LRO Preliminary Design Review
Payload Systems Overview
Arlin Bartels
Mary Reden
Leslie Hartz
Joanne Baker
Stan Scott
Tom Jones
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
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Payload presentation intro
• Instrument teams successfully completed their PDRs in the
Sept/Oct 2005 time frame
– More detailed discussion and current RFA status at end of
presentation
– CDs of Instrument PDR packages, RFAs and resolution, and other
pertinent documentation included in your review package materials
• The Instrument PDRs focused on the Instrument-specific
design details, without placing the Instruments within the
LRO Mission context
– “Down and in” from the Instrument <-> Spacecraft interface
– Intra-instrument design aspects
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Payload presentation scope
• This presentation intended to demonstrate how the
Instruments are technically managed and integrated as a
payload as part of the Orbiter
– “Up and out” from the Instrument <-> Spacecraft interface
– Inter-instrument systems engineering aspects affecting all Instruments
• Instrument teams available for splinter during PDR to address
any Instrument-specific questions
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Presentation Agenda
• Payload Complement Overview
– Instrument Summary
– Orbiter Accommodations Overview
• Systems Engineering Overview
–
–
–
–
–
–
–
–
Requirements flow down and verification
Technical Resource Allocations (mass, power, data volume)
ICDs and other documentation status
Risk Management
Schedule
Safety
Operations and Ground Segment
I-PDR RFA Summary
• Backups
– Instrument development update
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Payload Complement
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LRO Payload Complement has high heritage
•
•
The LRO Payload complement consists of six instruments selected by NASA
HQ for contribution to the LRO Mission Level 1 science objectives and one
technology demonstration payload of opportunity added in early 2005
Instruments selected for high maturity, relatively low risk, with heritage from
(multiple) previous missions
Instrument
Heritage Predecessor Most Recent Mission Heritage Current Status
CRaTER
IPS
POLAR (GSFC)
Diviner
MCS
MRO (JPL)
LAMP
Alice
New Horizons (APL)
LEND
HEND
MO (JPL)
LOLA
MLA
MESSENGER (APL)
LROC
CTX (NAC) and
MARCI (WAC)
MRO (JPL)
Mini-RF
Forerunner
Payload Systems Mgmt Overview 2006-02-08
Launched 2/1996;
mission still operational
Launched 8/2005;
enroute to Mars
Launched 2/2006;
enroute to Pluto
Launched 4/2001;
mission still operational
Launched 8/2004;
enroute to Mercury
Launched 8/2005;
enroute to Mars
Chandrayaan (ISRO) Forerunner PDR 12/05/05
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Payload summary and teams
INSTRUMENT
CRaTER
Cosmic Ray Telescope
for the Effects of Radiation
DLRE Diviner Lunar
Radiometer Project
SPONSORSHIP
PI: Harlan Spence, BU
IM: Rick Foster, MIT
MEASUREMENT
EXPLORATION BENEFIT
Tissue equivalent response to
radiation
Safe, high performance, lighter
weight space vehicles
Better than 500m scale maps
of temperature, surface ice,
minerology
Determines conditions for systems
operability, resource including
water-ice location
Maps of frosts in permanently
shadowed areas, etc.
Locate potential water-ice on the
surface, image shadowed areas
Maps of hydrogen in upper 1
m of Moon at 10km scales
Locate potential water-ice in lunar
soil
~50 m scale polar topography
at < 10 cm vertical, roughness
Safe landing sites and surface
navigation
1000’s of 50cm/pixel images
(125km2), and entire Moon at
100m visible, 400m UV
Surface Landing hazards and some
resource identification
X&S-band Radar imaging and
radiometry
Demonstrate new lightweight SAR
and communication technologies,
locate potential water-ice
ISE: Bob Goeke, MIT
PI: David Paige, UCLA
IM: Wayne Hartford, JPL
ISE: Marc Foote, JPL
LAMP
Lyman-Alpha Mapping
Project
PI: Alan Stern, SwRI
LEND
Lunar Exploration
Neutron Detector
PI: Igor Mitrofanov, IKI
IM: Ron Black, SwRI
ISE: Dave Slater, SwRI
Deputy PI: Roald Sagdeev, UMD
IM: Anton Sanin, IKI
ISE: Maxim Litvak, IKI
LOLA
Lunar Orbiter
Laser Altimeter
PI: David Smith, GSFC
Co-PI: Maria Zuber, MIT
IM: Glenn Jackson, GSFC
ISE: John Cavanaugh, GSFC
LROC
Lunar Reconnaissance
Orbiter Camera
PI: Mark Robinson, Northwestern
University
IM: Scott Brylow, MSSS
ISE: Mike Caplinger, MSSS
Mini-RF
Technology
Demonstration
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PM: Bill Marinelli, NAWC
DPM: Dean Huebert, NAWC
PI: Chris Lichtenberg, NAWC
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Instrument Accommodation Overview
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LRO Payload Complement
Diviner
CRaTER
LROC WAC
LOLA
LAMP
LROC NAC’s
Mini-RF
LEND
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Driving instrument accommodation requirements
being met
•
Non-trivial driving requirements to fully accommodate Instruments
– Field-of-View
• No impingements in Field-of-View
• No impingements in stray light Field of Regard
– Thermal
• Radiator view factors to space for thermally isolated Instruments
• Location constraints for thermally coupled Instruments
• Sun avoidance in Field-of-View
– Alignment
• Optical instruments requiring high pointing accuracy
• Ability to co-register the two LROC NACs, as well as LOLA to one or both NACs
– Handling/I&T
• Purge and GSE access during ambient Observatory I&T
• Instruments need to be able to place optical cubes/surfaces for alignment
verification at orbiter level
• Ability to accommodate any required targets during Observatory thermal vacuum
•
Two post-PDR accommodations trade studies
– Diviner placement in Y-axis (assessing thermal margin in current location)
– Mini-RF placement/orientation (flexibility for operational enhancement)
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Instrument accommodation summary
S/C Body
Mounted
Thermally
Coupled
CRaTER Instrument
X
X
Diviner Instrument
X
Diviner Electronics
X
Instrument
LAMP Instrument
OB Mounted
X
X
X
LEND Instrument
Thermally
Isolated
X
X
X
LOLA Instrument
X
X
LOLA Electronics
X
X
LROC NAC #1
X
X
LROC NAC #2
X
X
LROC WAC
X
X
LROC SCS
X
X
Mini-RF Antenna
X
Mini-RF E-box #1
X
X
Mini-RF E-box #2
X
X
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Payload Systems Engineering
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Payload Systems Engineering Scope
• Payload System Engineering addresses both instrument to
S/C interfaces, inter-instrument dependencies/constraints and
insight/oversight into intra-Instrument Systems Engineering
– S/C Interfaces and Driving Requirements
•
•
•
•
•
Requirements flowdown from Level 1 to Level 3
Coordinate accommodations issues between Instruments and Spacecraft
Monitor and resolve technical resource allocation issues
Oversee ICD development between Instruments and Spacecraft
Coordinate specifications, action items and technical giver/receiver items
(e.g., math models)
• Oversee instrument integration into Orbiter and Orbiter-level test
program
– Payload Inter-Instrument Issues
• Operational constraints between Instruments
• Co-alignment of Instruments
• Orbiter-level issues affecting multiple Instruments
– e.g. Pointing/jitter, electromagnetic compatibility, mechanical Loads and
contamination control
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Inst requirement flowdown begins with the Level 1’s
Objectives
Find Safe Landing
Sites
Locate Potential
resources
Life in Space
Environment
New Technology
LRO Requirements
LRO Data Products
M30 - Topography grid
LOLA - Global DEM
M40 - Topo resolution
LROC, LOLA - low res images
M80 - Surface Features & hazards
LROC, LOLA, DLRE - hi res images, roughness,
slopes, rock abundance
M50 - Surface Temp
DLRE - Surface Temp Maps
M60 - Images of PSRs
LAMP, LOLA - Albedo maps, topo
M70 - Sub/Surface ice
LAMP, LEND- ice&frost density maps
M90 - Polar Illumination
LROC, LOLA, DLRE - illum maps
M100 -Regolith Resources
LROC, DLRE – IR/UV/Vis images
M110 - Hydrogen mapping
LEND - Hydrogen density maps
M10 - Radiation Environ
CRaTER, LEND - LET Spectra
M20 - Radiation on Human equiv
tissue
CRaTER - Tissue equiv LET Spectra
P160 - Technology demo
Mini-RF – N/A (no formal L1 data product)
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LRO requirements flowdown to L2
LRO Level 1 Requirements
ESMD-RQMT-0010
Project Requirements
Measurement Requirement
Instrument Specific Expecte
LRO Mission Requirements
Document
431-RQMT-000004
Allocations
Electrical Spec
Mechanical Spec
Thermal Spec
Mini-RF
LROC
LOLA
LAMP
LEND
CRaTER
Diviner
Operations
•
For “Level 2 family” of requirements
documents, linkage is maintained between
Project Office and Instruments
–
Level 2
Performance & SOC
Requirements
Contamination
Radiation
Mission Assurance
–
–
Launch Vehicle
Instrument teams responsible for generating
their Level 2 Performance Requirements –
“Instrument Requirements Document” –
(IRD)
LRO Project responsible for generating MRD
including driving constraints from
Instruments
Instrument teams review and sign the MRD,
LRO Project reviews and signs the IRDs
Spacecraft, Instrument & Ground
Level 3 Requirements Documents & ICDs
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Instrument L2 and L3 requirements
management is a shared responsibility
L1 Requirements
controlled at HQ/RLEP
PI Institution
Verification
Tracking
Database
L2 Instrument Requirements
Controlled jointly by LRO Project &
PIs
LRO Project Oversight
LRO Project Level
Verification
Tracking
Database
Requirements
Documents in
LRO CM
L3 Instrument Requirements
Controlled at PI Institution
LRO Project Insight
Instrument
Verification
Matrix
LRO Project verification of
Insturment Level 2 compliance
Legend
Instrument Responsibility
Project Responsibility
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Requirements Verification (1)
• Payload Systems works with Mission Systems to ensure a
cohesive verification plan across all payload elements
• Each Instrument is expected to have verified all performance
requirements at delivery to LRO, except as negotiated with
LRO Project Office.
– Instrument verification plans require LRO Project Office review and
approval
– Any verification requirements which are deferred to orbiter level are
documented in the Instrument PAIPs and/or verification plans and be
agreed-upon prior to beginning official test sequence by LRO Mission
Assurance
– Any required waivers are resolved at time of occurrence rather than at
delivery
– Instrument requirements which are re-verified by test at orbiter level
during orbiter I&T campaign are captured in LRO I&T Plan
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Requirements Verification (2)
• Requirements expected to be deferred to orbiter level of
assembly for verification include:
– Acoustics
– Pyroshock
– Failure-free operating hour (FFOH) and total operating hour
requirement
• Some requirements can only be verified at the Orbiter level
– Instrument-Instrument coalignment
– System-level EMC
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Pointing and Alignment Verification
•
Draft orbiter alignment and optical metrology plan includes all
Instruments and tech demos
– All optical Instruments have alignment cubes for theodolite verification
– Installation and initial alignment knowledge accounts for 1g / ambient
temperature installation
– All Instruments are referenced to the LRO master optical cube on the optical
bench to which LROC, LOLA and LAMP are mounted, and by extension to
the primary star tracker which is also mounted to the optical bench
– Inter-instrument co-alignment is explicitly verified for:
• LROC NAC-NAC Co-alignment
• LOLA co-registration to LROC NAC(s)
•
Pointing and alignment at orbiter level of assembly performed by
integrated system analysis:
– Alignment verification over temperature will be performed by integrated
STOP analysis in conjunction with Instrument teams
– Jitter analysis to include dynamic torque disturbances
•
Each Instrument has ability to do absolute alignment verification during
post-launch commissioning and calibration phase.
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Payload Technical Resource Allocations
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Payload Technical Resource Allocations
•
•
Starting point for Instrument technical resource allocations was the Instrument
proposal Current Best Estimates (CBEs)
Allocations (CBE + contingency held by Instruments) then established per GSFCSTD-1000, tailored by Instrument heritage
– Tailoring process described in TRA, 431-SPEC-000112.
•
Instrument CBE status tracked monthly
Establishment of TRA’s
Instrument CBE’s in
original proposals
LRO technical
evaluation of CBE’s
LRO assessment of
instrument maturity
Initial resource allocations,
including margin and reserve,
generated
Technical resource allocations
formally documented in
431-SPEC-000112.
Revision of TRA’s
Formal requests for revised
allocation submitted to LRO
*
Instrument teams refine design
and assess sufficiency of
allocations
* Process for revising allocations
documented the LRO SEMP (431-PLAN000005).
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Instruments have managed mass responsibly to date
M-PDR
CBE (kg)
Allocation
(kg)
9.6
12.4
12.3
15.3
24%
Mini-RF
10.4
10.5
10.5
12.6
20%
LROC (3)
13.6
13.6
14.4
16.5
14%
CRaTER
5.6
5.3
5.3
6.4
21%
Diviner
9.0
9.8
11.0
12.0
8%(1)
LAMP
5.0
4.9
5.0
5.3
6%(1)
LEND (3)
23.2
23.2
23.7
23.7
0%(2)
Instruments
Subtotal
77.2
79.7
82.2
91.8
LOLA
Proposal
CBE (kg)
Current
M-SRR
CBE (kg)
Instrument
Contingency (%) (4)
(1)
Lower-than-standard margins for Diviner and LAMP are acceptable due to high degree of build-to-print heritage (actual measured values)
(2)
Agreement with LEND team is that they will design their collimator/shield design to allocation. Requirements can be met within allocation.
(3)
Mass liens being tracked: LEND final dressout, LROC thermal mass
(4)
Percentages computed against CBE per GSFC Gold Book
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Instruments have managed electronics power
dissipation responsibly to date
Nominal Instrument operational power draws, not including spacecraft mounted heaters
Component
Proposal
CBE OAP
(W)
Current
Contingency
%(2)
M-SRR
CBE
OAP (W)
M-PDR
CBE
OAP (W)
Allocation
OAP (W)
5.1
5.1
9.0
76%
22.0 (3)
19.4
26.6
37%
CRaTER
6.9
Diviner
11.0
LAMP
4.3
4.4
4.8
4.9
2%(1)
LEND
9.8
9.0
9.0
13.0
44%
LOLA
26.2
34.3 (3)
35.7
39.4
10%
LROC
22.0
24.0
20.0
26.4
32%
Total
80.2
98.8
94.0
119.3
(1): Per SwRI Lunar Terminator Sensor Internal Peer Review, both sides of Lunar Termination Sensor (LTS) now simultaneously powered;
LAMP has yet not requested an allocation increase due to high build to print heritage with measured values from New Horizons Alice. LRO
holding lien for additional allocation pending LAMP LVPS testing.
(2) Percentages computed against CBE per GSFC Gold Book
(3) CBE increase includes LRO-directed C&DH design changes to accommodate LRO C&DH bus and Diviner operational heaters
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Daily Data Volume dominated by LROC/NAC
LRO Daily Data Volume (Gbits)
40.84 16.61
515.13
LROC NAC
LROC WAC
Other 5 Instruments combined
Note – based on 16 NAC pairs per orbit max.
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Interface Control Documents (ICDs) and
Other Instrument Document Status
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Interface Control Documentation (1)
• Each Instrument’s interface requirements are captured in four
separate ICDs generated by the engineers responsible for
implementing the interface
– Data, Electrical, Mechanical, Thermal
• All ICDs are controlled through the formal LRO
Configuration Management (CM) process.
• Additional mechanical interface details are captured in
instrument-generated Mechanical Interface Drawings (MIDs)
– Although generated by Instrument teams, LRO Project Office will
still review/sign the MIDs to ensure proper communication
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ICD release status on track
CRaTER Diviner
LAMP
LEND
LOLA
LROC
Mini-RF
Data ICD
Electrical ICD
Mechanical ICD
Mechanical I/F Dwg (MID)
Thermal ICD
Completely through CM CCB and all signoffs
Currently in CM CCB for signoff, final release expected by 2/17
Mature draft currently, expected release by 3/15
GSFC to generate with instrument input
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Other Instrument Docs at PDR level
CRaTER
Diviner
LAMP
LEND
LOLA
LROC
Mini-RF
IRD
PAIP / Mission Assurance Plan
Contamination Control Plan
Risk Management Plan
Verification Plan
Data Management Plan
Released through Instrument CM system
Preliminary versions released; finals due at I-CDR
In Process
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Risk Management
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Risk management is taken seriously by both the Instrument
teams and payload systems
•
•
Instrument-generated Risk Reports are provided on a monthly basis
Payload Staff reviews each reported risk and criticality (Consequence and
Likelihood) as assigned by the instrument teams
– Periodically reviewed with Instrument teams during status telecons
– 76 instrument-generated risks are currently tracked in the payload risk
management database
•
In addition, Payload Systems Team performs continuous risk
management assessment of the instruments, payload as a whole, and
interfaces to the S/C.
