# Presentation: 220749

## Section 18

_***Section 18***_
***Attitude Control System***

*** ******. . . Paul Sanneman***
*EO-1 ACS Lead Systems Engineer, Swales Aerospace, Inc.*

## Contents

***Contents***

*** ******Attitude Control Subsystem Block Diagram ***
***Summary of Requirements***
***Test and Verification Program***
***Critical Problem Reports***
***Special Topics***
***Independent Safehold (RFA #4.01)***
***Star Tracker Issues (RFA #14.16)***
***Nitrogen Purge of Gyroscopes (PR #468-20-4)***
***Independent Phasing Review***
***RWA Performance - Still in Family (RFA #15.06)***
***Can ACS de-tumble S/C from 5 deg/sec (RFA #9.22)***

## ACS Architecture Block Diagram

***ACS Architecture Block Diagram***

## ACS Level III Requirements (1 of 3)

***ACS Level III Requirements (1 of 3)***

***The ACS shall be capable of despinning the spacecraft from [5.0,1.5,0.5] deg/sec tip-off rates to within reaction wheel capability in 120 minutes via magnetic torquing***
***The ALI boresight shall be pointed along the geocentric nadir vector of the ECI J2000 coordinate frame***
***The ACS shall employ magnetic torquer coils to control reaction wheel momentum***
***The ACS shall control attitude errors to within 30 arcsec, 3-sigma, per axis, during the nadir-pointed science observations (as allocated by the System Pointing Budget) ***
***The ACS shall perform on-board attitude determination to within [54, 108, 54] arcsec, 3-sigma w.r.t. the ACS [X,Y,Z] axes, during the nadir-pointed science observations (as allocated by the System Pointing Budget) ***

## ACS Level III Requirements (2 of 3)

***ACS Level III Requirements (2 of 3)***

***The ACS shall provide a ground commandable slew maneuver capability which shall follow an attitude/rate trajectory based on uploaded parameters***
***The Safe Pointing Mode desired attitude shall orient the spacecraft such that the Sun vector is normal to the solar array cell-side when the SA is at the 0 deg position***
***Safe Pointing Mode pointing accuracy with respect to the Sun shall be ********25 degrees, 3-sigma, two axes; third axis about the sun line shall be rate damped only***
***The ACS shall utilize reaction wheels and magnetic torquer coils with sufficient momentum management for Safe Pointing mode (to avoid thruster operations)***
***The ACS shall provide attitude control to within 6 degrees, 3 sigma, per axis, via thrusters during delta-v maneuvers***

## ACS Level III Requirements (3 of 3)

***ACS Level III Requirements (3 of 3)***

***The ACS shall provide solar array closed-loop Sun tracking during orbit day with an accuracy of 5 degrees, 3-sigma and an open-loop rewind profile during eclipse***
***The ACS shall autonomously drive the solar array to the 0 deg position during the Sun Acquisition and safe hold pointing control modes***
***The ACS shall be capable of biasing reaction wheel speeds at a pre-determined orbital position to reduce S/C vibrations during a science observation***
***The ACS shall calculate two steering angles and issue pointing commands to the X-Band Phased Array electronics via the 1773 bus at 1 Hz***
***The ACS shall acquire and downlink telemetry from the IRU and AST at rates sufficient to support on-orbit calibration of ALI pointing ***

## EO-1 S/C Pointing Budget after Hyperion Addition

## S/C Pointing Budget Notes

***S/C Pointing Budget Notes***

***1)******	******ACS estimates  represent expected EO-1/Hyperion nadir pointing capability (including demonstrated star tracker performance)******	***
***2)******	******Conversions between angles and meters based on 705 km altitude***
***3)******	******GPS Uncertainties based on 'Phase 1' results of EO-1 Flight GPS Tensor Test Report (6/22/98)***
***4)******	******HSA-to-ACS on-orbit alignment uncertainty estimate based on post-processing of ACS attitude (30 arcsec), GPS navigation data (50 m), and HYP image recognition/processing (45 arcsec or 150 m)***
***5)******	******This budget does not include potential prediction errors (1 to 3 day predict) due to development of the absolute time command sequence (i.e., the ground targeting commands to ACS and the science observation commands)***
***Along-track error expected to increase by 2700 m after 3 days (3-sigma, worst case);  accepted by instrumenters and science community***

