PSR Section 3

1
Section 3 Verification Process
. . . Mark Perry Swales Aerospace
2
Verification Process Agenda

• Environmental Test Approach
• Verification Matrix
• Satellite Environmental Test Flow
• Satellite Environmental Test Documentation
• Summary of Satellite and Component Test Time
• Functional and Performance Testing
• Environmental Testing Overview
• LV Separation Shock / Vibration
• Acoustics
• EMI/EMC
• Magnetic Survey
• RF Compatability
• Thermal Balance/Thermal Vacuum
• Alignment
• Mass Properties
• X-Band Near Field Assessment
• End to End Instrument Optical Testing

3
Environmental Test Approach

• EO-1 Verification Process is described in the
  EO-1 Verification Plan and Environmental
  Specification (SAI-SPEC-158)
• Verification Matrix identifies the box and system
  level testing
• Some flight box level testing was deferred to
  observatory level testing based on qualification
  by similarity or a second in a series of like
  hardware
• System level environmental testing was with all
  flight hardware with exception of the flight
  battery or where configuration did not allow
  example the solar array was not included with the
  Observatory during Thermal Vacuum
• Satellite Environmental Flow comprised the
  following
• Shock (2) Firings
• Vibration - 3 axis sine sweep at protoflight
  levels (1.25 x flight limits)
• Acoustic at protoflight levels
• EMI/EMC
• Thermal Balance
• Thermal Vacuum

4
Final System Level Spacecraft Verification
5
Spacecraft Verification Flow
5/3
5/11
7/14
7/15
7/24
7/26
7/26
7/30
7/31
8/5
8/5
8/6
7/16
7/23
1773 Sensitivity Test
1st CPT (PreShip to GSFC) SA Deployment on S/C
OTDR F/O measurements, S/W Tests
Partial 2nd CPT MOC Tests
Acoustic Test
Weigh Satellite 3 Axes Vibration 2 Shock Tests -
Modal Tap Test AST/IRU Alignment
CPTMOC
Acoustic
CPT
F/O
Modal/Align
Vib/Shock
F/O S/W
8/7
8/8
8/29
9/4
9/5
9/16
9/17
10/1
10/1
10/20
10/30
10/21
8/17
8/19
EMI/ EMC Radiation Emissions Test
M5, RSN, PSE S/W Test Regression Test
TB/TV 4 Cycles - CPT HOT - CPT Cold - MOC
Simulation - Safe Hold 24 Hrs - 1773
Sensitivity - Functional Test
Alignment Measurement
RF Compatibility CPT0-CPT11
S/C Timing Test 1773 Margin Test Functional
Test ACE/XPA/PSE S/W
ACE/XB S/W Load Test
Alignment
RF Compatibility and CPT
EMC
F/O, Timing
S/W
S/W Test
TB TV
11/1
Pre-Ship
11/9
11/17
11/8
11/18
11/30
12/11
11/29
Verification Flow -CPT -F/O Sensitivity -Modal
Survey -Alignment Measurements -Weigh
S/C -Vibration -Shock -Acoustic -RF
Compatibility -EMC/RE -Thermal Balance -Thermal
Vac -Instr End to End - Magnetic Survey
(Static) -EMC/CE -Pre-Ship CPT -Mass Properties
ALI Hyperion ETE Test Hyperion VNIR
Characterization XB NF Test
1773 Bus Buster MOC SIMS
Magnetic Survey RCS Press/Leak EMI/EMC CE Test
Self Compatibility
Pre-Ship CPT
Mass Properties
Ship to WTR
Instrument ETE X-Band Characterize
S/W, MOC Test
EMC
Mass PROPS
CPT
6
Verification Documentation

• EO-1 Verification Plan Environmental
  Specification SAI-SPEC-158
• CPT SAI-PROC-679A
• F/O Sensitivity TD WOA-462
• 1773 Sensitivity TD WOA-496
• Alignment SAI-PLAN-206
• Modal Test SAI-PLAN-341
• Shock Test SAI-PLAN-341 SAI-PROC-739
• Vibration Test SAI-PLAN-307 SAI-PROC-739
• Acoustic Test SAI-PLAN-320
• RF Compatibility SAI-PLAN-XXX, WOA-410/590 (STDN
  - 408.1)
• Mass Properties SAI-PLAN-319 (Test Postponed)
• TB/TV SAI-PLAN-273 SAI-PROC-619
• Instrument End-to-End SAI-PLAN-358
• Magnetic Test SAI-PROC-768
• EMI/EMC Test SAI-PROC-752
• X-Band Near Field Test SAI-PLAN-368