– 12 Payload System-generated risks are currently tracked in the payload risk
management database in addition to the instrument-generated risks
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Payload risk management is incorporated into
Project risk management
•
All Moderate (yellow) and High (red) risks are then entered into the
project risk tracking tool
– No High (red) risks identified at Instrument or payload level
– 21 Moderate instrument risks
– 4 Moderate payload systems risks
•
LRO Project Management assesses these risks and determines which to
elevate to a new, Project-level risk
– Full risk tracking spreadsheet is included as back-up
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Risk Management Summary
Total
Risks
Risks
Risks
Reported Reported
at I-PDR Since PDR Tracked
Risk Mgmt
Plan No.
Risk
Matrix
Format
CRaTER
32-01202
4x 4
1
0
1
1
0
0
0
Diviner
JPL D-32940
5x 5
6
10
16
7
5
0
4
LAMP
Contained in
PAIP-05-15-11239
5x 5
6
1
7
3
2
0
2
LEND
TBD
5x 5
6
0
6
3
3
0
0
LOLA
LOLA-PLAN-0007
5x 5
10 *
2
36
26
8
0
2
LROC
M SSS- LROC-DID- 7071
5x 5
6
0
6
3
2
0
1
Contained in
PAIP
431-PLAN-000181
5x 5
4*
N/A
4
1
3
0
0
Payload
431-PLAN-000193
Systems
5x 5
n/a
12
12
8
4
0
0
39
25
88
52
27
0
9
Mini-RF
Total
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(Top 10 only)
(PDR 3/1/06)
Arlin Bartels – Payload Systems
Low
Moderate
(Green) (Yellow)
High
(Red)
Retired
Overall
trend
since IPDR
32
Payload Systems Risk Focus Matrix
(C x L) Trend
5
L
I 4
K
E
L 3
I
H
O 2
O
D
LRO-PLS-10
Risk ID
Approach
3x4
LRO-PLS-10
M
Heritage Thermal Design
Obsolescence
4x2
LRO-PLS-12
M
Mini-RF Radar Interference with
other instruments
4x 2
LRO-PLS-13
M
NAC Co-alignment verification
over temperature deferred to
Orbiter level
3x1
LRO-PLS-03
M
Operational conflicts between
Instruments
2x2
LRO-PLS-07
M
Payload EMI/EMC Conducted
Emissions
LRO-PLS-13
LRO-PLS-07
LRO-PLS-12
LRO-PLS-03
1
1
2
3
4
Risk Title
5
CONSEQUENCES
Criticality
High
Med
Trend
Decreasing (Improving)
Increasing (Worsening)
Unchanged
Low
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Approach
M - Mitigate
W - Watch
A - Accept
New Since Last Month
33
Payload Systems Risk Focus (1)
Rank, Title,
Criticality (C x L)
Risk Statement
Approach & Plan
Status
If: Analysis of the Lunar thermal environment
Mitigate
and Spacecraft accommodations demonstrates that Early exchange of thermal models; IDT's
heritage thermal designs are obsolete…
focused on this issue and working closely
with LRO thermal team. Interface
accommodation engineer assigned fullThen: there may be significant programmatic
time to work mech/therm instrument
impacts
interfaces
Decreasing
(Improving)
2. Mini-RF radar inteference with
Instrument operations
Crit: 4 x 2
If: Unexpected Mini-RF radiated emissions cause Mitigate
interference to other Instruments…
Get as much information about Mini-RF
antenna patterns (especially backlobes) as
possible ASAP and provide to Instrument
Then: Either Mini-RF or Instrument operations
teams for susceptibility assessment.
may be constrained on-orbit.
Unchanged
Initial info exchange
between Diviner and
Mini-RF complete.
Working with Mini-RF
team to get further
information.
3. NAC Co-alignment verification
over temperature deferred to
Orbiter level
Crit: 4 x 2
If: Unable to verify NAC co-alignment over
temperature by analysis with sufficient margin…
Unchanged
Integrated STOP
analysis to be complete
by M-CDR
1. Heritage Thermal Design
Obsolescence
Crit: 3 x 4
Med
Med
Med
Payload Systems Mgmt Overview 2006-02-08
Then: there may be significant programmatic
impacts (redesign or elaborate test)
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Mitigate
1) Optical Bench baselined as low-CTE
M55J composite
2) Perform integrated system-level STOP
analysis with LROC and Orbiter and
assess margin on alignment.
RTMM available for
Instruments; no
showstoppers to date
with Instruments
34
Payload Systems Risk Focus (2)
Rank, Title, Criticality
Risk Statement
Approach & Plan
Status
4. Operational conflicts between
Instruments
Crit: 2 x 2
If: Operational conflicts between Instruments are Mitigate
identified which cannot be solved by operational 1) Perform orbiter-level integrated jitter
workarounds
analysis including all sources of
dynamic torque disturbances
2) Develop ConOps as soon as feasible to
flush out any other possible
Then: either hardware redesign may be required
conflicts
or science may be impacted
Unchanged
Integrated jitter
analysis to be
complete by MCDR.
Instrument PIs to
sign off on ConOps.
5. Payload EMI/EMC Conducted
Emissions
Crit: 3 x 1
If: Instrument conducted emissions exceed
allowable levels
Unchanged
Heritage CE being
assessed by LRO,
will be sent to
Instrument teams for
susceptibility
evaluation
Low
Low
Payload Systems Mgmt Overview 2006-02-08
Mitigate
1) Get heritage info from Instrument teams
and see if tailoring of 461 for LRO is
Then: Compatibility issues between Instruments appropriate.
may arise
2)Have instruments assess susceptibility
against heritage test results from other
Instruments;
3)Conduct early EMI/EMC testing with
instrument power supplies
4) Assess feasibility of early compatibility
testing ASAP including LOLA, Diviner
and Mini-RF
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Instrument Risk Focus Matrix update from
Instrument PDRs
(C x L) Trend
5
L
I 4
K
E
L 3
I
H
O 2
O
D
1
NGIN Risk ID
Approach
Risk Title
4x3
LEND-05
W
Delay in NASA-FSA IA signoff
3x4
LROC-04
M
LROC Thermal
Environment/Design
4x2
LOLA-22
M
LOLA RTAX-S FPGA Part
Qualification (Note: also
tracked at Project Level)
2x4
MRF-03
M
Mini-RF Transmitter Interference
with Spacecraft
2x4
LAMP-01
M
LAMP EEE Parts Delivery
2x3
DLRE-05
M
Diviner susceptibility to Mini-RF
emissions (Note: risk
elevated to Payload Risk
level)
1x1
LOLA-28
M
Thermal Analysis schedule
impact on Beryllium
procurement
LAMP-01
Mini-RF-03
LROC-04
LEND-05
DIV-05
LOLA-22
LOLA-28
1
2
3
4
5
CONSEQUENCES
Criticality
High
Med
Trend
Decreasing (Improving)
Increasing (Worsening)
Unchanged
Low
Approach
M - Mitigate
W - Watch
A - Accept
New Since Last Period
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
36
Instrument Risk Focus (1)
Rank, Title, Criticality
(c x l)
1. Delay in NASA-FSA
Implementation Agreement
signoff
Crit: 4 x 3
Risk Statement
Approach & Plan
If: Unexpected delays in signoff of
LEND Implementation Agreement
occur…
Then: LEND development and delivery
will be impacted
Med
2. LROC Thermal
Environment design
Crit: 3 x 4
If: Thermal environment in current
orbiter configuration obsolesces earlier
LROC thermal design…
Then: Possible redesign may impact
LROC Programmatics may be affected
Watch
1)
Contact HQ for path forward and to
ensure that IA is on track
2)
Work with LEND team to resolve any
issues which arise on either the NASA
or Roscosmos side
Increased
IA has completed NASA review
and is being tranlated to Russian.
Will meet with LEND PI and HQ
to review language during PDR.
Mitigate
1)
Move to top focus of Payload team
2)
Provide Orbiter reduced model to
LROC team
3)
Ensure close contact between LROC
Thermal and LRO Thermal; assess
any driving requirements which can be
relaxed
4)
Establish mass lien with Systems
Increased
Mitigate
1)
Implement GSFC Center-wide plan
per NASA OLD and Actel
recommendation
Unchanged
GSFC Center-plan has been
agreed upon with processing
specified in procurements. Status
being monitored across Center by
SS&MA.
Med
3. LOLA RTAX-S FPGA Part
Qualification
If: RTAX-S FPGA part fails
qualification program…
Crit: 4 x 2
Med
Then: Backup part with inferior
performance will need to be used and
programmatics will also be affected
Change since I-PDR
Move to beta-gamma
configuration obsolesced
elements of PDR-era design.
However, significant progress
being made in redesign. TIM
held at LROC on 11/15 to assess.
Interface temperatures at bench
made more benign at LROC
request. Mass lien for radiators
established with systems.
Also tracked as Project-level risk (part used on Spacecraft as well)
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
37
Instrument Risk Focus (2)
Rank, Title, Criticality
(c x l)
4. Mini-RF Transmitter
Interference with Spacecraft
Crit: 2 x 4
Med
5. LAMP EEE Parts
procurement
Crit: 2 x 4
Risk Statement
Approach & Plan
If: Mini-RF transmitter interferes with
LRO S-Band Transponder operation…
Then: Either redesign will be required,
or Mini-RF operation may be
constrained
Mitigate
Unchanged
1)
Work with GSFC to complete RF Assessment ongoing.
interference assessment
2)
Implement operational constraints
on Mini-RF
If: EEE Parts procurement/delivery is
delayed
Then: Schedule slack will be reduced.
Mitigate
1) Put EEE Parts on order with project
concurrence ASAP.
Med
6. Diviner susceptibility to
Mini-RF emissions
Med
7. LOLA Thermal Analysis
schedule impact on Beryllium
procurement
Low
Change since I-PDR
If: Diviner is susceptible to Mini-RF
Emissions in current Orbiter layout
Decreased
EEE Parts put on order with
Project concurrence after I-PDR.
Part delivery schedule to date is
acceptable
Mitigate
1)
Assess Mini-RF Emissions and
Diviner susceptibility
Then: Either Diviner/Mini-RF will have 2)
Perform trade study of moving
to be moved or operational constraints
Diviner or Mini-RF
will need to be implemented
3)
Implement operational constraints
on Mini-RF
Increased
First information exchange
between Mini-RF and Diviner
complete, more information from
Mini-RF needed. Post-PDR trade
study to assess ability to move
Diviner in (-Y) directin away
from Mini-RF
If: LOLA Thermal Analysis is not
sufficiently complete to size radiator by
beryllium procurement target date
Decreased
Thermal analysis completed
without driving the beryllium
procurement date
Mitigate
1)
Expedite thermal analysis
Then: LOLA schedule may be delayed
due to long-lead procurement
Elevated to Payload-systems risk (other teams with similar concern)
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
38
Instrument Schedule Overview
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
39
LRO Instrument Schedule
1/30/06
2005
CY
1
2
3
4
5
6
7
8
9
10
11
1
2
2006
3
LAMP
4
PDR
1
2
CDR
2007
3
4
PER
Instrument Reviews
Lunar Terminator Sensor
Interface Simulator
Flight Unit
Flight Unit Integration & Test
Deliver Flight Unit to GSFC
4
PSR
CDR
Contingency
(62 days)
Ready for Delivery
CRaTER
PDR
CDR
Contingency
(52 days)
Ready for Delivery
PDR
D-PDR
CDR
PER
PSR
Contingency
(51 days)
Ready for Delivery
LROC
PDR
CDR
IDR
Contingency
(43 days)
Ready for Delivery
LEND
Mini- RF
3
Ready for Delivery
PDR
Instrument Reviews
Detail Designs (Optics, Elec, EGSE, Flt S/W, Mech,
Ther.)
Fab & Test (Optics, Elec, EGSE, Flt S/W, Mech,
Ther.)
Flt Unit Integration & Test/Calibration
Deliver Flight Unit to GSFC
LOLA
2
Contingency (162 days)
Diviner
12
13
14
15
Instrument Reviews
16
EM Model Fab & Test
17
Flight Unit
18
Flt Unit Integration & Test/Calibration
19
Deliver Flt Unit to GSFC
20
21
Instrument Reviews
22
Detail Designs (Laser/Optics,Elec, S/W, Mech,
23
Fab & Test (Laser/Optics, Elec, S/W, Mech, Ther)
24
Flt Unit Integration & Test
25
Deliver Flight Unit to GSFC
26
27
Instrument Reviews
28
Detail Designs (NAC,WAC,SCS,S/C Interface,GSE)
29
Parts Procurement/Screening
30
Fab & Test (NAC,WAC,SCS,S/C Mount, GSE)
31
Flt Unit Integration
32
Deliver Flight Unit to GSFC
33
34
Instrument Reviews
35
Lab Unit Development & Test
36
Elec. Instr. Simulator (Dev, Test, Delivery)
37
Eng. Unit (Dev,Test, Delivery)
38
Qual Unit (Dev, Test)
39
Flt Unit (Dev, Test, Calibration)
40
Deliver Flight Unit to GSFC
41
42
Instrument Reviews
43
Design, Development, I&T
44
Emulator Delivery to GSFC
Systems
2006-02-08
45Payload
Deliver
Flight Mgmt
Unit toOverview
GSFC
1
PDR
CDR
Contingency
(42 days)
Ready for Delivery
Instr. Selection
PDR
CDR
PER
Emulator
Arlin Bartels – Payload Systems
Contingency (30 days)
40
Instrument Delivery Schedule Considerations
•
•
All instruments status slack against 10/15/2007 delivery date for
programmatic tracking purposes
All instruments currently show the recommended schedule delivery
margin (1 month slack per year remaining)
– LAMP scheduled for delivery 2/2007, all others showing 7-8/2007 delivery
dates
•
•
Instrument schedules have credibility because most schedules have their
basis in heritage instrument actuals.
While I&T has a preferred order of integration, the modular nature of the
LRO design can accommodate any order of Instrument delivery
–
•
Integration flexibility in the event one (or more) Instrument delivers late.
Instruments will be integrated onto the Spacecraft after they have passed
both their post-ship functional (required) and risk-mitigation testing with
the flat sat setup (as time allows)
– Risk-mitigation testing will focus on database and procedure checkout so as
not to tie up orbiter I&T
– Pending eventual delivery dates, other risk mitigation testing will be
performed as time allows prior to installing Instruments on Orbiter.
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
41
Integrated Instrument Operations (flight
and ground)
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
42
LRO Operating Modes are straightforward
•
•
•
•
•
Majority of Instruments are (nadir-staring)
mappers which operate over entire orbit
No power constraints on Instrument operation
during nominal data taking
LAMP takes data pole-to-pole (dark side);
autosafes on bright side via LTS
LROC takes images pole-to-pole (bright side);
ground loads command NAC imaging per
specified targets
LROC only Instrument which will request LRO
to maneuver off-nadir
–
–
•
Will continue to look for possible operational
conflicts as designs mature
–
–
•
When off-nadir, the spacecraft attitude info will be
provided to other Instrument teams
Baseline is that non-nadir operations will occur no
more than three times per day
Possible Diviner/LROC interaction which can be
easily mitigated by operational workarounds
Data collection conflict between NACs and MiniRF (Spacewire bandwidth constraints)
Periodic calibration activities will be performed
in conjunction with Orbiter orbital stationkeeping maneuvers
Payload Systems Mgmt Overview 2006-02-08
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43
Instrument Post-launch Commissioning activities
being defined
•
•
Instrument Post-Launch commissioning requests captured in LRO
Calibration Plan #431-SCI-000393 and then ingested into ConOps
Due to short cruise phase, most instruments will not be operated during
cruise phase
– Both CRaTER and LEND will be turned on during cruise phase
•
Formal commissioning of Instruments will happen after lunar orbit
capture after Spacecraft health is verified and outgas is complete:
– Aliveness/Functionality
– “Absolute” alignment verification versus star tracker as function of
temperature, consistent with thermal constraints and available targets
– Special commissioning (sensitivity, focus) tests against dark sky and lunar
surface as required
•
Required outgassing time for each to be determined analytically via
analysis and fed into commissioning plan
– Will coordinate outgassing with Instruments in systematic way to reduce
cross-contamination (e.g., LAMP doors will stay closed until LROC
decontamination heater operations are complete)
•
LRO will settle into routine operations and periodic
stationkeeping/instrument calibration when outgassing and
commissioning are concluded.
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
44
Ground Segment MOC/SOC responsibilities have
been defined
• LRO is responsible for the Mission Operations Center
(MOC), the various Instrument teams are each responsible
for their own Science Operations Centers (SOCs).
• LRO Project is responsible for negotiating the MOC-to-SOC
ICDs with the Instrument teams, the Instrument teams are
responsible for negotiating the SOC-to-Planetary Data
System (PDS) ICDs, with LRO Project insight.
• Instrument SOC design reviews to be integrated into the LRO
Ground Segment Single Design Review, currently scheduled
for Oct. 2006.