## ACS Component Capabilities Summary

***ACS Component Capabilities Summary***

## EO-1 ACS Staff

*** ******	****** ****Name**	**	**Company**	**	**Position**	**	**   Role**	**            Years Exp*
***	******  ******Paul Sanneman ******	******Swales Aerospace******	******ACS Lead******	******	******  Core team******	******	******14***
***Teresa Hunt******	******Swales Aerospace******	******ACS Analyst/Engineer******	******  Core team******	******	****** 3***
***Marcelo Gonzalez******	******Swales Aerospace******	******ACS Analyst/Engineer******	******  Core team******	******	******14***
***Dave Speer******	******Litton Adv. Systems******	******ACE Engineer******	******  Core team******	******	******18***
***Kathie Blackman******	******the Hammers Co.******	******ACS Analyst/Engineer******	******  Core team******	******	****** 8***
***Craig Woodruff******	******Swales Aerospace******	******ACS Flt Ops Engineer******	******  Core team******	******	******11***
***Paul Parry******	******MRJ Tech. Sol’ns******	****** ACS Engineer******	****** Independent******	******16***
***Seth Shulman******	******CSC******	******	******ACS Flt Ops Engineer******	******   Support******	******	******12***
***Kurt Woodham******	****** Litton Amecom******	****** ACS Analyst******	******   Support******	******	******10***
***Scott Miller******	******Litton Amecom ******	****** ACS Analyst******	******   Support******	******	******14***
***M. Shamma******	******Swales Aerospace******	****** ACS Analyst******	******   Support******	******	******12***
***Eric Holmes******	******NASA/GSFC******	******	******ACS Engineer******	******   Oversight******	******	******12+***

- ACS/ACE Software Engineers experience also extensive - see Flight S/W section

## ACS Technical Reviews

***ACS Technical Reviews***

***ACS CDR in June 1997***
***S/C CDR in June 1997***
***Safehold Mode CDR on April 17, 1998***
***ACS Hyperion Delta-CDR on Dec 12, 1998***
***ACE Safehold Mode Code Walkthrough on March 26, 1999***
***ACS Phasing Review on Dec 1, 1999***
***ACS FSW Acceptance Test Reviews (numerous 1999)***
***Independent Verification***
***Mr. Paul Parry of MRJ Technology Solutions served as a detailed independent ACS engineering consultant from Sept 1998 through Dec 1999***
***Mr. Eric Holmes of NASA/GSFC GNC Center has conducted an in-depth review of component data packages, S/C I&T procedures for ACS and continues in an oversight role***

## Attitude Control Subsystem Verification

***Attitude Control Subsystem Verification***

***ACS Component functional tests performed as part of integration***
***Verified Attitude Control Electronics interfaces as part of component integrations***
***MV ACS FSW interfaces to ACE verified using both override command capability and as part of phasing tests***
***MV ACS FSW control mode transitions, algorithm performance, and software performance within MV tested extensively in FSW Lab and with a selected subset on the flight MV using VirtualSat simulator***
***Includes GPS data processing, X-Band antenna steering, Solar Array Drive control, and Pulsed Plasma Thruster commands***
***ACE RSN Safehold Mode control software verified using extensive command/telemetry test, SHM phasing tests, and entry/exit test***
***Detailed MTB-to-TAM magnetic coupling data collected multiple times***

## ACS Phasing Tests Summary

## ACS Components Verification

**18 - **

## Accumulated Power On Time for EO-1 S/C & Components

## ACE RSN Safehold Mode Testing

***ACE RSN Safehold Mode Testing***

***The Attitude Control Electronics (ACE) contains the Safehold Mode control software for Sun Pointing operation during spacecraft anomalies***
***Spacecraft-level testing has validated all aspects of SHM***
***Ground commanded entry/exit***
***MV ACS Fault Detection & Correction commanded entry***
***ACE autonomous entry due to error in “I’m OK” handshake***
***ACE SHM control of Solar Array Drive ( to 0 Deg then Stop)***
***ACE SHM control paths (CSS, IRU to RWA, TAM/MTB)***
***ACS Component power on/off as appropriate for SHM ***
***Cold/Warm Restart functions***
***SHM control parameters commands and telemetry***
***Safehold Mode exercised for 24 hours during T/V testing***

## Reaction Wheel Assemblies

***Reaction Wheel Assemblies***

***RWA performance has been verified in both air and vacuum***
***Control torques issued by override and closed-loop software***
***Overall run-up and coast down trends have been verified with vendor and are summarized on next page***
***Thermal/Vac testing of overspeed protection provided below  ***