7
Additional S/C Testing

• Testing not included in original Verification
  Plan
• ALI Hyperion End to End Testing (Optical
  Testing)
• Hyperion VNIR Characterization
• 1773 Bus Buster Test
• MV Stress Testing
• S/C Timing Test
• X-Band Near-Field Test
• Post-Environmental Flash Test of Solar Arrays

8
Projected Power On Time
9
EO-1 Functional and Performance Testing
10
CPT Design and Description

• Goal is to test every function and to measure
  performance for each subsystem and for the
  complete EO-1 system
• With the exception of some RF and RCS testing,
  all tests are conducted from ASIST STOL
  procedures to facilitate repeatability and reduce
  errors
• The CPT contains about 100 individual tests.
  These tests were planned to perform as much
  parallel testing as possible without sacrificing
  the integrity of the individual tests.
• All nominal and critical commands (about 500)
  tested during the CPT
• The rest of the on-orbit commands are tested
  outside the CPT, such as during
  software-acceptance tests
• Additional commands, which were used for
  integration, development, test and diagnosis,
  will be deleted from the database.
• The CPT is comprised of 12 sections, as defined
  in SAI-PLAN-659A

11
CPT Sections

• Sections 0 and 1 Power-on defaults, launch,
  ascent, deployment, and sun acquisition
• Sections 2, 5, and 6 Exhaustive s/c tests ACS,
  Power, CDH, RF, S/A drive, deployment
  redundancy
• Section 3 Non-imaging instrument and WARP
• Section 4 EFF
• Sections 7 and 8 Exhaustive imaging
• Sections 9 and 10 Restart, safemode, and
  load- shedding
• Section 11 Orbit-timeline

12
Non-CPT System Performance Tests

• These are special tests that are incompatible
  with the CPT
• Require special satellite configuration which
  would inhibit other tests
• Require special GSE
• Limitations such as maximum pressure cycling
• Specific Non-CPT Tests
• End-to-end optical tests verify instrument
  alignment and optics (require EO-1 in a
  horizontal attitude)
• 1773 tests bus buster/SEU and optical
  performance (special GSE)
• X-band ACS-to-X-band pointing tests and
  near-field test (special GSE)
• GPS constellation tests requires special GSE and
  is incompatible with other systems tests using
  ephemeris and time.
• ACS virtual satellite (simulates ACS components)
  software testing
• FDC testing TSMs, RTSs, WARP, PSE, and ACS
• Solar-array mechanical deployment (to exercise
  the s/a array drive and associated software, the
  s/a is not mounted during CPT)
• RCS pressure testing propulsion tank is limited
  to a few cycles
• Heater tests requires cold-soak
• MOC simulation tests system tests in additional
  operational scenarios
• LFSA and PPT (category III payloads) tested
  separately
• Flash test of solar arrays (requires special
  equipment and location)

13
Baseline CPT

• Conducted August 31st to September 4th
• Generated 160 problem reports, dominated by
  procedure errors
• 120 PRs defined as procedural errors
• 20 software errors
• 3 GSE problems
• 14 incorrect limits and data-base errors
• 8 open PRs, 5 are pending verification as of 8
  December 1999
• Main deficiency of baseline CPT was that final
  flight s/w was not ready
• PSE (FDCmisc.), HK RSN (GPSmisc.), and TSM/RTSs
  are biggest changes
• Mitigated by two later CPTs (12/6/99 and 2/00),
  which use final flight software
• Subsystem reports and analysis
• Initial problems, mostly software, already
  identified, corrected, and verified
• Reports and analysis nearly complete all known
  results are nominal no problems discovered that
  require hardware modification

