST Next Generation Space Telescope

NGXO Next Generation X-Ray Observatory

NMP New Millennium Program

NOAA National Oceanic and Atmospheric Administration

OBC Onboard Computer

PC Personal Computer

PI Principal Investigator

POES Polar Orbiting Environmental Satellites

QUEST Quaternion Estimation

RMS Root-Mean-Square

RTADS Real-Time Attitude Determination System

RWA Reaction Wheel Assembly

RXTE Rossi X-Ray Timing Explorer

SA Selective Availability

SABER Sounding of the Atmosphere Using Broadband Emission Radiometry

SAGE Stratospheric Aerosol and Gas Experiment

SAIL Satellite Artificial Intelligence Laboratory

SAMPEX Solar, Anomalous, and Magnetospheric Particle Explorer

SAO Solar Astronomical Observatory

SBIR Small Business Innovation Research

SMEX Small Explorer

SOHO Solar and Heliospheric Observatory

SOMO Space Operations Management Office

SPC Spectroscopic and Polarimetric Coronagraph

SPECS Submillimeter Probe and the Evolution of Cosmic Structure

SPS Standard Positioning Service

STK Satellite Tool Kit

SWAS Submillimeter Wave Astronomy Satellite

TARA Two Axis Rate Assembly

TDRS Tracking Data Relay Satellite

TDRSS Tracking Data Relay Satellite System

TRACE Transition Region and Coronal Explorer

TRMM Tropical Rainfall Measuring Mission

ULDB Ultra Long Duration Balloon
B3

URL Uniform Resource Locator

VO Visualization Option

VSEP Visiting Student Enrichment Program

WIRE Wide-Field Infrared Explorer

WWW World Wide Web
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13. ABSTRACT (Maximum 200 words)

4. TITLE AND SUBTITLE

1. AGENCY USE ONLY (Leave blank)

6. AUTHOR(S)

9. SPONSORING / MONITORING AGENCY NAME(S) AND ADDRESS (ES)

11. SUPPLEMENTARY NOTES

Applied Engineering and Technology Directorate

Flight Dynamics Analysis Branch

January 2000 Technical Memorandum

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) 1999. 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 analysis results and mission experience

derived from the various mission support roles. The primary focus of the FDAB is to provide expertise

in the discipline of flight dynamics, which involves spacecraft trajectory (orbit) and attitude analysis,

as well as orbit and attitude determination and control. The FDAB currently provides support for

missions involving NASA, government, university, and commercial space missions, at various stages

in the mission life cycle.

Flight Dynamics Analysis Branch End of Fiscal Year 1999 Report 572

T. Stengle and F. Flores-Amaya

2000–00275–0

TM—2000–209485

Flight dynamics; EOS; FUSE; GOES; IMAGE; InFocus; Landsat; MAP; 29

Spartan; Triana; WIRE; ERBS; Lunar Prospector; TRMM.