## RWA Run-up / Coast Down Trend Table

***Trending data is consistent over time.  ***
***RWA-Y run-down time variation still in family (Special Topic)***

## Autonomous Star Tracker

***Autonomous Star Tracker***

***A number of AST-related issues have arisen due to continued testing of the MAP ASTs and PR generated during EO-1 testing***
***See Special Topic on AST Issues for details of most items***
***EO-1 AST has experienced toggling in the number of stars***
***Star Field Stimulator has only 17 stars - some toggling to 16, 18 seen with some orientations of the SFS***
***16 star events are caused by GSE  configuration/cleanliness***
***18 star events related to SNR for very dim stars - not expected to be an issue on-orbit, response procedures ready if needed (Open PR #674-20-41)***
***AST thermal design has been modified based on analyses and results of S/C Thermal/Vac testing ***
***Silver teflon tape replaced with MLI blanketing to protect 1773 transceiver during powered off COLD case (i.e. SHM)***
***Thermal Electric Cooler operation validated during Thermal/Vac***

## GPS Constellation Simulator Test

***GPS Constellation Simulator Test***

***GPS Constellation Simulator provides end-to-end test capability of spacecraft hardware/software related to GPS***
***Based on a representative EO-1 orbital scenario, RF signals are generated and used to stimulate the flight GPS antennas ***
***GPS Tensor outputs time and navigation vectors consistent with the scenario, which are utilized in the MV C&DH and ACS***
***Used to verify phasing, magnitude, and units, as well as the generation of the ACS attitude commanding***
***GPS Tensor Warm start command capability has also been proven***
***Ground command procedures for initializing the GPS Tensor under development; recent testing has verified MOC capability to properly deliver necessary parameters to S/C and S/C to GPS***
***GPS Tensor component level performance validated following all S/C environmental testing***

## Thermal/Vacuum Test Results

***Thermal/Vacuum Test Results***

***Attitude Control Subsystem performance validated as part of CPT segments and via other special tests***
***All CPT associated tests run at hot and cold plateaus***
***RWA Overspeed protection verified at hot and cold plateaus***
***AST thermal design characterized; analysis model updated***
***AST Thermal Electric Cooler operation verified***
***IRU internal heater control of gyro platform validated***
***Extended duration operation in ACE Safehold Mode***
***ACE operated with maximum MTB current for hot case ***

## Changes Since S/C Thermal/Vacuum

***Changes Since S/C Thermal/Vacuum***

***ACS Components***
***AST removed and X-rayed; subsequent removal/replacement of EMI Filter with additional clearance between part and PWB***
*Debris found under EMI Filter and removed during repair*
***Removed IRU to prevent helium exposure during MAP testing ***
***ACE Software***
***ACE RSN S/W update to A/D conversion implementation***
***MV ACS Software***
***Ephemeris epoch time changes from MET to UT to improve navigation propagation accuracy (MV clock drift)***
***Modified Warm restart logic to ensure consistent recovery time***

## Critical PR Status (1 of 3)

***Critical PR Status (1 of 3)***

***PR #858-30-2 IRU Data Checksum Error during Stim Command***
***Due to asynchronous nature of the ACE-IRU interface,  an error can occur in one data packet during a command transmission***
***One packet corrupted out of 25 packets/sec does not affect measured body rate used by control law***
***IRU Stim commands are not used on-orbit***
***PR #782-20-1 MV Warm Start during build 4.5 Testing***
***During an alternate method for performing a leap second adjustment, the ACS code delta-time was near zero; this caused an FPU exception and MV Warm start***
***Constraint manual documents the anomaly and provides the preferred method using the ATS to perform a leap second change***

## Critical PR Status (2 of 3)

***Critical PR Status (2 of 3)***

***PR-790-30-5  ACE failed to enter Safehold due to SHM inhibit***
***An original ACE S/W requirement for ground GSE-based inhibit for SHM entry was included in code although not needed ***
***After reported anomaly, the feature was completely removed from the ACE S/W and verified (PTR 868)***
***PR-794-20-6 MV Warm Start due to arccos function error***
***Lunar vector not properly normalized prior to usage, which resulted in attempting to perform an arccos of value slightly higher than 1.0***
***MV ACS code updated to perform normalization (PTR869)***
***PR-790-30-3 ACE Entered Safehold following MV Warm Restart***
***Safehold timeout set to 10 seconds during this testing***
***MV Warm restart took longer than 10 seconds, causing SHM***
***MV ACS task synchronization during Warm start modified in Ver 4.8***
***Subsequent testing on 3/13/00 demonstrated ************  6 sec Warm restarts***

## Critical PR Status (3 of 3)

***Critical PR Status (3 of 3)***

***PR-468-20-4 IRU Parametric Drive Voltage***
***See Special Topic ***
***PR-680-20-1 ACEPROPONOFR value went Red High***
***Thruster current sum is dependent on Bus Voltage***
***Limit applied  during test was incorrect***
***Constraint maual documents Bus Voltage dependency***
***This telemetry item is NOT used on-board***
***PR-943-40-1  Filter Table Jam did not work properly***
***Telemetry Output code had limited range for APID***
***MV FSW patch on 23 MAR 2000 to increase the allowed range was implemented and verified***