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
45
SOC activities have started ramping up
• Most SOCs have had a kickoff TIM with the LRO Project:
–
–
–
–
–
–
–
CRaTER: still to be scheduled
Diviner: 11/10/2005
LAMP:
10/14/2005
LEND:
10/18/2005
LOLA:
9/13/2005
LROC:
10/3/2005
Mini-RF: 10/28/2005
• LRO Data Working Group (LDWG) has begun holding
biweekly meetings as well
• LROC SOC hosted peer review on 11/29/2005
– LROC SOC most complicated of all SOCs due to data volume and
target generation requirements
– Other SOC teams will assess the feasibility of a peer review for their
SOC
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
46
Payload Systems Engineering
Summary
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
47
Payload Systems Engineering Summary
• All requirements well understood and flowed down from
Mission Level 1 to Instrument Level 3
– Requirements verification planning at PDR level of maturity
• Instrument accommodations on the Observatory meet all
Instrument driving requirements.
• Instrument Technical Resource allocations are fixed and
appear to be adequate.
• ICDs and other documents are on track
• Risk management is being proactively performed at both the
Instrument and Payload level
• All Instruments showing adequate schedule margin
• Operations being considered during design phase
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
48
I-PDR RFA status
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
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Instrument PDR Summary
•
All six instruments completed successful PDRs in the Sept/Oct 2005 time frame
–
•
•
Mini-RF PDR scheduled for March 2006
RFAs due for closure at I-CDRs
LRO Project audits and monitors closure status in IIRT system
–
For Diviner and LAMP, RFAs first closed in institution’s systems process and then in official IIRT system process
LRO Instrument PDR RFA Summary and official IIRT status, 2/02/2006
Review
Date
Total #
of RFAs
Responses
submitted by
Instrument
RFA's Closed by
Review Team
Open
LOLA PDR &
ΔPDR
6/16/2005
10/6/2005
59
52
42
17
LROC PDR
9/8/2005
18
13
10
8
LAMP PDR
9/12/2005
14
11
8
6
Diviner PDR &
Thermal ΔPDR
9/14/2005
11/15/2005
47
28
23
24
LEND PDR
9/30/2005
3
0
0
3
CRaTER PDR
9/29/2005
5
2
0
5
Payload Systems Mgmt Overview 2006-02-08
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50
Wrapup
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
51
Payload Summary
• Taken as a group, LRO Payload is development is going
(atypically) smoothly to date :
– Instruments have been able to preserve heritage with minimal
obsolescence
– Team leads have stayed largely intact since selection
– Documentation maturity is appropriate for this stage
– Instrument teams employing rigorous risk management philosophy
– Grass-roots costing exercises have credibility
– No Instrument currently in danger of missing 10/15/2007 delivery
date
– All instruments currently on track for CDRs in the 4-6/2006
timeframe
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52
Backup charts
Payload Systems Mgmt Overview 2006-02-08
Arlin Bartels – Payload Systems
53
Cosmic Ray Telescope for the Effects of
Radiation (CRaTER)
PI: Harlan Spence, BU
IM: Rick Foster, MIT
ISE: Bob Goeke, MIT
CRaTER Rapid Prototype Model
Payload Systems Mgmt Overview 2006-02-08
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54
CRaTER Development Summary
•
•
Significant design changes since PDR: none
Current Status
–
Placed purchase order for EM semiconductor detectors
•
–
–
•
Science team has successfully conducted tests of prototype detectors at
LBNL 88” cyclotron
Successful test with Tissue Equivalent Plastic (TEP) test apparatus at
Massachusetts General Hospital’s proton beam facility (few tens of MeV up
to 230 MeV)
Significant near-term work
–
–
–
•
•
140 u detectors due 4/2006, 1000 u detectors due 6/2006
Initial heavy-ion run at Brookhaven National Lab in 3/2006
Fabrication/test of Engineering Model Instrument
Design/fab of GSE
Significant Open Issues: None
Target CDR Date: June 2006
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Arlin Bartels – Payload Systems
55
Diviner
Lunar Radiometer Experiment (DLRE)
PI: David Paige, UCLA
IM: Wayne Hartford, JPL
ISE: Marc Foote, JPL
MCS Flight Instrument
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56
Diviner Development Summary
•
•
Significant design changes since d-PDR: none
Current Status
–
–
–
Budget agreement reached with LRO Project
Mission Assurance agreement with Project on path to closure
Thermal design making excellent progress
•
–
–
–
–
•
Oxford building first set of EM filters
FPGA logic designs on track
Flight software ready to begin unit test
EEE Parts 96% kitted
Significant near-term work
–
MCS Actuator Life Test completion
•
–
–
•
Currently at 2.9 M cycles, test scheduled to complete in mid-March at 3.5 million cycles
Thermal Development Test
Test EM LWIR (50-200 u) filters in full system-test
Significant Open Issues
–
–
•
Thermal shield no longer needed
Ongoing trade study will ensure physical location of DLRE on LRO meets thermal requirements.
Diviner considers proximity to Mini-RF a major concern without antenna patterns available from
Mini-RF
Target CDR date: 4/18-19/2006
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57
Lyman-Alpha Mapping Project (LAMP)
PI: Alan Stern, SwRI
IM: Ron Black, SwRI
ISE: Dave Slater, SwRI
New Horizons Alice Flight Instrument
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58
LAMP Development Summary
•
•
Significant design changes since PDR: none
Current Status
– Electronics Engineering Models (form, fit and function)
• Redundant LVPS undergoing final test
• HVPS module checkout underway
• Two each C&DH modules with updated LAMP flight software undergoing test and checkout
– Obtained good results from prototype testing of LTS electronics.
– Procurement – 19 EEE parts left to order. All flight PCBs (coupon tested) in house.
• EEE Parts spare parts being released by New Horizons to LAMP
– Flight mechanical fabrication well underway
– Hardware subcontracts in place and on track
•
Significant near-term work
–
–
–
–
–
•
•
Complete mechanical and thermal analysis.
Complete scattered light analysis based on spacecraft configuration.
Complete preliminary reliability analyses.
Delivery of Interface Simulator to GSFC.
LTS EM fabrication and characterization.
Significant Open Issues: none
Target CDR date: 5/2-3/2006
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59
Lunar Exploration Neutron Detector
(LEND)
PI: Igor Mitrofanov, IKI
IM: Anton Sanin, IKI
ISE: Maxim Litvak, IKI
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60
LEND Development Summary
•
•
Significant design changes since PDR: none
Current Status
–
–
–
–
–
–
•
Selected pure 10B for the neutron collimation module
Testing of LEND laboratory prototype is in progress
Selected exact detector types
Proceeding with FPGA logic design
95% of electronic parts are selected and suppliers have been chosen
LEND Engineering Team remains at GSFC the week after PDR to complete
ICDs and documentation
Significant near-term work:
– Perform laboratory unit tests with different types of collimation modules
– Interface tests between LEND/LRO electrical simulators
•
Significant Open Issues
– Need to bring closure to NASA-FSA IA document
•
Target CDR date: 6/2006
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61
Lunar Orbiter Laser Altimeter (LOLA)
PI: David Smith, GSFC
Deputy PI: Maria Zuber, MIT
IM: Glenn Jackson, GSFC
Deputy IM: Ron Zellar, GSFC
ISE: John Cavanaugh, GSFC
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62
LOLA Development Summary
•
Major Design Changes since PDR:
–
–
•
Main Electronics Box (MEB) now thermally isolated from Instrument Module (IM)
Multiple small(er) changes to incorporate high-precision Laser Ranging (LR) tracking system
Current Status
–
Engineering unit Time-to-Digital Converter (TDC) ASIC has undergone significant testing
•
–
–
–
EM energy detector has been tested with breadboard laser
Beryllium billet procurement imminent
Majority of LOLA optics procurements complete
•
•
•
•
EM version scheduled to arrive in March
All laser optics procurements complete
– Design complete for laser beam expander and Laser Electronics Assembly
Significant near-term work:
– Breadboard Digital Unit (DU) board-level test
• Verifying FPGA and TDC ASIC performance
– Complete STOP analysis with new Orbiter configuration
– EM laser systems test
– Power Control Assembly (PCA) test with breadboard laser
– Structural analysis with EELV loads
– FM laser diodes life tested
– Receiver optics thermal test
– MLA receiver optics dynamics tests
– EM analog board test
– Diffractive Optic Element (DOE) thermal test
Significant Open Issues: None
Target CDR date: June 2006
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63
Lunar Reconnaissance Orbiter Camera
(LROC)
PI: Mark Robinson, Northwestern University
IM: Scott Brylow, MSSS
ISE: Mike Caplinger, MSSS
Mars Color Imager
Mars Reconnaissance Orbiter
Context Camera
Mars Reconnaissance Orbiter
MSSS Spaceflight Imaging Systems
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64
LROC Development Summary
•
Design Changes since PDR:
– Now providing four 1A switched services instead of one big one
– External radiators added to NACs and WAC
– WAC Baffle redesigned to constrict straylight Field-of-Regard (S/C accommodation
issue)
•
Current Status
–
–
–
–
–
•
Significant near-term work
–
–
–
–
•
LROC team making excellent progress and on track for CDR
NAC CDR (with MSSS vendor) scheduled for late February
WAC CDR (with MSSS vendor) scheduled for early March
MRO parts (WAC electronics, NAC electronics) transferred from JPL to GSFC
Electronics breadboards or brassboards complete for all assemblies
Completion of NAC, WAC radiator design
Updating thermal models for NAC, WAC to reflect new S/C locations, radiators
Mechanical analysis of new WAC baffling
Mechanical analysis of NAC adapter plate and corresponding optical sensitivity
Significant Open Issues
– Completing thermal design in beta-gamma orbiter configuration
•
Target CDR date: 5/3/2006
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65
Mini-RF Technology Demonstration
PM: Bill Marinelli, NAWC
DPM: Dean Huebert, NAWC
PI: Chris Lichentenberg, NAWC
Sample Antenna Element Matrix
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66
Mini-RF Development Summary
•
•
Design changes since PDR: n/a
Current Status
– Draft PAIP submitted
– National Telecommunications and Information Administration (NTIA) forms required
by the U.S. Dept. of Commerce nearly complete
•
Significant near-term work:
– Firm up mechanical and thermal ICDs
– Refine mechanical and thermal math models
•
Significant Open Issues
– Interface temperatures for coupled electronics boxes are under negotiation with LRO
(ranges being made more benign)
– Antenna Placement Trade Study to be completed
• FOV is blocked during portions of each orbit by the LRO Solar Array
•
Target CDR date: September 2006 (TBD)
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Mini-RF MOA Status
• Draft Memorandum of Agreement between NASA HQ
Exploration Systems Mission Directorate (ESMD) and Space
Operations Mission Directorate (SOMD) is being developed
(January 2006)
• LRO Project is refining details of the Mini-RF to Spacecraft
allocations and will provide inputs to ESMD
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68
LRO PDR
Data Management Overview
Stan Scott
Karen North
February 8, 2006
Agenda
•
•
•
•
Data Management Role
LRO Data Working Group (LDWG)
Data Management Plan Status
Additional Documentation
Stan Scott - Data Management Overview
2
Data Management Role
• Serve as LRO Project POC for measurement data systems
and PDS
• Approve SOC deliverable documents
• Assist SOC managers with any peer reviews they may
convene
• Complete LRO Project Data Management Plan (DMP)
• Coordinate and approve SOC DMPs
• Coordinate SOC-PDS ICDs
• Provide Project perspective in measurement data format
definition
• Coordinate SOC-PDS interface testing
• Help coordinate SOC and any PDS testing with ground
system
Stan Scott - Data Management Overview
3
LRO Data Working Group (LDWG)- 1
• LRO is not like traditional planetary science missions due to
quick flight- SOC planning, development, and testing
activities must be completed before launch
• LDWG is biweekly Project telecon with SOCs & PDS
• Main topics involve SOC data processing & PDS archive
• Project, SOC, & PDS can keep each other up-to-date
• Issues discussed with solution coordination/tracking
• Development & testing schedules are coordinated
• Meeting notes distributed & action items tracked
Stan Scott - Data Management Overview
4
LRO Data Working Group (LDWG)- 2
• First meeting held on December 19, 2005
• Two additional meetings held before PDR
• Issues/topics discussed to date include:
– ITs wish to use monthly LOLA data for analysis
• DEM data
• Time-sequenced profile data
– IT requirements for SPICE data and source
• LOLA to provide NAIF with monthly reconstructed ephemerides
– Scope and approach for LRO Archive Plan
– Began discussion of SOC - PDS ICDs
• E-mail listserv and document sharing web site established
Stan Scott - Data Management Overview
5
LRO Data Working Group (LDWG)- 3
LDWG Participants:
ITs:
CRaTER
Diviner
LAMP
LEND
LOLA
LROC
Mini-RF
Larry Kepko /BU
Tim Schofield /JPL
Joel Parker /SwRI
Ken Ramey,
Kate Crombi,
Karl Harshman /UA
Greg Neumann /GSFC
Ernest BowmanCisneros,
Joe Digilio /NU;
Eric Eliason /UA
Howard Taylor,
Helene Winters /JHU
PDS:
LRO Lead
Keith Bennett /WU
Engineering Node Betty Sword /JPL
Geosciences Node Diviner/LEND/LOLA:
Susie Slavney,
Jen Ward/WU
Imaging Node
LAMP:
Rafael Alanis /JPL
LROC:
Chris Isbell /USGS
NAIF Node
SPICE:
Chuck Acton /JPL
Planetary Plasma CRaTER:
Interactions (PPI) Steve Joy /UCLA
LRO Project:
Project: Stan Scott, Karen North, Arlin Bartels
Flight Dynamics: Mark Beckman
Science: John Keller
Stan Scott - Data Management Overview
6
LRO Data Working Group (LDWG)- 4
PDS Nodes/Subnodes/Data Nodes
DIVINER
New Mexico
State
University
Cornell
University
ATMOSPHERES
Jet Propulsion
Laboratory
Stanford
University
University of
Arizona
ENGINEERING
GEOSCIENCES
Washington
University
LEND
LOLA
Arizona State
University
RADIO
SCIENCE
Mini-RF
University of
California at
Los Angeles
Navigation and
Ancillary
Information
Facility (NAIF)
PLANETARY
PLASMA
INTERACTIONS
Southwest
University
(San Antonio, TX)
CRaTER
University of
Iowa
Jet Propulsion
Laboratory
University of
Hawaii
University of
Maryland
SMALL BODIES
Planetary Science
Institute
IMAGING
RINGS
SETI Institute
United States
Geological Survey
(Flagstaff, AZ)
Arizona State
University
LAMP
LROC
University of
Arizona
GSFC/MOC
FDF
Data Management Plan Status- 1
• The LRO Data Management Plan (DMP) addresses:
– Roles and responsibilities for measurement data
processing, archive, and distribution
– MOC-SOC & SOC-PDS data flows
– Data handling steps including processing, reprocessing,
and validation
– Summary descriptions of measurement data products
• DMP schedule
– Draft version available
– Final version due for Mission CDR
Stan Scott - Data Management Overview
8
Data Management Plan Status- 2
•
•
The SOC Data Management and Archive Plan (DM&AP) includes:
– Detailed description of measurement data products
– Team data management functions and responsibilities
– SOC configuration management process
– SOC data processing environment
– Algorithm peer review process
– Data release policy
– Archive Plan
• Archive generation
• Validation and peer review
• Data transfer and distribution
SOC DM&APs schedules
– Draft version due date: Mission CDR - 30 days
– Final version due date: Mission CDR + 90 days
Stan Scott - Data Management Overview
9
Additional Documentation
Other measurement data-related documentation includes:
• ICDs
– MOC-SOC ICD (LRO Project document)
– SOC-PDS Discipline Node ICDs (jointly developed)
– LOLA SOC-LRO SOCs ICD (LRO Project coordination)
• SOC documents for PDS
–
–
–
–
Data Product Software Interface Specifications (SIS)
Archive Volume SIS
High-level data set descriptions
Several other small deliverables
• Small LRO Project deliveries to PDS
– High-level mission description
– High-level spacecraft description
Stan Scott - Data Management Overview
10
Back-up Chart
PDS Node Descriptions
•
•
•
•
•
•
•
•
The Engineering Node -- provides systems engineering support to the entire PDS,
handling global aspects such as standards (data, software, documentation, operating
procedures), technology investigations, coordination and development of system-wide
software, coordination of data ordering and distribution, catalog development and
implementation, and maintenance of the PDS catalogs.
Atmospheres Node -- Responsible for the acquisition, preservation, and distribution of all
non-imaging atmospheric data from all planetary missions.
Geosciences Node -- Maintains data sets that are relevant to the geosciences discipline, the
study of the surfaces and interiors of terrestrial planetary bodies.
Planetary Imaging Node -- Maintains and distributes the archives of planetary image data
acquired from NASA's flight projects with the primary goal of enabling the science
community to perform image processing and analysis on the data.