14
Functional Test

• Used between moves and after environmental tests
  to verify all EO-1 functions
• Performed after transportation to GSFC and
  planned for after transport to launch site
  (omitted if CPT performed).
• Performed between vibration and TV tests and
  after TV test.
• Performed (with some modification) at the launch
  pad
• Objective is to test every wire and connection on
  EO-1
• Much of the software is tested as a by-product
• The functional test is a subset of the CPT tests
• Maximum use of well-developed tests
• The functional does not include tests that are
  exclusively performance tests

15
System Testing During Thermal/Vacuum Test

• Each CPT section conducted at both hot and cold
  plateaus
• Since hot/cold plateaus are only 24 hours long,
  CPT was not continuous at each plateau
• Functional conducted during pumpdown and at the
  end, just before chamber break
• 150 PRs, 80 are closed and verified as of 8
  December 1999
• All critical PRs closed
• Nadir-deck heater short was only hardware anomaly
• Several new software errors discovered
• Some procedure errors discovered

16
Post-Environmental CPT Preliminary Results

• Conducted December 6-9, 1999
• All flight hardware and software configuration
  except
• LFSA not mounted
• Will probably modify HK RSN software
• Will fuse LVPC non-critical power services
• Several table parameters to be modified for
  flight (telemetry filter, TSM, RTS)
• Except as noted in PRs, all results reviewed to
  date are nominal
• 20 PRs initial assessment (through CPT5)
• 6 were procedure or operator errors (already
  corrected)
• 4 were nominal conditions initially perceived as
  problems
• 4 GSE problems (3 FEDS and 1 GPS simulator)
• No hardware problems
• 5 PRs were software problems
• Unexpected safehold entry due to M5 warmstarts (2
  PRs)
• Deployment may re-start if HK RSN has a warm
  start (statistics are 1-in-50 known cause not
  necessary to fix)
• Unexpected WARP error (status?)
• 1 RTS-table error
• 1 not investigated

17
Environmental Testing Overview
18
LV Separation Shock

• Test Configuration (Building 7 Test Cell 040 )
• Flight-Type Payload Attach Fitting Clamp Band
  System (Boeing)
• Spacecraft supported from crane and PAF dropped
  on to foam pad
• Spacecraft fully configured and powered in
  launch mode
• Hyperion and LFSA not electrically integrated
  (not powered during launch)
• Thirty one tri-axial accelerometers used to
  monitor shock levels
• Time History plots
• Shock response Spectra Plots, Q10, 100-2000Hz
• Two Firings performed
• Results
• All responses, with exception of two on Nadir
  deck, well within specified environments
• Conclusions
• Environment benign
• Aliveness test indicated no anomalies

19
Vibration Test

• Test Configuration (Building 7 Test Cell 040 )
• Flight-Type Payload Attach Fitting Clamp Band
  System (Boeing)
• MAP Spacecraft vibration test fixture (force ring
  transducers)
• Spacecraft fully configured and powered in
  launch mode
• Thirty three tri-axial accelerometers used to
  monitor shock levels
• Three axis sine sweep vibration to protoflight
  levels (1.25 x Flight Limit Loads)
• Results
• No physical damage to structure No shift in
  Frequency pre- versus post-vib
• Conclusions
• Spacecraft Structure qualified for flight
• Coupled Loads Analysis model updated and
  submitted to Boeing for VLC base banding
  (fundamental freq. increased over predicts)
• Verification Coupled Loads results confirm all
  margins positive
• Aliveness test indicated no anomalies

20
Acoustic Test

• Configuration (Building 7 Acoustic Chamber)
• Spacecraft lifted off dolly and suspended by
  chamber crane
• Thirty three tri-axial accelerometers used to
  monitor shock levels
• Spacecraft fully configured and powered in
  launch mode
• Boeing Delta 7320-10 Levels (Overall 141.1 dB)
• Protoflight levels (one minute duration)
• Results
• No physical damage to structure
• Protoflight level achieved (within tolerance)
• Responses well within specified component test
  levels
• Conclusions
• Spacecraft Structure qualified for flight
• Aliveness test indicated no anomalies

21
EMI/EMC Test

• Configuration
• Spacecraft fully configured with exception of
  solar array removed
• Radiated Emissions Testing Radiated
  Susceptibility (GEVS Range/ELV) Testing
  performed in GSFC EMI facility (performed pre TV)
• Power Profile Testing Conductive Emissions
  Testing performed in EO-1 Clean Tent (Performed
  post-TV)
• Results
• RE RS fully characterized and all exceedences
  documented
• CE fully characterized and all exceedences
  documented
• RF Hat Coupler Leakage tested and verified
• Conclusions
• All appropriate Waivers generated and evaluated
  as acceptable
• See Special Topic at the end of this section