## RFA #15.07  S/C Magnetic Dipole

***RFA #15.07  S/C Magnetic Dipole***

***RFA  EIRR#7 from PSR (15.07)******Please provide the data which includes the estimated magnetic dipole for the total spacecraft including operating solar arrays.  Define a corresponding assessment of ACS impacts, if any.***
***S/C static magnetic test indicated over 5,000 Gamma @ 1m***
***Residual dipole moment estimated near 9 Am******2****** (worst case)***
***Analysis by TECSTAR presented at Solar Array CDR 12/8/1998***
***Worst-case analysis resulted in 0.004 Gauss @ 20 inches (uncompensated)***
***Mr. J. Scheifele of GSFC/695 estimated this to be 10 Gamma @ 1m after acounting for compensating effect of SA wiring***
***Operating solar array less than 1% of S/C main body dipole***
***ACS analyses/simulations have verified the expected MTB dipole margin and no impact to pointing performance***
***Operations tools exist for estimating S/C residual magnetic field and trending it with respect to time, if necessary***

## ACS Flight Readiness

***ACS Flight Readiness***

***ACS components and hardware are ready***
***ACS and ACE software are ready***
***ACE Safehold Mode is ready***
***Ops preparations/training are in work***
***Remaining Work***
***Re-integrate IRU (mechanical, purge, electrical, alignment)***
***Magnetic phasing check to be scheduled near last CPT***
***Final CSS check after Solar Array attached/deployed ***
***AST alignment following final mounting***

***Upon completion of above remaining work,***
***Attitude Control Subsystem is READY***

## Special Topic:

***Independent Safehold Mode (RFA #4.01)***

*** ******. . . Paul Sanneman***
*EO-1 ACS Lead Systems Engineer, Swales Aerospace, Inc.*

## Overview

***Overview***

***Safehold Mode Implementation***
***Safehold Mode Reviews***
***ACE SHM Testing***

## ACS Architecture Block Diagram

***ACS Architecture Block Diagram***

## ACS Components Science Pointing

***ACS Components Science Pointing***

## ACS Components - Safehold Mode

## ACE RSN Safehold Mode Testing

***ACE RSN Safehold Mode Testing***

***The Attitude Control Electronics (ACE) contains the Safehold Mode control software for Sun Pointing operation during spacecraft anomalies***
***Spacecraft-level testing has validated all aspects of SHM***
***Ground commanded entry/exit***
***MV ACS Fault Detection & Correction commanded entry***
***ACE autonomous entry due to error in “I’m OK” handshake***
***ACE SHM control of Solar Array Drive ( to 0 Deg then Stop)***
***ACE SHM control paths (CSS, IRU to RWA, TAM/MTB)***
***ACS Component power on/off as appropriate for SHM ***
***Cold/Warm Restart functions***
***SHM control parameters commands and telemetry***
***Safehold Mode exercised for 24 hours during T/V testing***

## ACE Safehold Mode Reviews/Status

***ACE Safehold Mode Reviews/Status***

***Safehold Mode CDR on April 17, 1998***
***Attendees included GSFC E-O1 S/C Systems Engineer Brian Smith and GNC Engineer Alan Reth***
***Alan Reth’s feedback “the ACS team has done an excellent job with EO-1 Safehold Mode”***
***Final ACE Safehold Mode code walkthrough March 26, 1999***
***ACE SHM software control algorithms also separately tested using the Hammers VirtualSAT simulation***
***Safehold Mode has been fully integrated with the rest of the S/C and is tested as part of the CPT***
***MOC Safehold Mode recovery training simulation performed using a configured SHM recovery procedure***

## Special Topic:

***Star Tracker Issues (RFA #15.07)***

* ****. . . Wayne Roher***
*Litton Advanced Systems*

## Autonomous Star Tracker (AST) Issues

***Autonomous Star Tracker (AST) Issues***

***MAP AST Issues since EO-1 AST delivery from Lockheed-Martin***
***Floating ACTEL MODE pin***
***Solder Ball under 5V converter***
***15V DC/DC converter***
***EMI Filter Short***
***C side CCD noise***
***AD811 design deficiency***
***Unverified EMI CS02 test failure***