Planetary Plasma Interactions Node -- Responsible for acquisition, preservation, and
distribution of fields and particle data from all planetary missions.
Rings Node -- Devoted to archiving and distributing scientific data sets relevant to
planetary ring systems.
Small Bodies Node -- Provides data sets and consulting expertise for comets, asteroids, and
interplanetary dust.
Navigation and Ancillary Information Facility (NAIF) Node -- Responsible for design
and implementation of the SPICE concept for archiving, distributing and accessing
observation geometry and related ancillary data used in mission design, mission evaluation,
observation planning and science data analysis.
Stan Scott - Data Management Overview
12
LRO PDR
Ground System
Rick Saylor
February 7, 2006
LRO Ground System Overview
•
Ground System (GS) consist of five main elements
–
–
–
–
–
•
Space Communications Network (SCN)
Mission Operations Center (MOC)
Flight Dynamics Facility (FDF)
NASA Integrated Network Services (NISN)
Mission Operations Team (MOT)
Functions:
– Provide Tracking, Telemetry and Command (TT&C) services
– Archive raw mission data for life of mission
– Receive, track, and deliver measurement files to the SOCs within 24-hours of
ground receipt
– Perform daily operations and monitor health and safety of the orbiter
– Provide network interfaces between ground system elements and external
elements
– Perform orbit determination, maneuver planning, and attitude determination
support. Generate flight dynamics products for operations and data
processing
Rick Saylor - Ground System
2
Ground System Architecture
Rick Saylor - Ground System
3
Space Communications Network
Rick Saylor - Ground System
4
Mission Operations Center
Rick Saylor - Ground System
5
Flight Dynamics
Rick Saylor - Ground System
6
NASA Integrated Services Network
Rick Saylor - Ground System
7
Mission Operations Team
Rick Saylor - Ground System
8
Ground System Analysis/Trade Studies
Trade Study/Analysis
Description
Outcome
Ground Network Support for LRO
Select ground network support for LRO for all
mission phases.
White Sands 18m Ka-band
station and USN sites for SBand support
Telemetry and Command System
Select Telemetry & Command system for orbiter I&T
and mission operations.
Integrated Test & Operations
System (ITOS)
Trending and Analysis System
Select trending and analysis system for mission
operations
Integrated Trending and
Plotting System
Mission Planning System
Select mission planning system for mission
operations.
FlexPlan
S-Band Tracking Data Improvement
Evaluated plan to support high accuracy tracking
requirement
S-Band tracking can’t support
high accuracy requirement
Launch & early mission coverage
Evaluate potential gaps in coverage of critical
activities shortly after launch and investigate options
Baseline SN coverage after
separation
Backup MOC Location
Select location that allows operations team to
maintain orbiter health and safety if MOC is offline
Started – Complete by GS
Peer PDR
Data Storage System
Select hardware implementation for data storage
within the mission operations center
To Be Started – Complete by
GS SDR
Rick Saylor - Ground System
9
Ground System Documentation Status/Plan
Document Title
Document Number
Status
LRO Ground System Development Product Plan
431-PLAN-000046
Released
LRO Detailed Mission Requirements Document
431-RQMT-000048
Released
LRO Telemetry and Command Formats Handbook
431-HDBK-000052
In Draft Review
LRO Ground System Interface Control Document
431-ICD-000049
Started – GS PDR
LRO CFDP Implementation Specification
431-SPEC-000078
Started – GS PDR
LRO Mission Flight Rules and Constraints
431-OPS-000309
Started - MCDR
LRO Ground System Mission Readiness Test Plan
431-PLAN-000079
Started – GS SDR
LRO Mission Operations Test Plan
431-PLAN-000308
Started – GS SDR
LRO Ground System Software Acceptance Test Plan
431-PLAN-000424
Not Started – GS SDR
LRO Flight Operations Plan
431-PLAN-000064
Not Started – GS SDR
LRO Telemetry and Command Database Management Plan
431-PLAN-000051
Not Started – GS SDR
LRO NISN Operations Support Plan
431-PLAN-000302
Not Started – GS SDR
LRO Launch and Commissioning Handbook
431-HDBK-000069
Not Started – MOR
LRO Operations Training Plan
431-PLAN-000080
Not Started – MOR
LRO Mission Operations Contingency Plan
431-PLAN-000068
Not Started - FORR
Rick Saylor - Ground System
10
Ground System Review Plan
Review
Status
RFAs
RFAs Close Date
Ground System Requirements Peer Review
Complete – Jan 11, 2006
33
March 20, 2006
White Sands/USN Peer Review #1
Complete – Jan 19, 2006
20
March 3, 2006
White Sands/USN Peer Review #2
March 2006
Ground System Peer Preliminary Design Review
June 2006
Ground System Single Design Review
November 2006
Mission Operations Review
June 2007
Flight Operations Readiness Review
June 2008
Rick Saylor - Ground System
11
Wrap-Up
•
•
Ground system development team is following the flight segment design
All major components of the ground system has been selected, largely an
integration and test effort
– SCN consist of current ground stations except White Sands S/Ka ground station.
• Multi-mission WS1 development is on schedule and meets LRO requirements
– MOC systems are based on existing software packages
• Requires only minor enhancements for LRO
•
Initiated TIMs with science operations centers
– Captured initial interfaces and reviewed mission data products
•
Successfully completed requirements peer review
– Included review of all level 3 requirements (~600 requirements)
– Received 33 RFAs, all considered minor
•
Successfully completed SRR/PDR on WS1/USN
– Received 20 RFAs, plan to revolve by March 3, 2006
•
Started operations team staffing
– Providing support to subsystem integration and test
– Developing detailed operations plans and concepts
•
Initiated development of the ground system ICD
– Defines ground system interfaces and products for GS elements and the instrument
SOCs
– Initial release is planned before GS PDR
Rick Saylor - Ground System
12
LRO PDR
Launch Vehicle
Tom Jones
February 8, 2006
Agenda
•
•
•
•
•
•
LRO Launch Vehicle Mission Requirements
Launch Services NLS-B
Launch Services NLS-L
Launch Services NLS-O
NLS Launch Services Acquisition Process
LRO EELV Interfaces
Tom Jones - Launch Vehicle
2
Launch Vehicle Mission Requirements
Level 1
Level 2
Level 3
Requirement
MRD-24
431-RQMT000397, Rev A
Vehicle Performance
Capability equal to or greater
than 2000kg. C3 > -1.8
km2/sec2
MRD-25
431-RQMT000397, Rev A
Insertion Accuracy
The maximum three sigma
velocity dispersion at
injection is 3 m/s.
RLEP-LROP50
MRD-26
N/A
DE-Spin
EELV alleviates requirement
RLEP-LROP50
MRD-27
431-RQMT000397, Rev A
Pointing Accuracy/Body
Rates
At spacecraft separation the
L/V pointing error shall be
less than 5 deg per axis and
the L/V induced body rates
after separation shall be less
than 1º, 2º, and 2º per second
about the spacecraft x, y, & z
axes, respectively.
RLEP-LROP50
MRD-28
431-RQMT000397, Rev A
Vehicle Interfaces
Defined in 431-RQMT000397, Rev A
RLEP-LROP40
RLEP-LROP50
Tom Jones - Launch Vehicle
3
Launch Services NLS-B
Delta IV
Delta II
2320-10
2326-9.5
2420-10
Delta III
2920-10L 2925H-9.5
3940-11
Tom Jones - Launch Vehicle
4040
4240
4450
4050H
4
Launch Services NLS-B
•
•
•
Mission Requirement: C3 = -1.8 km2/sec2
Launch Vehicle Performance: Delta-IV 4040-12 @ 2855 kg
Ground Rules
– 3-sigma probability of Stage 2 commanded shutdown, plus additional
reserves as determined by the LSP.
– 1194-4 payload adapter.
– Launch from SLC-37 at CCAFS (Cape Canaveral Air Force Station).
– 185 km (100 nmi) circular park orbit at 28.9 deg. inclination.
– 185 km (100 nmi) escape orbit perigee.
– Performance shown is applicable for declinations between 28.9 deg. and 28.9 deg.
Tom Jones - Launch Vehicle
5
Launch Services NLS-L
Atlas V
Atlas III
Numbering Scheme:
1st digit: 4 = 4m Fairing
5 = 5m Fairing
2nd digit: # of Strap-on
Solid Rocket
Boosters
3rd digit: 1 = (SEC) Single
Engine Centaur
2 = (DEC) Dual
Engine Centaur
AIII-B
40X
50X
51X
52X
Tom Jones - Launch Vehicle
53X
54X
55X
6
Launch Services NLS-L
•
•
•
Mission Requirement: C3 = -1.8 km2/sec2
Launch Vehicle Performance: Atlas V 401 @ 3565 kg
Ground Rules
– 3-sigma mission required margin, plus additional reserves as determined by
the LSP.
– Launch from SLC-41 at CCAFS (Cape Canaveral Air Force Station).
– Performance values assume harness, logo, reradiating antenna, 3 payload
fairing doors.
– Payload mass greater than 9000 kg (19,841 lbs) may require mission unique
accommodations.
– Type B2 payload adapter.
– 4-meter Extended Payload Fairing (EPF).
– 185 km (100 nmi) minimum park orbit perigee altitude.
– 185 km (100 nmi) minimum escape orbit perigee altitude.
– Performance shown is applicable for declinations between 28.5 deg. and 28.5 deg.
Tom Jones - Launch Vehicle
7
Launch Services NLS-L
•
•
•
Mission Requirement: C3 = -1.8 km2/sec2
Launch Vehicle Performance: Atlas V 501 @ 2785 kg
Ground Rules
– 3-sigma mission required margin, plus additional reserves as determined by
the LSP.
– Launch from SLC-41 at CCAFS (Cape Canaveral Air Force Station).
– Performance values assume harness, logo, reradiating antenna, 3 payload
fairing doors.
– Payload mass greater than 9000 kg (19,841 lbs) may require mission unique
accommodations.
– Type B2 payload adapter plus type C2 spacer.
– 5-meter Short Payload Fairing.
– 185 km (100 nmi) minimum park orbit perigee altitude.
– 185 km (100 nmi) minimum escape orbit perigee altitude.
– Performance shown is applicable for declinations between 28.5 deg. and 28.5 deg.
Tom Jones - Launch Vehicle
8
Launch Services NLS-O
•
•
•
Mission Requirement: C3 = -1.8 km2/sec2
Launch Vehicle Performance: No advertised capability.
Ground Rules
– Recently added Orbital Sciences Corp. as a provider to the NLS contract.
Tom Jones - Launch Vehicle
9
Launch Services NLS Acquisition Process
•
Nominal competition under NLS.
Considerations:
– Estimated 6 month procurement process if single procurement
• Spacecraft project completes generic IRD and submits to KSC LSP MIT
• LSP, with spacecraft project, develops the Request for Launch Service Proposal
(RLSP) and submits to Boeing (NLS-B), Lockheed Martin (NLS-L), and Orbital
Sciences (NLS-O)
• Vendors may decline to offer a solution or submit a launch service proposal
within 30 days of the RLSP.
• Proposals are evaluated by LSP with spacecraft customer and recommendation
presented to Flight Planning Board for authorization.
• Launch Service Task Order (LSTO) is awarded to the contractor that provides the
best value in launch services to meet the Government’s requirements based on
technical capability/risk, reasonableness of proposed price, and past performance.
– Volume buy possibility if other missions are also scheduled to issue an LSTO
within the same timeframe.
• This presents opportunity for a quantity discount.
Tom Jones - Launch Vehicle
10
Launch Services NLS Acquisition Process
•
Schedule
– Release Request for Launch Service Proposal (RLSP) January 2006.
– Proposal evaluation March-May 2006.
– Award Launch Service Task Order (LSTO) by end of June 2006 (~L-28
Months)
•
Spacecraft Inputs
– Initial Interface Requirements Document (IRD) submitted 12/16/2005 to KSC
– Final input submitted to KSC January 01/11/2006 in support of RFP release
by the end of January 2006.
Tom Jones - Launch Vehicle
11
Launch Services Spacecraft Interfaces
•
Spacecraft requires a EELV class vehicle
– Separated spacecraft mass of 2000kg with a minimum energy (C3) shortcoast transfer for that day.
– Based on a Periselene altitude above lunar south pole of 257 km ± 5 km, True
Lunar inclination of 90 deg ± 0.2 deg and Lunar beta angle at lunar solstice of
0 deg ± 20 deg.
•
Spacecraft Requires a 1194mm clampband separation interface
– Two 61 pin umbilical connectors.
– PLF and LV surfaces to which the SC is exposed shall be cleaned and
inspected, and maintained to Level 450A per IEST-STC-CC1246D. All
interface surfaces including the EELV Launch Vehicle Adapter shall meet
this requirement prior to final mating.
•
Fairing
– At least a 4m fairing is required.
– With 3 doors locations TBD.
•
•
T-0 GN2 purge Grade C is required
After contract award the IRD (431-RQMT-000397) will become the basis
for the EELV Interface Control Document which will supersede the IRD.
Tom Jones - Launch Vehicle
12
LRO PDR
Mechanical Overview
Giulio Rosanova
February 8, 2006
LRO Mechanical Team Personnel List
Area of Responsibility
Mechanical System Lead Engineer
S/C Structure Lead Engineer
HGAS Lead Engineer
Mechanical GSE Lead Engineer
Gimbals Lead Engineer
Structural Analysis Lead
Loads/Dynamics Analyst
S/C Structural Analyst
Shop Planner
Test Engineer
Lead Technician
Senior Technician
Technician
SAS Lead Engineer
Lead Designer (PM)
Senior Designer (DPLY)
Designer (Misc.)
Designer (AM)
Designer (IM)
IM Structural Analyst
DPLY Structural Analyst
Structural Analysis
Technician
Technician (Detailed to PROP Group)
Name
Code/Org
Giulio Rosanova
543
Gordon Casto
543
Greg Martins
543
Darian Robbins
543
Kamal Thakore
544
Greg Clarke
542
Craig Stevens
542
Wayne Chen
542
TBD
547
TBD
549
Steve Patton
543
Karl Schuler
543
TBD
547
Mike Hersh
Swales
Dan Hayward
ManTech
Suk Yoon
Swales
Joe Green
Swales
Robert Robinson ManTech
Ginger Bronke
Swales
Heather Borowski Swales
Bryan Rizzo
Swales
Shelly Conkey
Swales
George Mooney
Swales
Hal Baesch
Swales
Giulio Rosanova - Mechanical Overview
3
LRO Mechanical PDR Agenda
•
Introduction (G. Rosanova)
– Requirements
– Design Overview
– Nomenclature
– Dimensional Layouts
– Launch Vehicle Interfaces
– ICDs
•
•
•
•
PM, AM, IM - Primary Structure (G.
Casto)
HGAS Mechanical (G. Martins)
SAS Mechanical (M. Hersh)
Analysis Mechanical (G. Clarke)
•
Programmatics (G. Rosanova)
•
MGSE (D. Robins) Æ I&T
Giulio Rosanova - Mechanical Overview
4
Lunar Reconnaissance Orbiter (LRO)
Mechanical System Level 2 Flow Down Significant Requirements
STRUCTURES
Level 2 Req.
Verification
Level 3: Requirements
Respon.
I
Identifier
A
D
T
Requirement
MRD-64
Verification
Testing
MRD-30 & 65
Mechanical
Environments
MSRD1, 2, 3
Primary Structure
The Mechanical System Team shall Design, Analyze, Build, and Test a Primary Structure that
includes a Propulsion Module (PM), Avionics Module (AM), and an Instrument Module (IM).
3
3
540
MRD-22
Extended
Mission MRD96 Thruster
Locations
S/C
Component
MICD’s
Instrument
MICD’s
MSRD5, 6 , 7, 8, 9,
10, 11, 12
Accommodations
The PM, AM, & IM shall accommodate a Hydrazine Propulsion System, including 897.5kg of
consumables (for 4 years), Fuel Tank, Pressure Tank, Thrusters, Fuel Lines, Regulators, and other
miscellaneous Prop components as well as the ACS Reaction Wheel Assemblies (RWA’s), LRO
Spacecraft Subsystem Avionics Components, the Diviner Instrument, the LAMP Instrument, Mini-RF
Tech Demo, CRaTER, LEND, LOLA, and LROC Instruments as well as the Star Trackers and IRU
per their MICD’s.
3
3
540
MRD-14
Nadir Pointing
MRD-62
Coordinate
Systems
MRD-71
Fields of View
MSRD13, 14
Instrument Placement
The Instruments shall be placed on the PM, AM & IM Structure such that they point in the Nadir
(+Z), (CRaTER +/-Z) direction and such that all Optical & Thermal fields of Regard are clear of any
obstruction.