22
Magnetic Survey Test

• Configuration
• Basis of Test Procedure was ACE, TRMM, WIND,
  Polar and XTE
• Static magnetic field measure via fixed
  magnetometers setup by GSFC Magnetics Group. Test
  conducted by GSFC MAG
• Performed in Building 29 SSDIF
• Performed during period of low activity in
  facility to reduce background magnetic noise
  levels
• Spacecraft suspended by crane using specially
  designed sling that minimized use of magnetic
  parts and which maximized distance from crane
  hook (14ft)
• Results
• Total Static (DC) magnetic field of Satellite
  fully characterized
• Conclusion
• DC Magnetic dipole of 9Am2 is higher than
  expected
• Magnetic Torque Bars have sufficient margin to
  compensate -- analysis in process

23
RF Compatibility Test

• Configuration (S-Band X-Band RF Hat Coupler to
  Compatibility Test Van)
• Demonstrated the operational capability of the
  command, telemetry and tracking data
  communications interface between the Ground
  Network (GN), Space Network (TDRSS), and Mission
  Operations Center
• Spaceflight Tracking Data Network (STDN-408.1)
  Procedure used
• Phase 1 RF Network Compatibility Testing WOA-410
• Measure the physical characteristics (Group
  delay, frequency, Bit Error Rate, carrier
  suppression) of the EO-1 RF System SN/GN via
  CTV only
• Phase 2 RF End-to-End System Test RF TDRSS
  Compatibility Test WOA-590 (CTV used as the
  interface between the spacecraft MOC)
• Verify that the MOC has the capability of
  commanding the spacecraft and receiving real-time
  playback housekeeping telemetry through the GN
• Verify that real time housekeeping telemetry may
  be received by the MOC at 2Kbps over the SN

24
RF Compatibility Test

• Results
• All S-Band X-Band characteristics verified
• Awaiting final CTV Test report
• Conclusion
• EO-1 RF Subsystem is compatible with the Ground
  Network Space Network

25
Thermal Balance /Thermal Vacuum Test

• Configuration (Building 7 Thermal Test Chamber
  238)
• EO-1 Satellite fully configured (LFSA not
  electrically integrated non-flight
  calorimeters) with all flight MLI Thermal
  blankets installed
• Demonstrate that the EO-1 Satellite and payloads
  can operate satisfactorily in all functional
  modes for the mission, at 10C beyond the hot
  cold extremes predicted for orbit
• Perform a complete Comprehensive Performance Test
  the hot cold plateau (distributed). Perform
  additional Performance tests throughout TB/TV
• Demonstrate the EO-1 satellite operates
  satisfactory in safe hold (24 hours)
• Demonstrate satisfactory operation of the EO-1
  Spacecraft and payload thermal control systems
• Verify the EO-1 Spacecraft and payload thermal
  models
• Verify the EO-1 Satellite meets contamination
  requirements
• Assess Instrument self-contamination and outgas,
  if necessary
• Provide opportunity for MOC to run flight
  operations procedures

26
Thermal Balance Thermal/Vacuum Test
27
Thermal Balance Thermal Vacuum Test
28
Thermal Balance Thermal Vacuum Test

• Results
• All objectives met
• Nadir Deck Heater Service Anomaly (see Special
  Topics)
• Demonstrated nominal operation of EO-1 Spacecraft
  and payloads at temperature extremes
• Demonstrated TCS for Satellite and payloads
• Excellent correlation with Satellite Thermal Math
  Model
• Contamination of EO-1 Satellite demonstrated to
  be well within limits however ALI out-gassing
  phenomenon which occurred during payload testing
  materialized at Satellite level. Impacts initial
  On-Orbit bake-out planned
• Conclusion
• EO-1 Satellite qualified for On-Orbit operation
• Post Thermal Vacuum Functional test was nearly
  flawless