## AST Mode Pin

***AST Mode Pin***

***Problem***
***The MODE input to one of the ACTEL FPGA devices was an inadvertently left open circuit due to a PWB layout design error on the AST I/O board***
***The MODE pin was found to be “floating” when Lockheed-Martin was attempting to trouble shoot the MAP AST I/O for other issues***
***Potential Symptoms***
***If the MODE pin had floated “high”, communications with the tracker would have been lost.***
***EO-1 AST test history at component and S/C level has no indication of this condition***
***EO-1 Recovery***

## EO-1 Recovery Plan

***EO-1 Recovery Plan***

***At our request Lockheed-Martin opened the tracker, installed the jumper.  We performed  workmanship vibe test and functional test on S/C without problems.***
***Potential overstress anomaly investigation***
***Using two commercial FPGAs, programmed by Lockheed-Martin, simulated the condition to potentially overstress one of the FPGAs.***
***Part sent to Lockheed-Martin and tested fine in their engineering model AST.***
***Lockheed-Martin further tested the part by forcing the mode pin “high”; result was tracker ceased to function as expected.  Upon removing the mode pin forcing function, the AST resumed normal operation***
***Conclusion***
***EO-1 position is to fly AS IS because:***
*1)**	**Lockheed-Martin Mission Success panel rescinded recommendation to replace FPGA.  Grounding mode pin is all that is needed*
*2)**	**Elaborate testing failed to reveal evidence of overstressed part.*
*3)**	**Residual risk is deemed negligible.*

## AST Solder Ball Concern

***AST Solder Ball Concern***

***Solder ball shorted a pin on the DC/DC converter on MAP1***
***Verified by X-Ray***
***MAP2, Orbital, and SIRTF power supply boards  X-rayed, no solder balls found***
***Considered random process failure***
***EO-1 was x-rayed, no solder balls found***

## AST 15V DC/DC Converters

***AST 15V DC/DC Converters***

***MAP experienced  failure with ± 15V DC/DC converter***
***MAP2 - Open failure during ATP at ambient after Thermal/Vacuum***
*LDC9905/9845*
*DPA indicate inadequate stress relief on internal bond wire, insufficient solder, and solid adhesive attachment to lid caused bond wire to pull away from pad in vacuum.*
*LM and GSFC have performed several DPAs of different LDC and found units to be acceptable.*
***Conclusion***
***EO-1 position is to fly AS IS***
*Based on X-ray performed on parts in early January.*
*No indication of solder balls under part*
*Sufficient clearance of internal wire to inductor (sufficient strain relief)*
*EO-1 has survived 450 hrs. of Thermal/Vacuum*
*Still an open issue with Flight Assurance Office.  Will discuss at Red Team Follow-up Review.  Potential residual risk candidate.*

## AST EMI Filter Short(RFA #14.16)

***AST EMI Filter Short****(RFA #14.16)*

***MAP2 experienced a short on +28V***
***U8 was flush mounted to the board creating a short between filter housing and +28V pad***
***LMMS has no recorded measurements of part clearance***
***RFA #14.16 requested a method be developed to measure the clearance***
***Recovery***
***Only known way to be sure of proper clearance is to measure.  EO-1 X-ray indicated no clearance.  Also indicated presence of solder balls with high potential for shorting.***
***Filter was replaced with proper clearance by Lockheed Martin Technician ***
***AST was vibrated a workmanship levels, returned to S/C, and tested with out issue***
***Negligible residual risk***

## AST C Side CCD Noise

***AST C Side CCD Noise***

***MAP2 found CCD “C” side noise increase up to 20 A/D counts during thermal cycling from -8° to 0°C***
***Root cause is physical routing of C side clock signal***
***Photograph of EO-1 board showed clock signal routing different from MAP2 board, less likely to cause noise in the output***
***EO-1 software does not permit direct noise measurement***
***Longer integration time results in a higher SNR.  99.4% probability of retaining lock under worst conditions.***
***Negligible residual risk***

## AST AD811 Design Issue

***AST AD811 Design Issue***

***Resistor bias of A/D converter could cause clipping of bright star image intensity***
***Bright stars above saturation limit would not be used.  Very small number of stars affected; will not effect performance of EO-1 AST***
***Negligible residual risk***

## AST Unverified EMI CS02 Test Failure

***AST Unverified EMI CS02 Test Failure***

***MAP2 experienced soft reset during EMI CS02 testing***
***MAP associated problem with EMI filter.  Solder ball found under filter. Map replaced EMI filter.  Subsequent testing has had no further resets.***
***MAP1 has not had failure***
***No failure during EO-1 testing (component level and spacecraft level)***
***No replacement is recommended by LMMS for other trackers as issue was associated with workmanship***
***Negligible residual risk***