3
3
540
MRD-49
Pointing
Allocations
MRD-70
Structural
Stability
MRD-72
Flexible Modes
MSRD15, 16, 17,
18, 19, 20
Structure Stability
The Mechanical System structure shall be sufficiently stiff and stable (thermal distortion) to maintain the
pointing and alignment requirements of the Spacecraft Components and Instruments, High Gain Antenna
System (HGAS), and the Solar Array System (SAS per the LRO Pointing and Alignment Specification 431SPEC-000113.
Giulio Rosanova - Mechanical Overview
4
3
3
540
5
Lunar Reconnaissance Orbiter (LRO)
Mechanical System Level 2 Flow Down Significant Requirements
STRUCTURES
Level 2 Req.
Verification
Level 3: Requirements
Respon.
I
Identifier
A
D
T
Requirement
MRD-29 Mass
Allocations
MRD-61
Margins/
Reserves
MSRD-21
Mass
The Mechanical System shall not exceed a mass allocation as stated in the LRO Technical Resource
Allocation Specification 431-SPEC-000112, and shall maintain adequate margin throughout system
development.
3
2
540
MRD-17
Momentum
Management
MRD-75 Mass
Properties
MSRD22, 23, 24, 25
Mass Properties
The Mechanical System shall manage the Mass Properties for the entire LRO Orbiter, such that CG migration,
CGCP-First Area Moments, and Products of Inertia are minimized to meet the Momentum management
requirement.
3
2
540
MRD-67
Thermal
Environments
MSRD-26
Thermal Radiators
The Mechanical System shall incorporate Thermal radiator panels to meet component thermal Requirements.
3
540
MRD-73
Accessibility
MSRD27, 28
Access
The Mechanical System shall incorporate removable structural panels, to facilitate access to internal
components. Ground Support Equipment of external components shall be accommodated.
MRD-66
Electrical
Environments
MSRD29, 30
Chassis Ground
The Structure shall serve as the Electrical Chassis Ground.
MRD-76
Support
Equipment
MSRD31, 32, 33
MGSE
The Mechanical System shall provide the LRO MGSE.
Includes Transportation Dollies, Lifting Fixtures, G-Negation Systems, and Protective Covers.
Giulio Rosanova - Mechanical Overview
4
3
3
4
540
2
540
4
540
6
Lunar Reconnaissance Orbiter (LRO)
Mechanical System Level 2 Flow Down Significant Requirements
LAUNCH VEHICLE
Level 2 Req.
Verification
Level 3: Requirements
Respon.
Identifier
I
A
2
3
D
T
Requirement
540
MRD-2 Launch
Vehicle
MRD-28
Vehicle
Interfaces
MSRD34, 35, 36,
37, 38
L/V Interface
The Mechanical System shall fit in the EELV Launch Vehicle Fairing Static Payload Envelope. The
Mechanical System shall be compatible with EELV Launch Vehicle Payload Attach Fitting (PAF).
MRD-1 Launch
Mass
MRD-30 & 65
Mechanical
Environments
MSRD-39
L/V Environments
The Mechanical System shall sustain the EELV Loads Environments at the LRO Maximum Launch Mass of
2000kg.
3
3
540
MRD-2 Launch
Vehicle
MSRD-40
Fundamental Modes
The Mechanical System shall be compatible with the EELV fundamental frequency requirements in the
Launch Configuration.
2
3
540
MRD-28
Vehicle
Interfaces
MRD-69
Contamination
Control
MSRD41, 42
Purge System
The Mechanical System shall Incorporate Instrument Dry Purge Lines on Orbiter up to T-0 (umbilical breakaway).
3
540
Giulio Rosanova - Mechanical Overview
4
4
7
Lunar Reconnaissance Orbiter (LRO)
Mechanical System Level 2 Flow Down Significant Requirements
DEPLOYABLES
Level 2 Req.
Verification
Level 3: Requirements
Respon.
I
Identifier
A
D
T
Requirement
MRD-7
Deployables
MRD-64
Verification
Testing
MRD-30 & 65
Mechanical
Environments
HGA MICD
Solar Array
MICD
MSRD43, 44, 45,
46, 47, 48
Deployable Mechanisms
The Mechanical System Team shall Design, Analyze, Build, and Test a High Gain Antenna Deployment and
Articulation System (HGADAS) that accommodates the S & Ka Band High Gain Antenna System, and a
Solar Array Deployment and Articulation System (SADAS) that accommodates the Solar Array System. The
HGADAS & SADAS shall consist of a Deployment Hinge, Release Mechanism, Boom, 2 Axis Gimbal Set,
and 2 Axis Gimbal Controller.
MRD-15 Solar
Array Tracking
MRD-16
Antenna
Tracking
MRD-62
Coordinate
Systems
MRD-71 Fields
of View
MSRD49, 50
Deployables Placement
The HGADAS shall be Deployed parallel to the –Z axis of LRO, and shall Articulate the High Gain Antenna
through a complete Hemisphere centered on the –Z axis, such that RF fields of Regard are clear of any
obstruction. The SADAS shall be Deployed parallel to the –Y axis of LRO, and shall Articulate through 0 to
+90 degrees about the Z axis and +/-90 degrees about the Y axis.
MRD-49
Pointing
Allocations
MRD-70
Structural
Stability
MSRD51, 52
Deployables Stability/Pointing
The HGADAS shall be sufficiently stiff and stable to maintain the pointing and alignment requirements of the
HGA as stated in the LRO Pointing and Alignment Specification 431-SPEC-000113. The SADAS shall be
sufficiently stiff and stable to maintain the pointing requirements of the Solar Array as stated in the LRO
Pointing and Alignment Specification 431-SPEC-000113.
Giulio Rosanova - Mechanical Overview
4
3
3
540
3
3
540
3
3
540
8
Lunar Reconnaissance Orbiter (LRO)
Mechanical System Level 2 Flow Down Significant Requirements
DEPLOYABLES
Level 2 Req.
Verification
Level 3: Requirements
Respon.
I
Identifier
A
D
T
Requirement
MRD-67
Thermal
Environments
MSRD-53
Deployment Temp.
The HGADAS & SADAS shall be capable of Deploying at extreme temperature as Specified in
the LRO Thermal Systems Specification (431-SPEC-000091).
3
3
540
MRD-72
Flexible
Modes
MSRD-54
Deployed Frequency
The Deployed fundamental frequency of the HGADAS & SADAS shall be greater than 0.5Hz.,
such that they do not couple into the ACS control system.
3
3
540
MRD-21
Mission
Duration
MRD-51
Continuous
Operations
MSRD-55
Gimbal Operation
The HGADAS & SADAS Gimbals shall operate continuously for 14 Months without failure. Life
test shall be conducted for verification.
4
540
Giulio Rosanova - Mechanical Overview
9
LRO Deployed
3 PANEL MODULAR
SOLAR ARRAY
HGA
INSTRUMENT
MODULE
LAMP
X
Mini-RF
PROPULSION
MODULE
LEND
Y
Giulio Rosanova - Mechanical Overview
Z
10
LRO Deployed
CRaTER
LOLA
AVIONICS
MODULE
INSTRUMENT MODULE
(OPTICAL BENCH)
HGA
LEND
LEND
X
Mini-RF
Y
3 PANEL MODULAR
SOLAR ARRAY
Z
Giulio Rosanova - Mechanical Overview
11
LRO Stowed
CRaTER
LOLA
LROC
LEND
PROPULSION
MODULE
THRUSTERS
Mini-RF
Giulio Rosanova - Mechanical Overview
12
LRO Stowed
STAR
TRACKERS
3 PANEL MODULAR
SOLAR ARRAY
THERMAL
RADIATOR
PROPULSION
MODULE
Giulio Rosanova - Mechanical Overview
13
LRO Nomenclature
•
PROPULSION MODULE (PM)
–
Structure (Al)
•
•
–
–
–
–
–
–
•
•
Stub Skirt, Deck, Bulkheads,
Panels
Radiators
Fuel Tanks (2)
Pressurant Tank
Thrusters - 20# (4), 5# (8)
Prop. Components / Plumbing
Lines
S-Band Omni Antennas (2)
CSSs (6)
AVIONICS MODULE (AM)
–
Structure (Al)
•
•
–
–
–
–
INSTRUMENT MODULE (IM)
– Composite Optical Bench
(OB)
– Flexures (3)
– Star Trackers (2)
– Instruments (On IM)
• LAMP, LOLA, LROC
– Instruments (on PM)
• CRaTER, Diviner,
LEND
– Tech-Demo (on AM)
• Mini-RF
Frame, Panel w/Heat Pipes
Radiation Cover
Radiators (VCHP)
S/C Avionics Boxes (13)
Reaction Wheel Assemblies (4)
S/C Main Harness
NOTE: “LRO SPACECRAFT” Refers to LRO Orbiter Without Instruments
Giulio Rosanova - Mechanical Overview
14
LRO Nomenclature
•
HIGH GAIN ANTENNA SYSTEM (HGAS)
– HGA Deployment & Articulation Sys.-(HGADAS)
• HGA Deployment System (HGADS)
– Hinges / Damper / Potentiometer
– Release Mechanism / Pyro Actuators
• Articulation System (AS)
– Y-X Gimbal Set
– Cable Wrap
– Gimbal Controller
– HGA Boom
–
–
–
High Gain Antenna Dish, S-Band Patch Antenna
Wave-Guide, Coax Cable, Rotary Joints
Laser Ranging
•
SOLAR ARRAY SYSTEM (SAS)
– Solar Array Deployment & Articulation Sys.-(SADAS)
• Solar Array Deployment System -(SADS)
– Hinges / Damper / Potentiometer
– Release Mechanism / Pyro Actuators
• Articulation System (AS)
– Az-El Gimbal Set
– Cable wrap
– Gimbal Controller
– Solar Array (S/A) Substrate
– Module Substrates
– Frame
–
–
Solar Cells
CSSs (4)
NOTE: “LRO SPACECRAFT” Refers to LRO Orbiter Without Instruments
Giulio Rosanova - Mechanical Overview
15
MGSE List
•
ORBITER / Propulsion Module
– ORB/PM Dolly/Turnover Fixture
– ORB/PM Lift Sling
– GSE Stub Skirt
•
Avionics Module
– AM Dolly/Turnover Fixture
– AM Lift Sling
•
Instrument Module
– IM Dolly/Turnover Fixture
– IM Lift Sling
•
Solar Array System
– SAS Dolly/Turnover Fixture/SC-Simulator
– SAS Lift Sling
– SAS G-Negation System
•
High Gain Antenna System
– HGAS Dolly/Turnover Fixture/SC-Simulator
– HGAS Lift Sling
– HGAS G-Negation System
•
Transporter System
•
Wiring Harness Mock-Up
Giulio Rosanova - Mechanical Overview
16
Dimensional Layouts (Stowed)
Giulio Rosanova - Mechanical Overview
17
Dimensional Layouts (Deployed)
Giulio Rosanova - Mechanical Overview
18
Launch Vehicle Interface
Combined EELV
Payload Envelope
T-0 Purge
System
Battery A/C
Fairing Door
Fill & Drain
Fairing Door
Giulio Rosanova - Mechanical Overview
TCP
Fairing Door
19
Mechanical / Instrument ICD’s
All MICD/MIDs and CAD Models will be signed by March 15, 2006
Giulio Rosanova - Mechanical Overview
20
Mechanical / Avionics Component ICD’s
All MICD/MIDs and CAD Models will be signed by March 15, 2006
Giulio Rosanova - Mechanical Overview
21
LRO PDR
Spacecraft Structure
Gordon Casto
February 8, 2006
AM/PM Design
•
•
•
•
•
All aluminum design, machined fuel tank decks with honeycomb vertical panels
OD of stub skirt based on PAF dimension, Vertical panel centerlines coincident with stub
skirt centerline
Major panels are 1.5” (3.1pcf aluminum honeycomb, .032” facesheets)
AM consists of –Y panel of structure
– Avionics on –Y side of AM panel
– Reaction Wheels on +Y side of AM panel
– Orient components on AM to satisfy:
• Connector side(s) of box facing direction of major harness path
• 9” clear area on connector side of box
• 1” clearance with PM tanks is desirable
– Provide for electrical conductivity/grounding
– E-box thermal conductivity. Heat conducts through cotherm/nu-sil coupling into
embedded CCHP’s in panel
– RWAs aligned to external (-y) reference at AM level of integration
Minimize I-xz then I-yz then I-xy in deployed configuration with mission fuel load (post
lunar capture)
Gordon Casto - Spacecraft Structure
2
Primary Structure
Gordon Casto - Spacecraft Structure
3
PM Primary Structure Element
Gordon Casto - Spacecraft Structure
4
Major Interfaces
• HGADS Interfaces
– Base hinge & release
points to vertical panel
• Solar Array Interfaces
– Gimbal bracket to main
deck
– S/A release point are on
centerline of prop modual
panels
Gordon Casto - Spacecraft Structure
5
Details/Major Interfaces
• Lower Deck Interface to Panels
Panel Closeout Beyond
Honeycomb Panel
Panel Fitting
Lower Deck
Gordon Casto - Spacecraft Structure
6
Details/Major Interfaces
• Stub Skirt to Lower Deck Interface
Lower Deck
Prop Cone/Close-out
48 Bolt Interface
Stub Skirt
Gordon Casto - Spacecraft Structure
7
AM Primary Structure Element
Gordon Casto - Spacecraft Structure
8
IM Design
•
Composite “Wine Box” Design
– Reduces mass (fewer fasteners), increases stiffness with continuous
bond lines
– Uniform CTE for Entire Structure
•
•
•
– Graphite material is M55/CyanateEster Q.I. sheets, (~0CTE) large
database
– Insert material is Ti with BR127 electrical conductive primer
– Simplifies Overall Thermal Stability Analysis
Straight Forward Assembly Concept. Tooling used to match bond inserts
to ICD requirements
Simplified procurement of graphite material (buy flat laminate .xx thick)
Can fabricate in-house or procure as build to print.
Gordon Casto - Spacecraft Structure
9
IM Primary Structure Element
Gordon Casto - Spacecraft Structure
10
Major Interfaces
• IM to PM Interface
– Flexures mount to panels
with stiffener behind panel at
flexure locations
Gordon Casto - Spacecraft Structure
11
AM Grounding / Faraday Shield
Chotherm or
other thermal
coupler
Component Ground strap
Copper Ground Strap
if required
Box Chassis
Aluminum skin to
primary ground
AM Panel No Heat Pipes Shown
Gordon Casto - Spacecraft Structure
12
IM Grounding / Faraday Shield
Component Ground strap
Thermal Blanket
Copper Ground Strap
Instrument/
Component
Gr\Ep Skin
IM Panel
Titanium Inserts
Gordon Casto - Spacecraft Structure
13
Harness Layout
Gordon Casto - Spacecraft Structure
14
Plumbing Layout
High & Low Pressure
Valves & Latches
Pressurant Tank
Fuel Tanks
Thruster
Gordon Casto - Spacecraft Structure
15
Structure Trades
• Major trade for current design.
– Tank to tank cylinder, tank to skirt cone, with AM and IM
and additional bracketry to support HGAS
VS.
– Decks and Panel design as Presented
• Selected based on mass, schedule & access
• Considered
– ISO Thermal aluminum IM vs. Low CTE IM
– 65” DIA PAF w/ side by side tank configuration
Gordon Casto - Spacecraft Structure
16
Spacecraft Bus Mass Allocations
SUBSYSTEM
S/C Bus Structure
COMPONENTS
Propulsion Module
Avionics Module
Instrument Module
Fasteners
Heat Pipes
Radiators
ALLOCATION
(kg)
289.4
155.3
54.0
22.5
15.0
18.0
24.6
Gordon Casto - Spacecraft Structure
CURRENT
BEST ESTIMATE
(kg)
258.9
138.0
48.0
20.0
15.0
16.0
21.9
MARGIN
(%)
11.8%
12.5%
12.5%
12.5%
0.0%
12.5%
12.5%
17
Mass Properties Stowed
X
Y
Z
Gordon Casto - Spacecraft Structure
18
Mass Properties Deployed
X
Y
Z
Gordon Casto - Spacecraft Structure
19
Fields of Regard Clearances Star Cameras
Gordon Casto - Spacecraft Structure
20
Fields of Regard Clearances LEND
Note: Collimated FOR not shown, no issues
-X
Mini-RF Antenna
+Y
+X
Gordon Casto - Spacecraft Structure
21
Fields of Regard Clearances LOLA
-X
Gordon Casto - Spacecraft Structure
22
Fields of Regard Clearances LAMP
-X
Gordon Casto - Spacecraft Structure
23
Fields of Regard Clearances WAC
-Y
-X
Gordon Casto - Spacecraft Structure
24
Fields of Regard Clearances LROC
-X
Gordon Casto - Spacecraft Structure
25
Fields of Regard Clearances CRaTER
Gordon Casto - Spacecraft Structure
26
Fields of Regard Clearances DIVINER
-Y
+X
Gordon Casto - Spacecraft Structure
27
Fields of Regard Clearances DIVINER
+Y
-X
Gordon Casto - Spacecraft Structure
28
Structure Component Verification
•
Proposed Coupon & Sample Testing
– Fiber Volume, Void Content, & (CTE if required)
•
•
•
•
•
•
•
•
IM Facesheet Tensile Test
Lap Shear Testing – Qualifies Surface Preparation
NDE Testing on Sandwich Bond – Qualifies Panel Consolidation Process
Flatwise Tension – Taken from Panel Cutouts, Trim, or Witness coupon
Insert Strength Testing - Pull Out & Tear Out (SDO DATA)
4 Point Bend Test – Facesheet Compression
Mole Kit–Postbond Inserts with EY3010 & Outgassing Test (SDO)
Quality Control Process
– Develop Written Procedures for unique or new processes
– Maintain Quality Records (Adh. Mix & Batch Data, Hardness Samples)
– In Process & Final Inspections
Gordon Casto - Spacecraft Structure
29
LRO PDR
High Gain Antenna System
Greg Martins
February 8, 2006
HGAS Key Level 4 Requirements
•
The HGA GIMBAL SYSTEM shall consist of a +/- 90 deg AZIMUTH GIMBAL
whose spin axis is coincident with the S/C Y-AXIS, and a +/- 90 deg
ELEVATION GIMBAL that rides on the AZIMUTH GIMBAL and has a spin
axis that is coincident with the S/C X-AXIS when the HGAS is deployed and
when the AZIMUTH GIMBAL is at its 0 deg location.