29
Alignment

• Pre Environmental
• Alignment Characterization performed at the
  Spacecraft Bus level to establish primary
  reference frame
• Map Master Reference Cube (MRC) and S/C primary
  datums
• Map Secondary Reference Cube (SRC) on IRU to IRU
  Primary Reference Cube
• Map MRC to SRC
• Alignment Performed at Satellite Level
• ALI AC co-aligned to within 2 Arc-minutes
• Post Environmental
• IRU/SRC to AC co-alignment verified following
  vibration (no measurable movement - lt15 arcsec)
• AC to AST alignment verified post vibration
  (measurable rotation shift assessed)
• IRU SRC will be mapped to the S/C MRC AST
  reference cubes at WTR
• ALI HSA reference cubes will be mapped to the
  S/C MRC at WTR
• AST to S/C MRC will be mapped to the S/C MRC at
  WTR

30
Mass Properties

• Configuration (performed prior to shipment using
  Miller Mass Props Table)
• EO-1 Satellite mass has been measured several
  times throughout environmental program
• 98.4 of components measured (Hydrazine Fuse
  Plugs)
• Projected total mass of 573 Kg is 15 Kg under
  flight allocation (588 Kg)
• Detail Solid model (PRO-E model) used to
  calculate CG and inertia values prior to
  verification on Miller Table
• X Y CG values to be measured Z CG value to be
  calculated using Solid Model
• Results
• Total Mass allocation met
• CG requirement is projected to be within
  established requirement

31
X-Band Phase Array Near Field Test

• Configuration (Performed in Building 29 SSDIF)
• Performed at component and Satellite level
• Satellite level testing performed while in
  Horizontal Configuration on Arronson Table
• Satellite fully configured
• Verify Antenna Pattern Effective Isotropic
  Radiated Power (EIRP)
• Verify Attitude Control System can dynamically
  command the antenna to known beam positions
• Results
• Antenna pattern consistent with component level
  pre-environmental
• EIRP measured at 57 dBm meets requirement
• Demonstrated ACS can dynamically command antenna
  beam to known position
• Conclusion
• X-Band Phase Array performance verified at the
  Satellite level

32
X-Band Phase Array Near Field Test
NF Scanner in Position in Front of the XPAA
During NF3
33
End-To-End Instrument Test

• Objective
• Demonstrate that the ALI and Hyperion survived
  the environmental test program without
  degradation to the imaging quality.
• Hyperion VNIR characterization test exercises the
  Hyperion/WARP interface at high scene signal
  levels and provides data to determine the
  presence of signal dependent noise
• Configuration also allows verification of the ALI
  aperture cover redundant deployment mechanism
  (HOPAs)
• Configuration also allowed for the change out of
  the HSA connector post prior to start of test
• Configuration (Satellite mounted on Aronson table
  in SSDIF)
• ALI Hyperion ambient imaging test and Hyperion
  VNIR characterization
• Satellite mounted on Aronson table in the
  horizontal configuration (X down), instruments
  aligned to the MIT/LL optical bench (used at
  MIT/LL during ALI instrument level verification)

34
End-To-End Instrument Test

• Results
• ALI aperture cover mechanisms and HOPAs were
  inspected and verified to the extent possible
  (note one PR was written subsequent test
  verified remaining HOPAs)
• ALI and Hyperion optical performance is
  consistent with pre-satellite environmental
  performance measured at MIT/LL and TRW - no
  significant change in image quality
• Hyperion VNIR characterization was performed
  successfully, but max signal level limited to
  approx. 1500 counts. Analysis in progress. ALI
  reports NO effect from Hyperion observing noisy
  scene (chopper test).
• Conclusion
• Both ALI and Hyperion optical performance
  NOMINAL. No effect from S/C environmental testing

35
End-to-End Instrument Test
36
Special Topic EMI / EMC Testing Rationale
37
EMI/EMC Testing Requirements

• Based on MIL-STD-461C in accordance with the
  procedures of MIL-STD-462 and amplified by the
  following documents
• NASA/GSFC GEVS, BOEING MELV Mission Spec. and WTR
  MSPSP
• Provides guidance on acceptable EMI/EMC levels
  for the launch site environment
• MSFC On-Orbit RF Environment Doc.
• Provides a summary report on radiated emissions
  expected at the EO-1 mission orbit
• Litton SSO Doc. AM-149-0020(155)
• Provides a specific tailoring of MIL-STD-461C
  limits to the EO-1 spacecraft
• SAI-SPEC-158, EO-1 Verification Plan and
  Environmental Spec.
• Specifies the requirements for verification and
  test of the EO-1 spacecraft and its components
  and subsystems