## What Happens If The AST Fails?

***What Happens If The AST Fails?***

***AST has been extensively tested on S/C and performance has been nominal.  All issues have been aggressively pursued.***
***Demonstrated reliability of product line is less than desirable.***
***Error prone manufacturing and test efforts***
***Similar AST on DSI failed 15 months into mission***
***Flight dynamics ground based attitude determination could achieve 1 to 5 degree pointing accuracy***
***Would allow Landsat 7 limited image comparison since we would be within its 15 degree swath width***
***Ground attitude estimate developed using TAM and gyro data***
***Current flight ACS S/W has capability to use daily uploaded attitude estimate, propagated using gyro data***
***Would require some ground software to prepare and uplink daily attitude data; would be developed if AST fails.  Sun sensor safehold used until software developed.***
***Pending DC/DC issue, this AST has negligible residual risk.***

## Special Topic:

***Nitrogen Purge of the Gyroscopes***
***(Critical PR #468-20-4)***

*** ******. . . Paul Sanneman***
*EO-1 ACS Lead Systems Engineer, Swales Aerospace, Inc.*

## SIRU Nitrogen Purge - Background

***SIRU Nitrogen Purge - Background***

***Litton Guidance & Control Space Inertial Reference Unit***
***Utilizes three Hemispherical Resonator Gyros (HRGs)***
***Quartz crystal hemisphere oscillated at specific amplitude and frequency***
***Input angular rate measured by nodal displacement on rim***
***Quartz crystal becomes permeated with Helium***
***Increases gaseous damping, requires larger drive voltage - Parametric Drive Voltage (defined as a negative voltage)***
***Not an issue prior to arrival at GSFC Bldg 7***
***Significant amounts of Helium released in Bldg 7 by environmental chamber testing processes***

## PDV Trends Since Initial Integration

## Summary of PDV History

## Gyro A PDV Slope Estimate

## PDV Trends Since S/C Thermal/Vac

## PD Trends Normalized to 60 C

## SIRU Nitrogen Purge History - 1999

***SIRU Nitrogen Purge History - 1999***

***Test activity next to EO-1 clean tent in early August 1999 vented helium into room***
***Battery AC intake near that vicinity, outlet inside S/C near SIRU***
***Dedicated nitrogen purge installed and started August 10, 1999***
***Flight purge modifications in November 1999 to use existing manifold for purge up until T-0 (similar to instruments)***
***Provides helium protection through the launch environment, especially during a launch scrub when 2nd stage tanks are vented***
***PDV slope near zero during November 1999***
***PDV slope declined starting in the first week of December 1999***
***Coincident with Chamber 225 helium leak test (prep for MAP test)***
***30 day degradation period followed exposure***

## SIRU Nitrogen Purge History - 2000

***SIRU Nitrogen Purge History - 2000***

***Gas samples taken on Jan 15, 2000 were analyzed for helium concentration***
***Bldg 7 air near EO-1 SC was  330.5 PPM  (normal air is <5 PPM)***
***Routine gas samples of Bldg 7 air initiated; Bldg 7 operations coordination***
***Transitioned from bottled GN2 to boil-off GN2 (1/7/00)***
***Confirmed bottle purity was not the source of exposure***
***Removed battery AC and transitioned to chiller plate (1/7/00)***
***Nitrogen flow rate increased to further dilute any helium passing into the purge line and to improve flow through IRU box***
***from 0.5 L/min to 1 L/min (1/7/00);  from 1 L/min to 2 L/min (1/10/00)***
***from 2L/min to 5 L/min (1/24/00)***
***5 of 6 chevron vent holes blocked to improve flow (1/27/00)***
***Significant air gaps in SIRU box enclosure also apparent***
***Copper tube used to enclose Teflon purge line (2/11/00)***
***Each end of copper tube wrapped with aluminum tape to close-off ***

## Helium Measurements in Bldg 7

## SIRU Removal From Spacecraft

***SIRU Removal From Spacecraft***

***Removed from S/C on Feb. 25, 2000***
***Rationale***
***MAP instrument testing re-started in Chamber 225 - ten dewars of liquid helium to be used for constant instrument testing***
***Significant concerns and uncertainty regarding the Dec 1999 He exposure***
***WARP re-integration delay allowed for IRU removal without seriously impacting S/C test activities***
***Method for helium venting to outside of Bldg 7 not reliable***
***Allow access for possible sealing of air gaps in IRU box enclosure***
***Storage conditions at Swales Aerospace Bldg 5015***
***New stainless steel convoluted inner tube with exterior braid***
***Boil-off GN2 from dedicated tank at flow rate of 0.5 L/min***
***No helium in use throughout building***