•
The HGA DISH shall be attached to the ELEVATION GIMBAL such that it
points along the S/C -Z axis when both the AZIMUTH GIMBAL and the
ELEVATION GIMBAL are at their 0 deg locations in HGAS deployed
configuration.
•
In Nominal mode, the contribution of HGAS operation to nominal Observatory
attitude control disturbance effects shall not exceed the contribution defined in the
LRO Pointing, Jitter, and Alignment Budget.
•
HGAS shall meet all pointing-related allocations given in the LRO Pointing and
Alignment Specification (431-SPEC-000113).
Greg Martins - HGAS
2
HGAS Key Level 4 Requirements
•
The Gimbal Actuators shall be designed for a 5500 cycle life (14 months)
•
The HGAS shall include a Laser Ranging Telescope per ICD drawing
GE2079400
•
The HGAS STOWED Frequency GOAL shall be a minimum of 35 Hz.
•
The HGAS DEPLOYED Frequency shall be a minimum of 0.5 Hz, with a
minimum GOAL of 1.0 Hz.
•
The HGADS shall incorporate a potentiometer in the deployment hinge to
monitor the 90 degree deployment to fully deployed position.
•
The Gimbal Actuators (motors) shall use redundant motor windings and
redundant position feedback mechanisms (Encoder / Resolver)
Greg Martins - HGAS
3
HGAS Pointing Budget
POINTING ERROR BUDGET
LRO High Gain Anntenna System
Parameter (3σ values)
Bias
1
Random/
Known
Gnd-to(Deg)
6-Dec-05
Random (Deg.)
Subsystem
Very Low Low Freq. High Freq. Requirement.
Freq.
3
A/T
2
Orbit
(Deg)
Req #
ACS/GN&C Knowledge/Command Errors
ACS pointing knowledge
Ephemeris accuracy
Algorithm accuracy
Hardware Alignment Errors
Antenna boresight error
HGADS I/F to S/C Ref error
5
Boom to HGADS I/F error
Gimbal to boom axis co-alignment error
Gimbal to Gimbal-HGA I/f alignment error
HGA to Gimbal-HGA I/f alignment error
Gimbal interaxial orthogonality
Launch/Deployment/Gravity Release
Errors
HGADS launch shift
7
0.50
0.50
0.02
0.02
HGAS/Mech
HGAS/Mech
A/T
T
T
0.13
0.13
0.14
0.14
0.02
0.02
0.02
0.02
HGAS
Com/Gimbal
Com/Gimbal
Gimbal
T
T
T
T
0.00
HGADS
A
0.10
HGAS
0.50
0.55
0.00
HGADS
Gimbal
Com
A/T
T
T
A
(0.00)
0.00
0.01
Com
Mech
A
A
(2.00)
0.01
Mech
A/T
0.04
HGAS/ACS
0.02
0.02
0.08
ACS
ACS
Gimbal
A/T
A
A
A
(2.00)
Dynamic Pointing Errors
Gimbal/boom dynamic interaction
ACS (RW) induced boom dynamics
Other S/C induced dynamics
Gimbal tracking error
Thermal Distortion
S/C ref through Boom to El bracket
Gimbal, El bracket to antenna
Antenna
Column Totals (RSS random, linear
alignment), total on orbit error (deg)
Column Totals (RSS), total on orbit error
(after compensation and on-orbit
9,10,11
calibration)
4
Com
6
HGADS to S/C reference launch shift
HGADS to S/C reference gravity release
A
A
A
0.01
6
Antenna launch shift
Antenna gravity release
ACS
ACS
ACS
0.15
6
HGADS/gimbal gravity release
HGADS Deployment error
Gimbal actuator interface launch shifts
0.02
0.01
0.01
0.02
0.03
0.02
0.01
0.02
1.72
0.75
0.00
0.05
0.03
0.02
0.02
0.03
0.10
0.03
0.03
0.10
(deg)
MAXIMUM ERROR ALLOWED
Greg Martins - HGAS
HGAS
Gimbal
Com
Total Uncompensated
error (deg)
Total Error
Post
Calibration
(deg)
• LRO Pointing
Accuracy
requirement (+/- 0.3
deg) same as SDO’s
• HGAS Lead has
direct control over
the highlighted
items
8
4
A
A
A/T
2.62
0.20
0.30
4
HGAS Overview
•
HGAS Components
– Deployment System
• 2 Latches requiring mechanical release
• 3 Restraint areas not requiring mechanical release
– Articulation System
• 2-Axis Gimbal (+/- 90 degrees, each actuator)
• 2 Rotating Cable Wraps (+/- 96 degree capability)
– RF Components
•
•
•
•
•
•
High Gain Antenna, Ka band
S-Band Patch Antenna
S-band coax cable
Ka Band waveguide
2 Duel axis RF Rotary Joints – one at each Gimbal Actuator
1 Ka Rotary Joint in Hinge
– Laser Ranging Telescope
• Small Telescope (0.2 kg)
• Fiber Optic Cable
– Thermal & Electrical
• Temperature control items
• Cable harness
Greg Martins - HGAS
5
Trade Studies
Trade Study
Options
Criteria
Outcome
Latch Release
Actuators
• Hi-Shear Sep Nut
• Qwknut
• G&H Sep Nut
•Resettable?
•Reliable/Flight Proven?
•Requires X-Ray after
reset?
•Uses 28 v, 5 amp circuit
Hi-Shear Sep
Nut
Gimbal
Configuration
•X/Y
• Az / El like SDO
Provide Hemispherical
coverage
X/Y
Gimbal
Actuator
• SDO Actuator
• Moog Type 6
• Load capability
• Schedule
• Thermal properties
TBD (1.5 weeks)
Analysis in work
Greg Martins - HGAS
6
HGAS Stowed
HGAS Latch
areas
Solar Array
HGAS Hinge
Greg Martins - HGAS
7
HGAS Deployed
HGAS
Greg Martins - HGAS
8
HGAS Stowed Configuration with Latches shown
UNRELEASED
S-Band Antenna
2.1 m
HGA
2-axis Gimbal &
Latch area
RF Wave Guide
Hollow Aluminum boom
Greg Martins - HGAS
Hinge
9
HGAS Gimbal and Latch Area
Titanium Bracket to reduce heat transfer
from HGA Horn to Gimbal Actuator
Dual Axis RF
Rotary Joints
Waveguide
straight all the
way to HGA
Rotary Cable Wraps on Gimbal Actuators
HGA
Thermal Isolator concept (Titanium or TBD material), provided by HGA
vendor, to reduce heat transfer from HGA Horn to Gimbal Actuator
Greg Martins - HGAS
10
HGAS Gimbal and Latch Area
Laser Ranging
Telescope
Z Latch
Z Restraint
X
Y
YZ Restraints (use V-guides)
YZ Latch
Greg Martins - HGAS
11
Upper Latch & Restraint Area DEPLOYED
Laser Ranging
Telescope
YZ Restraints
(V-guides)
FOV shown going
through hole in HGA
dish
Z Latch bolt
retracted
X
Y
Z Restraint
Pad
Greg Martins - HGAS
12
“Z” Latch X-section
“ Z ” Bolt Retractor Assembly
(SDO’s)
X
• Bolt, ¼ inch dia
Z
• Belleville washers for pre-load
• Spring for retraction
• Swivel
Flex Mount
• Swivel Bearing (mono-ball)
• Hi-Shear Sep-Nut, free to rotate
• Flex-Mount
(Flexes only in +/-X)
Latch designed to avoid bolt bending during thermal expansion
Greg Martins - HGAS
13
“YZ” Latch
Same concept as Z Latch…
• Except this has V-guide
• “V-Guide” may help with
integration alignment
more than a “Cup-Cone”
Pyro’s
Y
Z
Greg Martins - HGAS
14
Deployment Hinge – Same as SDO Hinge
Deployment Damper inside Wire Wrap Drum
Constant Force
Spring
Cable Harness (runs
up side of Boom)
RF Rotary Joint
Potentiometer
Greg Martins - HGAS
15
Cable Harness Makeup
Greg Martins - HGAS
16
HGAS Allocations
SUBSYSTEM
High Gain Assy.
COMPONENTS
High Gain Antenna
WG-34 Ka Band Waveguide
HGA Hinge
HGA Gimbals
HGA Deploy Latches
HGA Gimbal Controller
Harness
HGA Boom
Thermal
ALLOCATION
(kg)
44.7
6.0
3.5
7.1
9.7
1.7
8.8
1.8
4.9
1.2
CURRENT
BEST ESTIMATE
(kg)
38.2
5.0
2.8
6.4
8.8
1.6
7.0
1.5
4.1
1.0
Greg Martins - HGAS
MARGIN
(%)
17.1%
20.0%
25.0%
10.0%
10.0%
10.0%
25.0%
25.0%
20.0%
20.0%
17
Schedule - Key HGAS Deliverables
Item
UID
Date
Complete HGAS Design & Analysis
4144
June 2006
Receive Latch Actuator Sep-nuts
5387
Feb 2007
Complete ETU HGAS assembly
4151
Feb 2007
Fit Check ETU HGAS to Flight spacecraft
5336
April 2007
ETU HGAS to ETU S/C for S/C Qual Testing
5671
June 2007
Receive HGA, Waveguide items
5372
Jun 2007
Receive Gimbal Actuators
5358
July 2007
Complete FLIGHT HGAS assembly
5371
Sept 2007
Complete FLIGHT HGAS Testing (env & funct)
5375
Jan 2008
Deliver FLIGHT HGAS to S/C I&T
5370
Feb 2008
Greg Martins - HGAS
Status
RFP goes out in March 06
18
LRO PDR
Solar Array Mechanical Design
Michael Hersh
February 8, 2006
Agenda
•
•
•
•
•
•
•
•
Introduction
Requirements
General Arrangement
Trade Studies
Design Details
Mass
Issues
Wrap-up
Michael Hersh - Solar Array System Mech
2
Solar Array Deployment and Articulation System
Panel size 2.8 x 1.47 m (9.2 x 4.5 ft)
3 Rigid Panels with
Solar cell Modules
Containment
Canister for
Restraint Bolt
Restraint System:
Flexure with
Separation-Nut
Release Device
Panel Hinge
with
Potentiometer
X
Z
Y
Panel Hinge
with Damper
Gimbal Elevation Actuator
with Cable Wrap
Sun Sensors
(4 on array)
Michael Hersh - Solar Array System Mech
S/C Interface
Bracket
Gimbal Azimuth Actuator
with Cable Wrap
3
Key Requirements
Category
Requirement
Mounting Platform,
Solar Cells
The SAS shall provide a mounting platform to accommodate cells producing 1849 W EOL; 3696
cells shall be grouped in 154 strings (24 cells per string)
Sun Tracking
• 2-axis gimbal; rotation-1 about the S/C Y-axis (azimuth); rotation-2 about an axis in the XZ
plane (elevation)
• Range of motion = 180 degrees for azimuth rotation; 90 degrees for elevation rotation
• Pointing accuracy = +/-2 degrees
• Tracking speed = 3.2 deg/min for azimuth rotation, = 1 deg/day for elevation rotation
• Gimbal life = 5500 rotate/rewind cycles
Structural Design
• Withstand launch: 14.5 g’s limit load any direction
• Stowed frequency goal 35 Hz minimum
• Deployed frequency goal 1 Hz minimum
Mechanism Robustness
Force and Torque margins >= 3
Jitter
Disturbance torques as required for pointing and alignment stability (431-SPEC-000113)
Thermal
•
•
•
•
•
Volume, Stowed
Fit in the Delta-IV and Atlas-5I Launch Vehicle Fairing Static Payload Envelopes.
Survival panel temperatures of –160 to +145 deg C
Survival panel temperatures of –125 to +135 deg C
Hinge damper operating temperature of –10 to +50 deg C
Gimbal operating temperature of –10 to +50 deg C
Restraint/Release operating temperature of –10 to +50 deg C
Michael Hersh - Solar Array System Mech
4
Key Requirements
Requirement
Telemetry
• Panel angle position during array deployment
• Actuator positions during panel articulation
Redundancy
The Mechanical System shall Adopt a Single String approach for the Mechanical/Structural and
Mechanical/Mechanisms design and development and perform adequate tests and analyses to verify design.
Selective Redundancy shall be allowed.