38
Critical EO-1 Subsystems that were tested prior
to S/C IT
39
EO-1 Spacecraft-Level EMI/EMC Test Results
40
Spacecraft EMI Tests Were Performed to Verify
Compliance With Mission Requirements, and
Subsystem Waivers

• RF Hat Coupler Leakage Test
• Verified acceptable RF levels with transmitters
  radiating into hat couplers (80-90 dB?V/M S-band,
  62.7 dB ? V/M X-band)
• RS testing at GEVS levels
• Verified acceptable susceptibility (up to 5 V/M)
  from 1KHz to 12 GHz
• RS testing at WTR/MELV levels
• Verified acceptable susceptibility (up to 24 V/M)
  from 14KHz to 40 GHz
• NB/BB RE testing, per MIL-STD-461C
• Verified acceptable emissions levels(exceedences
  observed reflect GSE and instrument noise)
• Radiated Self-compatibility testing
• Verified S-Band receivers compatibility with
  X-Band transmitter
• Transient Tolerance test
• Verifies power bus ripple spec. (1.5 Vpp) under
  worst-case operating modes
• Power Profile test
• Verifies spacecraft self-compatibility under
  various modes of operation

41
EO-1 EMI/EMC Testing Summary

• Subsystem testing, per MIL-STD-461/ LSSO
  AM-149-0020(155) exceedences observed, but will
  not effect performance (waivers generated)
• Satellite testing is compliant per GEVS, MSPSP
  and Boeing MELVS specs
• Some exceedences were observed at RE satellite
  level tests, but does not affect satellite or
  instrument performance at the launch site, or
  on-orbit
• Based on testing to date, satellite operates
  nominally with all subsystems and instruments in
  worst-case modes

After reviewing the satellite performance and the
test data, the project concludes that the
satellite is ready for shipment to the launch
site.
42
Waivers
43
Waivers
44
Waivers
45
Waivers
46
Waivers
47
Waivers
48
Special Topic Electrical Testing of Deployables
49
Electrical Testing of Deployables

• Reason for Topic
• Verification of sufficient testing of deployables
• Summary of Topic
• Deployables tested at subassembly level (S/A on
  test stand with flight HOPS), and spacecraft
  level (S/A on spacecraft, spacecraft initiating
  deployment). Final testing to include
  deployments and verifications of fusing/wiring
  changes.
• Ramifications of Topic
• Adequate testing of the HOP actuator interfaces
  ensure proper deployment of the solar array
  on-orbit
• Recent configuration change due to the addition
  of redundant fusing requires additional testing

50
Electrical Testing of Deployables

• Previous testing of Deployables at S/C level
• Array deployed multiple times on test-stand using
  EGSE to activate Flight HOPS
• Array deployed at Swales prior to shipment to
  GSFC (old wiring) with the spacecraft performing
  an end-to-end validation
• Array pop and catch at GSFC following vibration
  and acoustic testing (old wiring) with the
  spacecraft performing an end-to-end validation
• Multiple testing of electrical deployment
  circuits using HOP simulator (old wiring) which
  fully tests circuits and software
• Testing of electrical deployment circuits during
  TVAC (old wiring) using HOP simulator

51
Electrical Testing of Deployables

• Planned testing of Deployables at S/C level
• Following completion of new wiring the entire
  flight interface will be tested to verify
  functionality of both primary and redundant paths
  on the flight hardware using GSE which simulates
  the HOP loads (12/3/99)
• All fuse plugs will be tested with the system,
  verifying actuator motion on both primary and
  redundant sides (week of 1/10/00 or earlier)
• The remaining solar array deployment will be
  conducted with the flight fuse plugs installed,
  on the redundant side with no EGSE in the loop
  (at launch site)
• Conclusions
• Remaining testing planned will guarantee system
  functionality on-orbit
• Finish with pop and catch at VAFB on February
  23, 2000

52
EO-1 Deployment History