## SIRU Return To Spacecraft

***SIRU Return To Spacecraft***

***Re-integration to be performed following completion of MAP instrument testing and Bldg 7 air sampling validation of low helium concentration***
***Expect this will be within the first week of April 2000***
***Re-integration will utilize a new aluminum-wrapped teflon tube inside spacecraft, new stainless steel tube from S/C to tank***
***Gyro Parametric Dive Voltages will be measured following electrical integration to the S/C to verify margin still exists***
***If  sufficient margin does not exist, IRU would be returned to vendor, ******Litton Guidance & Control, for replacement of all three gyros***
***PROM replacement on system electronics module would also be required which increases risk; full IRU re-calibration sequence also then needed***
***Baseline gyros’ performance was excellent - most likely that  the ******replacements will not achieve same performance***

## Outlook

***Outlook***

***Mechanical and launch vehicle I/F engineers have focused activity for IRU protection throughout remaining S/C I&T at GSFC, during transport and launch site processing***
***The increased PDV is NOT a performance issue - just needs to stay within (-5V) capability***
***About one volt of margin required for full 10 deg/sec rate capability***
***Desire PDV values of 3.5 V (@25C) at time of SC mount on LV***
***PDV Voltages on 2/25/00: A=-2.81, B=-2.35, C= -2.65  @23C***
***Purge installation on launch vehicle has been improved to the extent allowable by the LV vendor/operator (see LV charts)***
***Purge flow rate to IRU separately metered at a flow rate of 5 L/min ***
***Expect Parametric Drive Voltage to remain stable (i.e. at current level) through launch and will slowly recover over the course of the mission - residual risk has been minimized***

## Special Topic:

***Independent Phasing Review***

*** ******. . . Paul Sanneman***
*EO-1 ACS Lead Systems Engineer, Swales Aerospace, Inc.*

## Independent ACS Phasing Review

***Independent ACS Phasing Review***

***Attitude Control Subsystem phasing has been diligently developed, reviewed and tested for over two years***
***Initial ACS Phasing document released in Dec 1997***
***Formal Phasing Review with Code 570 on 1 Dec 1999***
***Final ACS Phasing document released in Dec 1999***
***Swales employed an independent consultant, Mr. Paul Parry of MRJ, who is familiar with the SSTI/Lewis ACS design and issues, to review ACS design, analyses, I&T, and phasing issues***
***Conclusion is ACS design and validation is adequate for this mission***
***ACS portions of the S/C CPT have been designed to validate the actuator command polarity and magnitude for all modes possible***
***Earth spin rate and magnetic field used as physical checks***
***Electrical, optical, and commandable stimuli also utilized***

## ACS Component Locations - View 1

***ACS Component Locations - View 1***

## ACS Component Locations - View 2

***ACS Component Locations - View 2***

## Independent ACS Phasing Review - Closure

***Independent ACS Phasing Review - Closure***

***GSFC Code 570 (GNC Center) participants were very satisfied with the EO-1 ACS phasing development, documentation, and verification***
***Open actions from Phasing Review***
***Execute all ACS CPT Phasing tests following final MV S/W load [planned as part of pre-ship CPT]***
***Execute CSS Phasing for both MV ACS SunAcq and ACE SHM following final solar array attachment (with flight CSS config) [planned as part of pre-ship activities]***
***Repeat compass/magnet checks of magnetic phasing [completed]***
***Review analyses of thruster control with respect to expected CG migration due to propellant usage [completed]***

## Special Topic:

***RWA Performance -- Still in Family***

*** ******. . . Paul Sanneman***
*EO-1 ACS Lead Systems Engineer, Swales Aerospace, Inc.*

## Background

***Background***

***RWA run-down or spin down time is used by the vendor and spacecraft engineers as an indicator of friction in wheel bearing/lubricant***
***RWA is not hermetically sealed so observed friction is dependent on air pressure (i.e. much higher ‘friction’ in air due to flywheel windage)***
***RWA specification includes drag torque (i.e. friction) versus wheel speed for both air and vacuum conditions***
***Drag torque is a function of speed due to viscous friction effects***
***Ithaco Acceptance Testing of RWAs in Thermal/Vacuum includes drag torque verification at both hot (+55C) and cold (-20C) extremes***
***RWA serial number to EO-1 spacecraft reference***
***SN 76674 is RWA-X;      SN 76675 is RWA-Y;       SN 76676 is RWA-Z***

## RWA Run-up / Coast Down Trend Table

## Summary of Processed Data Plots

***Summary of Processed Data Plots***

***RWA rundown data from performance test executions was post-processed to estimate the drag torque as a function of wheel speed***
***Same process and time interval as used by Ithaco during ATP***
***Drag torque measurements plotted together with performance specification ***
***All values of drag torque are well within spec and agree well***
***Expected temperature variation due to lubricant viscosity***
***Plotted results on subsequent pages***