Power
• Gimbal average power consumption shall < 6 watts per actuator
• Gimbal maximum power consumption shall < 13 watts per actuator
Mass
87 Kg maximum
Requirements captured in 431-RQMT-000491, Solar Array Deployment and Articulation System Level 3
Requirements (in process)
Michael Hersh - Solar Array System Mech
5
Size of Solar Array
•
•
•
•
Y
154 solar-cell strings
24 cells per string
3696 cells
~10m^2 of power-generating area
X
Z
2.8 m
(9.2 Ft)
4.22 m
(13.85 Ft)
Michael Hersh - Solar Array System Mech
6
General Arrangement
•
•
•
Configuration
– One articulated wing
– Wing consists of three panels; number of cells = 3696; total array area ~ 10m2; power-generating
capability = 1849 watts EOL
– Dual-axis gimbal mounted directly to Spacecraft (no boom)
– Four Panel Hinges: two at each hinge line connecting center panel to two outer panels
– Outer-panel cells face outward in stowed configuration
– Panels spring-driven with viscous dampers to dissipate energy
– Deployment controlled by cam latches
Articulation
– Deployed wing rotates about two axes, with power going thru rotary joints at each axis
– Array rotation accomplished using actuators similar to the XTE/TRMM units; possible vendors
include Moog (Flight heritage) or Starsys (identical to SDO units currently being flight-qualified);
competitive procurement
– Rotation required is 180 degrees for azimuth actuator, 90 degrees for elevation actuator
– Position knowledge via incremental encoders (actuators) and potentiometers (panel hinges)
Stowed System
– Array is semi-kinematically mounted
• Total of 5 attachment points to S/C; 4 on panels + gimbal attachment
– Each Restraint Device utilizes a Separation Nut with redundant Pyro Initiators to release the array
Michael Hersh - Solar Array System Mech
7
Solar Array Deployment Sequence
Deployment
–
–
–
–
Initiation
• Deployment actuators remove restraints between S/C and array
• Restraint bolts pulled to outer panel and captured
• Panel-to-panel latches prevent panels from immediately unfurling
Elevation Gimbal Rotation
• Elevation gimbal rotates 90 degrees
Panels unfurl
• Latch releases inside panel (after 90-degree gimbal rotation)
• Kick-off spring helps to initiate panel motion (low force, long stroke)
• Torsion spring at hinge drives hinge open
• Rate of opening is controlled by rotary viscous dampers
• When inside panel reaches predetermined angle (~60 degrees), latch releases middle panel, which is
then likewise driven open
Panel Lockup
• Travel ends as hinges contact hard-stops
• Latches lock hinges rigidly in position
• Power is removed from damper heater circuits
Orbit Day
–
Tracking - the SA drive actuators are commanded to rotate at constant velocities (azimuth ~3.2 deg/min,
elevation ~1 deg/day) to maintain pointing at the sun (open loop) as the spacecraft circles through the orbit
day
–
Rewind & Reacquisition - the SADA slows to a stop, and then rewinds so that the array will be in the proper
orientation at orbit dawn
Orbit Night
Michael Hersh - Solar Array System Mech
8
Solar Array Deployment Sequence
2
1
3
Release S/C restraints;
begin driving gimbal actuator-2
4
5
Inner panel unlatched when gimbal
actuator reaches predetermined
angle (~90 degrees)
6
Michael Hersh - Solar Array System Mech
Middle Panel Unlatched When Inner
Panel Reaches Predetermined angle
(~ 60 degrees)
9
Solar Array Deployment Sequence
8
7
10
9
11
12
Michael Hersh - Solar Array System Mech
10
Solar Array Motion to Track Beta Angle
Beta 0 Beta 10
Beta 20
Beta 30
Beta 40
3 Months
Beta 50
Beta 60
Beta 70
Beta 80
Beta 0
Beta 90
Beta 10
Beta 20
Beta 30
Y
Beta 40
Beta 50
Beta75-90
Z
Beta 60-75
S/C
Coordinate
System
Michael Hersh - Solar Array System Mech
11
Array Motion to Track Sun During an Orbit
Beta 10
Beta 0
Beta 20
10 Days
(1 degree/day)
Gimbal-2
180-degree rotation
and rewind in 113
minutes to track
sun during orbit
Gimbal-1
Beta 30
Beta 50
Beta 40
Michael Hersh - Solar Array System Mech
12
Array Motion to Track Sun During an Orbit
Beta 60-75
Beta 75-90
Michael Hersh - Solar Array System Mech
Sun Safe
13
Array – WAC Field Of Regard (FOR) Issue
Wide Angle
Camera (WAC)
Wide Angle
Camera (WAC)
Array at
Beta-60
Orientation
Small
Clearance
WAC Field of Regard (FOR)
110 degrees cross-track by
50 degrees along-track
Michael Hersh - Solar Array System Mech
14
Trade Studies
Trade Study
Options
Criteria
Outcome
Potential Impingement of Solar
Array into Wide Angle Camera
(WAC) Field of Regard
• Place Array at end of deployable boom
• Place WAC on boom
• Solve via operational constraints
• Mass
• Number of required launch restraints
• Complexity of on-ground
deployments
• Power Generation
• Thermal
Solve via operational
constraints
• Operate normally from beta-0 to beta-60
• From beta-60 to beta-75, hold array at
beta-60 elevation and track normally
with azimuth rotation
• From beta-75 to beta-90, park array in
beta-90 position
Lightweight or Conventional Array
• Ultraflex
• Multi-panel rigid array
• Design Maturity/Heritage
• Simplicity/robustness
• Mass
Multi-panel rigid array
Rigid-Panel Substrate Design
• Honeycomb panel
• I-Beam Frame with Modules
•
•
•
•
•
I-Beam Frame with
Modules
Gimbal-Axes Configuration
• Azimuth Elevation
• XY
• Best suited to LRO orbit - smooth
motion; minimum necessary gimbal
motion
Schedule
Structural Integrity
Required Panel Thickness
Simplicity
Design Maturity/Heritage
Michael Hersh - Solar Array System Mech
Azimuth-Elevation
(Azimuth = S/C Y-axis)
15
Solar Panel Substrate Design
• Modules (substrates, cells, wiring) are bonded to
frame after delivery from cell vendor
• Panel = Frame + Modules
• 3 panel frames, 77 modules
• Frame material M55J composite, quasi-isotropic
(QI) lay-up, I-beam construction, bonded assembly,
depth ~ 38.1 mm (1.5 inch )
• Module Substrate = honeycomb panel, 6.35 mm
(.25 inch) thick, M55J QI facesheets, aluminum
core
ules
s
odule
25 M
s
odule
M
7
2
• Schedule benefit – fabrication of frame and modules
are completely independent
• Heritage: SMEX, WIRE, TRIANA
• Module size 44.9 x 28.4 cm (17.68 x 11.18 inch)
• Module = 2 strings = 48 cells
od
25 M
1 Module
Michael Hersh - Solar Array System Mech
16
Solar Panel Substrate Design
• Open Frame allows backside of modules to radiate thermal
energy
• Fittings for hinges, restraints, sun sensors, GSE handling
points, etc., are titanium for best CTE match; bonded to
frame
Michael Hersh - Solar Array System Mech
17
Solar Array Restraint Configuration
X
X
(cup/cone)
Z
Z
Panels
restrained in
local Y and Z
directions only
X
Z
Z
Y
LOCAL X
Array Restraint Concept
SADA attach to S/C
XYZ
Michael Hersh - Solar Array System Mech
18
X
Solar Array Restraint Configuration
Current Analytical results, Natural Frequencies
• Stowed ~ 35 Hz
• Deployed ~ 0.7-0.95 Hz (design and FEM
evolving)
Separation-Nut Cross-Section and Function
PYRO INITIATOR
HOUSING
A few Sep-Nut Designs with
Redundant Initiators
(Hi-Shear Corp.)
BOLT EJECTOR
NUT SEPARATOR
NUT (3 SEGMENTS)
RESTRAINING
CYLINDER
5.72 CM
(2.25 INCH)
5.72 CM
(2.25 INCH)
Michael Hersh - Solar Array System Mech
19
Panel-to-Panel Hinges
•
•
•
•
•
•
•
•
•
•
Derivative of COBE, XTE, TRMM design
Spherical bearing to allow for thermal distortions, tolerance build-up, and
inaccuracies in the assembly
Hinge line consists of a pair of hinges, identical except for mounting
interfaces to damper/potentiometer
Axial clearance at one hinge to accommodate thermal distortions on-orbit
Positive latch on each hinge with no deadband (backlash), roller follows
cam to produce virtually no latch friction during deployment
Titanium 6Al-4V housings and latch pin (with Tiodize V surface
treatment)
Torsion Spring at all hinges (redundancy)
Potentiometer for position telemetry during testing and deployment
Rotary viscous damper for energy dissipation
Redundant rotational path: shaft thru ball and ball thru race
Michael Hersh - Solar Array System Mech
20
Panel-to-Panel Hinges
Rotary Viscous
Damper to
control rate of
panel unfurling
Thermostats
bonded to housing
(not shown)
Potentiometer to
monitor panel
deployment
Pin-Slot Coupling
Adjustable Stop
Strip Heaters on
Damper Housing
(not shown)
Structure
Torsion Spring
Flexible
coupling
Spherical Bearing
Hinge with Damper
Hinge with Potentiometer
Michael Hersh - Solar Array System Mech
21
Articulation System
•
Actuator
– Purchased item
– Similar to units used for previous, successful GSFC missions
– Size comparable to Moog Type 6 or SDO Starsys actuator; it appears that both of these can meet or
exceed LRO performance requirements for step size, torque capability (both powered and
unpowered), stiffness (bending and torsional), load capability, slew speed, etc.; Initial CoupledLoads Analysis to be completed in February
– TRMM, XTE, and EO-1 utilized a so-called Moog Type 5.5
• Life-tested to 10 million degrees-of-travel
• EO-1 surpassed 10 million degrees-of-travel in orbit
• LRO requirement for azimuth unit is 2 million degrees-of-travel (5500 cycles, 14 month
mission); requirement for elevation unit is 420 degrees-of-travel (1.17 cycles)
– SDO is in the process of qualifying the Starsys actuator
– Weight = 4 Kg
•
Cable Wrap Mechanism
– Heritage: XTE/TRMM/EO-1 SADA Cable Wrap Mechanisms (GSFC design)
– Will size unit to provide 50% more rotation capability than required; requirement is +/- 96 degrees
(position of hard-stops)
– XTE/TRMM design successfully life tested to >10 million degrees-of-travel; LRO requirement for
azimuth unit is 2 million degrees-travel (5500 cycles, 14 month mission)
Michael Hersh - Solar Array System Mech
22
Solar Array Drive Assembly (Typical
Configuration)
(12.7 CM)
Michael Hersh - Solar Array System Mech
23
Solar Array Cable Wrap Mechanism
Cross section View
0.25 mm (0.01 inch)
Kapton Separator
Spiral Ribbon
Harness
Used successfully on XTE,
TRMM, & EO-1 Missions
Michael Hersh - Solar Array System Mech
24
XTE Solar Array Cable Wrap Mechanism (ETU)
(similar arrangements used for TRMM and EO-1)
Front View
Rear View
Harness transitions from circular bundle
into two flat, woven harnesses, which
are wrapped around each other in a
spiral fashion, and exit the housing
Harness bundle passes
through actuator quill shaft
and enters cable-wrap
A kapton sheet is placed between the
spirals for minimum resistance of cables
contacting each other
The cables are clamped at the inside
and outside diameters; all motion
occurs in the cable-wrap cavity
Michael Hersh - Solar Array System Mech
25
Wires from Solar Array to Pass through
Actuators and Cable Wrap Mechanisms
QU Cond AWG
28pr 56
20
2pr
4
24
4pr
8
20
6pr
12
24
3pr
6
26
3pr
6
24
2pr
4
24
4s
4
24
2pr
4
22
2pr
4
22
2pr
4
26
3pr
6
22
3pr
6
22
3pr
9
24
3pr
9
24
2pr
4
22
2pr
4
26
3pr
6
22
3pr
6
22
3pr
9
24
3pr
9
24
180
82pr + 4 GND
S/C
Boom Hinge
Power
Performance Monitors (2)
CSS
Thermistor PSE data
Motor Drive Deploy
Motor Drive Deploy
Motor Limit Switch TB
GND
Heater Power (Deploy)
Heater Power
Thermistor
Motor Drive
Motor Drive
Excite, decode
Excite, decode
Heater Power
Thermistor
Motor Drive
Motor Drive
Excite, decode
Excite, decode
Total
AZ/Yaw
Rotary Joint EL/Pitch Rotary Joint SA
Panel
Panel
CSS
Panel
Motor
Motor
Switch
GND
Heater
Heater
Thermistor
Motor Primary
Motor Redundant
Encoder Primary
Encoder Redundant
Heater
Thermistor
Motor Primary
Motor Redundant
Encoder Primary
Encodr RedundantResolver Red
56
4
8
12
6
6
4
4
4
104
gimbal-2
56
4
8
12
6
6
4
4
4
4
4
6
6
9
9
142
gimbal-1
142 conductors in harness
bundle to pass through center of
the Azimuth actuator
Representative harness mockup built: bundle diameter ~ .87
inches
Michael Hersh - Solar Array System Mech
26
Two-Axis Actuator Controller
1553
BUS A
1553
BUS B
1553 Interface
R/T
FPGA
(Interfaces to R/T
And 3 Phase
Motor Driver
TLM ckt)
Prim/Red.
3 Phase Motor
Drivers
Prim.
Motor
And
Encoder
Red.
Motor
And
Encoder
Prim./ Red. Encoder
Interface ckt
28V A
28V B
DC to DC Converter
And
EMI Filters
FPGA
(Interfaces to R/T
And 3 Phase
Motor Driver
TLM ckt)
Prim.
Motor
And
Encoder
Prim/Red.
3 Phase Motor
Drivers
Red.
Motor
And
Encoder
Prim./ Red. Encoder
Interface ckt
Michael Hersh - Solar Array System Mech
27
Two-Axis Actuator Controller
Interfaces
• +21V TO +35V power input from S/C
• +28V for motor drivers
• MIL-STD 1553 command and data interface and ‘simple’ parallel bus interface
Functions
• Interpret commands
• Gather and convert telemetry
• Drive actuators
• Close loop for 2 HGA antenna actuators or 2 Solar Array Drive actuators.
Command and Telemetry
• Commands (per axis): Direction, Actuator Enable, Pulse Rate, Go to Position, Initialize,
Mode of operation, Telemetry Request, Micro-stepping
• Telemetry (per axis): Actuator status (Home, End of Travel), Actuator position, Actuator
voltage and Current, Box level Voltage and Current
Actuators
• 3-Phase stepper motor and capable of micro stepping
Sensors
• Absolute or incremental optical encoder
• Resolver
Michael Hersh - Solar Array System Mech
28
Solar Array System Allocations
SUBSYSTEM
Solar Array Assy.
COMPONENTS
Solay Array Panels
S/A Gimbals
SA Deploy Latches
SA Gimbal Controller
Harness
S/A Boom
Thermal
ALLOCATION
(kg)
87.0
56.4
12.5
5.0
8.1
1.8
2.1
1.2
CURRENT
BEST ESTIMATE
(kg)
75.8
49.1
10.8
4.5
7.0
1.5
1.8
1.0
Michael Hersh - Solar Array System Mech
MARGIN
(%)
14.8%
15.0%
15.0%
10.0%
15.0%
15.0%
15.0%
20.0%
29
Cell Shadowing Evaluation
Issue
Technical Resolutions
4 of 154 solar-cell strings potentially shaded by restraint
containment-canisters when array is oriented 10-to-15
degrees from sun (beta-70 to beta-75); 2.6% reduced power
• Minimize height & diameter of canister
• Short eclipse time may render this a non-issue
• Revisit power requirements for these beta angles
Shadow with array at beta60 and sun at beta-75
Michael Hersh - Solar Array System Mech
30
Wrap-Up
•
•
Frame + Module construction = significant schedule benefit
All components based on designs having extensive GSFC flight heritage
Michael Hersh - Solar Array System Mech
31
LRO PDR
Mechanical Analysis
Greg Clarke
February 8, 2006
Mechanical Analysis Agenda
1. Loads and Environments
(L&E)
(Requirements)
- Mechanical System
Specification
2. Orbiter Analysis (OA)
-
Frequencies
Base Drive
Stress
STOP
Vibroacoustics
Shock
Future Work
3. Instrument Module and
Deployables Analysis
(IM&DA)
-
Frequencies
Stress
IM Module STOP
Vibroacoustics
Shock
Future Work
4. Coupled Loads Analysis
(CLA)
Greg Clarke - Mechanical Analysis
2
1. L&E
Requirements Documents
1. LRO Mechanical Systems Specification (LRO 431-RQMT000012)
2. Atlas Launch System Mission Planner’s Guide, September
2001
3. Delta IV Payload Planner’s Guide, October 2000
4. LRO Integration and Test Plan (LRO 431-PLAN-000100)
5. LRO Pointing and Alignment Specification (LRO 431SPEC-000113)
Note: LRO adheres to the Goddard Golden Rules; this means
GSFC-7000 for Mechanical Analysis and Verification.
Greg Clarke - Mechanical Analysis
3
1. L&E
• The LRO Mechanical Systems Specification (LRO 431RQMT-000012) includes:
• Launch Loads and Environments
–
–
–
–
–
Limit Loads (static and low frequency transient)
Sine
Vibroacoustics
Shock
Pressure Profile
• On-orbit Environment
– Jitter
– Thermal
Greg Clarke - Mechanical Analysis
4
1. L&E Quasi-Static/Design Limit Loads
• Orbiter quasi–static limit loads were developed from
PPG.
• Subsystems (PM, HGAS, and SAS) design limit
loads will be derived by base driving the observatory
to PPG net CG.
• Quasi-static IM, instruments and components limit
loads developed from mass acceleration curves.
Greg Clarke - Mechanical Analysis
5
1. L&E – Limit Loads – Primary Structure
7
6
5
4
Axial (g)
3
2
1
0
-1
-2
-3
-2.5
-2.0
-1.5
-1.0
-0.5
0.0
0.5
1.0
1.5
2.0
2.5
Lateral (g)
Atlas V 401
Delta IV Medium
Positive axial load denotes compression
Lateral loads may act in any direction
CG Limit Loads for the LRO Primary Structure
Greg Clarke - Mechanical Analysis
6
1. L& Instrument and Component Limit Loads
LRO Mass Acceleration Curve
40
35
Acceleration [g]
30
25
20
15
10
0
20
40
60
80
100
120
140
160
180
200
Mass [Kg]
MASS ACCELERATION CURVE
Source: Mechanical Systems Specification (431-SPEC-000012)
Greg Clarke - Mechanical Analysis
7
1. L&E – Orbiter Sine Vibration
Orbiter Sine Vibration
Peak Acceleration [G's]
10
Lateral (A5) LL
1
Lateral (D4) LL
Freq. [Hz] Lateral (A5) LL
Lateral (D4) LL
5
0.4
0.7
64.9
0.4
0.7
65.1
0.5
0.7
79.9
0.5
0.7
80.1
0.55
0.7
84.9
0.55
0.7
85.1
0.6
0.7
100
0.6
0.7
Sw. Rate
TBD
TBD
0.1
1
10
100
Frequency (Hz)
Note: 1) LRO will be tested up to 50 Hz and analyzed for this environment from 50
to 100 Hz (there are no LV coupled load events between 50 and 100 Hz).
2) This spec will be used for testing as well as input for Base Drive Analysis
to recover design limit loads( e.g PM subsystems).
Source: Mechanical Systems Specification (431-SPEC-000012)
Greg Clarke - Mechanical Analysis
8
1. L&E – Orbiter Sine Vibration
Orbiter Sine Vibration
Peak Acceleration [G's]
10
Thrust (A5) LL
1
Thrust (D4) LL
Freq. [Hz]Thrust (A5) LLThrust (D4) LL
5
0.6
1
19.9
0.6
1
20.1
0.8
1
29.9
0.8
1
30.1
0.6
1
74.9
0.6
1
75.1
0.7
1
79.9
0.7
1
80.1
0.9
1
100
0.9
1
Sw. Rate
TBD
TBD
0.1
1
10
100
Frequency (Hz)
Note: Limit level for the Delta IV is 0. 5 in DA from 5-6.2 Hz
Note: 1) LRO will be tested up to 50 Hz and analyzed for this environment
from 50 to 100 Hz
2) This spec will be used for testing as well as input for Base Drive
Analysis to recover design limit loads( e.g PM subsystems).