## RWA ATP Drag Torque Summary

***Linear estimates from measured ATP data, both Hot (55C) and Cold (-20C) plotted for each RWA***
***Note that RWA-Y drag performance here is consistent with S/C TV data***

## Slide 69

## Slide 70

## Slide 71

## Slide 72

## Slide 73

## Slide 74

## Slide 75

## Slide 76

## Slide 77

## Statements From Ithaco Space Systems

***Statements From Ithaco Space Systems***

***Bill Bialke - Ithaco Director of Engineering***
***EO-1 RWA drag torques as measured during the S/C Thermal/Vac test are consistent with the ATP and well within the historical family for these wheels, including recent measurements on MARS-2001 Program***
***Increased drag at cold temperature is related to excess grease in between the retainer and its guiding land.  A web of lubricant can form in this area, with a chord length which can vary widely depending on lubricant distribution.***
***The Rheolube 2000 grease used in the design has exhibited excellent retention characteristics, which can increase run-in time on some units***
***Experience at Ithaco predicts that additional extended run-in will most likely bring units in family, even at cold temperatures***
***The higher than average drag torque at cold temperature does not effect reliability or lifetime***

## Recent Data from Mars 2001 Program

***Recent Data from Mars 2001 Program***

***Data from Mars-01 Type A RWAs also shows similar increased drag torque in one RWA at cold temperature only***
***At room temperature and above, drag torque is in family, which confirms that the difference is lubricant related***
***Lubricant is the only drag torque contributor with temperature sensitivity***
***Out-of-family unit was accepted, since it represents an over-lubricated situation, which does not affect life or reliability***

## Special Topic:

***Can ACS De-Tumble S/C From 5 Degrees / Second? (RFA #9.22)***

*** ******. . . Paul Sanneman***
*EO-1 ACS Lead Systems Engineer, Swales Aerospace, Inc.*

## Launch Vehicle Separation Summary

***Launch Vehicle Separation Summary***

***Most recent spacecraft separation angular rates [X,Y,Z]***
***X-axis rate highest due to S/C stowed CG w.r.t. Y-axis***
***[ 3.9 , 0.5 , 0.0] deg/sec***   ***mean case ***
***[ 5.1 , 1.5 , 0.5] deg/sec***   ***worst case (mean + 3-sigma)***
***Initial total momentum range is 12-17 Nms***
***Angular rates, not momentum, affected by solar array deployment***
***Time to stabilize not dependent on SA deployment***
***Conservation of angular momentum    H=Iw***
***Separation location w.r.t. eclipse is dependent on launch date***
***Affects final sun acquisition time since controller may begin using coarse sun sensors just prior to eclipse entry ***

## ACS Initial Acquisition Sequence

***ACS Initial Acquisition Sequence***

***Nominal MV control mode sequence***
***BDOT Magnetic despin***
***Autonomous transition to Sun Acquisition based on momentum threshold (uses IRU rate measurements)***
***Sun Acquisition time dependent on transition time with respect to eclipse***
***In case of anomaly and Safehold Mode entry, no BDOT is performed - SHM performs Sun Acquisition***
***Expect to saturate RWAs (not a problem, higher power)***
***Magnetic unloading will still dump the momentum***

## BDOT Momentum Dumping 

***Various polarities of imparted rates ******from LV separation of****** [************ 5.1 ************ 1.5 ************ 0.6] deg/sec***.
***Average momentum decrease is approximately 8 Nms per hour***

## Representative Sun to Solar Array Angle

***Solar Array Error for [- - -] Tip Off Rates***

***BDOT duration of 2.0 hours***
***BDOT-to-SunAcq Transition near 21:30***
***Calculation made whether in eclipse or not***

## SHM vs. BDOT Momentum Dumping

***Launch vehicle imparted momentum ******will be dumped by either ******ACS control mode sequence***

## Conclusion

***Conclusion***

***ACS has been designed to perform an autonomous sun acquisition with the worst case tip-offs rates from the LV***
***High-fidelity simulations and ACS flight software acceptance testing have demonstrated expected timeline and performance***
***Battery depth of discharge for worst case scenario is acceptable (see Power System charts)***
***Communications issues related to spinning spacecraft are being addressed***
***S/C level training simulation for tumbling case was performed on March 9 to flush out ACS/COMM/PWR issues***
***Final S/C mass properties and launch date will be used to update the ACS initial acquisition scenarios prior to launch***