Source: Mechanical Systems Specification (431-SPEC-000012)
Greg Clarke - Mechanical Analysis
9
1. L&E -Instrument and Component Sine
Vibration
Instrument and Component Sine Vibration
10
Peak A cceleration [G 's]
Freq.[Hz]
5
50
100
Sw. Rate
PF / Qual.
8
8
8
4 / 2 o/m
Accep. Level
6.4
6.4
6.4
4 o/m
PF / Qual.
Accep. Level
1
1
10
100
Frequency (Hz)
Note: 1) Instruments and components must be tested up to 50 Hz and
analyzed for this environment from 50 to 100 Hz.
2) This spec was derived from a Base Drive Analysis using
peak enveloping.
Source: Mechanical Systems Specification (431-SPEC-000012)
Greg Clarke - Mechanical Analysis
10
1. L&E - Acoustics
Delta
5.
0
50
0.
0
80
0.
12 0
50
.
20 0
00
.
31 0
50
.
50 0
00
.
80 0
00
.0
31
0.
0
20
5.
0
12
.0
80
.0
50
.5
145.0
140.0
135.0
130.0
125.0
120.0
115.0
110.0
105.0
31
S P L (d B )
LRO LVs Payload Acoustic Flight Limit Levels
Frequency (Hz)
Delta IV Medium 140.0 OSPL
Atlas V 401 137.9 OSPL
IV
Frequency 140.0 OSPL
Medium Atlas V 401
137.9 OSPL
25.0
31.5
40.0
50.0
63.0
80.0
100.0
125.0
160.0
200.0
250.0
315.0
400.0
500.0
630.0
800.0
1000.0
1250.0
1600.0
2000.0
2500.0
3150.0
4000.0
5000.0
6300.0
8000.0
10000.0
0.0
119.5
122.5
125.2
126.3
128.0
129.0
130.0
130.0
130.0
130.0
130.0
129.5
128.0
125.0
123.0
121.0
119.5
118.0
116.5
115.0
113.5
112.0
110.5
109.0
107.5
106.0
114.000
119.000
125.100
123.500
125.000
125.500
127.000
127.000
128.200
128.000
127.500
127.000
127.000
125.500
123.000
120.500
117.500
115.000
113.000
115.000
114.000
112.200
113.000
115.000
114.500
115.500
116.5
Overall
140.0
137.9
Durations: 60 sec for Acceptance / Protoflight Test, 120 sec for Qual. Test
Source: Mechanical Systems Specification (431-SPEC-000012)
Greg Clarke - Mechanical Analysis
11
1. L&E – Acoustics
• Some instruments (LROC, LOLA, and Mini_RF)
and the flight observatory are required to be
acoustically tested.
• IM hardware is protoflight; AM, PM, HGAS and
SAS have ETU’s (require Qual./Acceptance
program).
Greg Clarke - Mechanical Analysis
12
1. L&E – Instruments and Components Random
Vibration
Components Random
1
P S D [g2/H z]
Freq. [Hz]
20
50
800
2000
Duration
PF / Qual
ASD [g2/Hz]
0.026
0.16
0.16
0.026
60 / 120 sec
Accept.
ASD [g2/Hz]
0.013
0.08
0.08
0.013
60 sec
PF / Qual
0.1
Accept.
0.01
10
100
1000
10000
Preliminary SEA indicates PSD
above GEVS. We are in the
process of completing analyses
and proposing necessary changes.
Frequency [Hz]
Note: Source: Mechanical Systems Specification (431-SPEC-000012)
Greg Clarke - Mechanical Analysis
13
1. L&E – LRO Shock Specifications
• As of today, the location-specific specifications are
the shock sources.
• Currently planning to develop specific shock levels
for Orbiter zones, instruments, and components
(semi-empirical/ virtual modes methods).
Greg Clarke - Mechanical Analysis
14
1. L&E – LRO Shock Sources
Sep Nut( SN9423-2)
Freq. {[Hz]
SRS Q=10
100
50
3000
4000
10000
4000
Shock Sources
10000
1000
Del ta IV (1194 PAF)
Atlas V (Type B1194 PAF)
Separ ati on Nut (SN9423-2)
Freq. [Hz] D4( 1194 PAF) A5(Type B1194 PAF)
100
150
100
1000
5000
2800
10000
5000
2800
100
These are fight limit levels,
Protoflight is 3 dB above
10
100
1000
10000
F r e que nc y [ H z ]
Greg Clarke - Mechanical Analysis
15
1. L&E – Pressure Profile: Delta IV and Atlas V
Delta IV Payload Fairing
Compartment Absolute Pressure
Envelope
Atlas V Typical Static Pressure
Profiles Inside the Payload Fairing
This spec applies to components and instruments w/ vent areas less
than 0.25 square inches of vent area for each cubic foot volume.
Source: Mechanical Systems Specification (431-SPEC-000012)
Greg Clarke - Mechanical Analysis
16
1. L&E On-Orbit Jitter
• Jitter is a joint effort between Mechanical and ACS groups to
verify compliance with Jitter allocation.
• Mechanical Analysis provides to ACS a set of eigenvalues
and eigenvectors derived from structural FEM.
• An updated set will be provided to ACS in time for the Peer
Review.
• There are no known jitter compliance issues at this time.
Greg Clarke - Mechanical Analysis
17
2. OA – Orbiter Finite Element Model
Three FEM Models (different fidelity and element types
required) :
1 Structural Model
- Stress Analysis, Normal Modes Analysis, Base Drive
1 Structural Model
- Sizing the Primary structure
1 Thermal Model
- Thermal Distortions
Greg Clarke - Mechanical Analysis
18
2. OA – Orbiter Finite Element Model
•
Finite element model (FEM) created in FEMAP and analyzed using MSC NASTRAN
•
FEM usage
– Determine loads for detailed stress analysis (basedrive, CLA)
– Analytical verification of fundamental frequency requirements
– Preliminary sizing of structural members
– Dynamic and static displacement checks
– Orbiter STOP analysis.
– Future derivation of instrument/component sine and random vibration levels
•
The (XYZ) center-of-mass location for the Orbiter with respect to the center of the interface (separation
plane) is: [38.12495, -3.532522, 0.459009] in
•
For all loads and normal modes runs the observatory dry mass was scaled up to obtain the orbiter 2000
kg (4409.24 lbm) maximum allocation.
•
Components and propulsion tanks modeled using lumped masses
•
Instrument models
– Smeared mass: Mini-RF
– Lumped mass: Crater, LAMP, LEND, LOLA LROC NAC, LROC WAC
– Physical model: Diviner
Greg Clarke - Mechanical Analysis
19
2. OA – Orbiter Finite Element Model
• Assumptions
– Instruments/components with fundamental modes greater than
75 Hz assumed rigid.
– Smeared non-structural mass captures items not physically
modeled (blankets, cable harness, fillets, inserts, etc.)
• Interfaces:
– FEM constrained in all DOF at launch vehicle interface
• Validity Checks
– All validity checks passed
Greg Clarke - Mechanical Analysis
20
2. OA – Orbiter Finite Element Model
Crater
IM
SAS
HGAS
Diviner
LEND
Greg Clarke - Mechanical Analysis
21
2. OA – Orbiter Finite Element Model
Propulsion
Tank
Propulsion
Tank
Avionics
Module
Greg Clarke - Mechanical Analysis
22
2. OA – Frequency
Axis
Thrust
Lateral
Delta IV
Medium
27
10
Atlas V
401
15
8
15%
Margin
31.05
11.5
PDR
32.96
21.17
Currently using 15% margin for PDR.
The first lateral in the thrust direction, is driven by the HGAS. Plan to revisit
this before the Peer Review.
Greg Clarke - Mechanical Analysis
23
2. OA – Orbiter Normal Modes
•
•
Mode 1 – 21.17 Hz
Lateral Bending: RZ
•
•
Mode 2 – 22.30 Hz
Lateral Bending: RY
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
24
2. OA – Orbiter Normal Modes
•
•
Mode 4 – 30.70 Hz
Torsion: RX
•
•
Mode 7 – 32.96 Hz
Thrust: X
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
25
2. OA – Orbiter Normal Modes
•
•
Mode 13 – 40.59 Hz
Torsion: RX
•
•
Mode 20 – 50.37 Hz
Torsion: RX
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
26
2. OA – Orbiter Normal Modes
•
•
•
•
Mode 27 – 58.19 Hz
Thrust: X
Mode 28 – 58.44 Hz
Thrust: X
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
27
2. OA – Orbiter FEM – SAS Deployed
•
•
•
•
The purpose of the deployed FEM is to verify the on-orbit frequency requirements
The SAS is expected to exhibit the lowest deployed frequency.
Deployed FEM identical to stowed FEM except the SAS is deployed
Free-free boundary conditions
MODE FREQUENCY
#
Rigid Body Modes
Exceeds on-orbit requirement of 0.5 Hz
Greg Clarke - Mechanical Analysis
1
2
3
4
5
6
7
8
9
10
(Hz)
0.000
0.000
0.000
0.000
0.000
0.001
0.782
0.844
0.894
3.615
28
2. OA – Orbiter Normal Modes – SAS Deployed
•
•
Mode 1 – 0.782 Hz
SAS Torsion
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
29
2. OA – Orbiter Normal Modes – SAS Deployed
•
•
Mode 2 – 0.844 Hz
SAS Bending RZ
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
30
2. OA – Orbiter Normal Modes – SAS Deployed
•
•
Mode 3 – 0.894 Hz
SAS Bending RX
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
31
2. OA – Orbiter Base Drive Analysis
•
The purpose is to investigate dynamic interactions of orbiter, assess quasistatic limit loads, and estimate interface loadings (SAS, HGAS, etc.) prior
to coupled loads analysis
•
FEM constrained at launch vehicle interface
•
Analysis was performed to 60 Hz to account for dynamic uncertainty
(DUF = 1.2)
•
Dynamic Amplification factor , Q, was used for all Base drive analysis.
•
Envelope of Delta IV and Atlas V PPG sine input specifications
•
Force limited – envelope of PPG net CG limit accelerations and orbiter
mass and CG used to develop limits for interface forces
Greg Clarke - Mechanical Analysis
32
2. OA – Orbiter Base Drive Analysis
Input limited at modes 1 and 2 in lateral directions
LRO Basedrive Limited Input Levels
1.2
1
Input Acceleration, g
•
0.8
X
Y
Z
0.6
0.4
0.2
0
0
10
20
30
40
50
60
70
Frequency, Hz
Greg Clarke - Mechanical Analysis
33
2. OA – Orbiter Base Drive Analysis
Results show that all instrument and component responses enveloped by MAC
Only flexible, massive subsystems and components have exceedences
Reductions to current levels (MAC) likely
LRO Basedrive Results Summary: Net CG Limit Loads
Instruments, & Components < 100 kg
100
Limit Load (g)
•
•
•
10
MAC
1
0.1
1
10
100
Mass (kg)
Greg Clarke - Mechanical Analysis
34
2. OA – Orbiter Base Drive Analysis
•
•
•
•
•
Summary table shows current
estimate of limit level loads
Basedrive responses above
MAC levels ignored
MAC limit load estimated
using CBE mass
All structure must survive the
maximum PPG static load (6.6
g) at a minimum
Results valid for current
configuration – modifications
to the current configuration
may cause changes in limit
loads
Item
IM Net CG
HGAS Net CG
SAS Net CG
Diviner Net CG
Propulsion Tank (2)
Pressurant Tank
Mini-RF Ebox
Mini-RF Xpander
Battery Box
Reaction Wheel (2)
C&DH
PSE
S-Band Transponder
PDE
TWTA
Ka Modulator
EPC
SA & HGA Gimbal Controller
IMU
Crater
Lamp
LEND
LOLA
LOLA Electronics Box
LROC NAC -X
LROC WAC
LROC Electronics Box
Star Tracker +X
SAS El-Gimbal Motor
Greg Clarke - Mechanical Analysis
Summary
CBE Mass
Expected Limit
lbm
kg
Load (g)
ID
Rationale
NET 151.51 68.71
14.9 Basedrive
NET
68.69 31.15
19.7
MAC
NET 151.70 68.80
15.1
MAC
NET
22.05 10.00
19.9 Basedrive
1002 1057.96 479.80
13.4
MAC
1005
33.96 15.40
23.9 Basedrive
1015
8.60
3.90
6.6
Static
1016
1.94
0.88
6.9 Basedrive
2015
66.15 30.00
9.9 Basedrive
2017
26.46 12.00
15.7 Basedrive
2020
46.75 21.20
8.9 Basedrive
2023
32.85 14.90
12.5 Basedrive
2024
11.03
5.00
10.0 Basedrive
2026
30.87 14.00
9.8 Basedrive
2027
2.21
1.00
6.9 Basedrive
2028
8.82
4.00
6.6
Static
2029
2.21
1.00
11.2 Basedrive
2030
15.44
7.00
10.7 Basedrive
2031
9.92
4.50
10.0 Basedrive
3027
12.35
5.60
7.5 Basedrive
3029
11.03
5.00
17.7 Basedrive
3030
52.26 23.70
11.0 Basedrive
3031
16.76
7.60
15.7 Basedrive
3032
6.62
3.00
14.4 Basedrive
3034
14.33
6.50
17.2 Basedrive
3035
1.54
0.70
22.4 Basedrive
3036
1.54
0.70
12.4 Basedrive
3037
6.62
3.00
17.6 Basedrive
162001
11.91
5.4
25.0 Basedrive
35
2. OA – Preliminary Orbiter Stress Analysis
• Starting from bottom of Orbiter and progressing upwards
Component
Load Case
Loads / Stresses
Allowable Loads / Stresses
Failure Mode
Factors
MS
3.4 G Axial + 2.0 G Lateral
σC, MAX = 3434 psi
σCR = 12730 psi
Compressive buckling
FSULT = 2.6
FF = 1.15
UF = 1.1
+ 0.13
-0.2 G Axial + 2.0 G Lateral
P = 1188 lb, V = 362 lb
PULT = 5700 lb, VULT = 3291 lb Tension-shear interaction
FSULT = 2.6
FF = 1.15
UF = 1.1
+ 0.37
Stub skirt fasteners ( lower )
( See Ref. 1 )
- NAS 1352 .2500-20
- 48X bolts at I/F to sep plane ( for GSE stub skirt )
-0.2 G Axial + 2.0 G Lateral
P = 893 lb, V = 347 lb
PULT = 5700 lb, VULT = 3291 lb Tension-shear interaction
FSULT = 2.6
FF = 1.15
UF = 1.1
+ 0.72
Bottom deck
( See Ref. 2 )
- AL 6061-T6
- Depth d = 2.00 in, fitting blocks 8.0 in X 4.0 in X 2.0 in
- Top skin t = 0.15 in, rib t = 0.090 in, ring to skirt t = 0.375 in
- Prop tank mounting flange t = 0.175 in, gusset t = 0.125 in
3.4 G Axial + 2.0 G Lateral
( TBD )
( TBD )
Stub skirt
- AL 6061-T6
- r = 23.5 in, h = 10.0 in, t = 0.180 in
( See Ref. 1 )
Stub skirt fasteners ( upper )
- NAS 1352 .2500-20
- 48X bolts at I/F to bottom deck
( See Ref. 1 )
Analysis of preliminary design shows peak Von Mises stresses < 15 ksi.
More design iterations on section properties and bottom deck I/Fs required.
Additional analysis iterations / consideration of other failure modes required.
Notes:
1) Ref. 1 is: 2006_01_19_Stub_Skirt_Sizing.pdf
2) Ref. 2 is: 2006_01_09_Bottom_Deck_and_Stub_Skirt_Bolts.pdf
3) FF is fitting / stability factor and UF is a factor to account for uncertainty in overall mass properties
4) Models used in the above analyses were set to 4100 lb ( 1860 kg ); future models will be set to 4409 lb ( 2000 kg ) with UF = 1.00
Greg Clarke - Mechanical Analysis
36
2. OA – STOP Analysis
•
Preliminary STOP analysis is performed by applying unit thermal
gradients (1.0 degC/m) along each axis as well as applying a unit bulk
temperature change. This preliminary analysis is used to determine if the
pointing of any of the components is sensitive to gradients in each of the
analyzed directions. A component was considered sensitive if the
resulting delta rotation was greater than 5.0 arc-sec.
–
–
–
–
No Components were sensitive to a unit bulk temperature change
The IMU is sensitive to a gradient along X
No Components were sensitive to gradient along Y
Components sensitive to gradient along Z
• Diviner
• IMU
•
A STOP analysis is planned for the observatory prior to CDR peer
review.
– Estimated date for delivery of temperatures: 3/1/06
– Estimated date to complete the STOP analysis: 3/15/06
Greg Clarke - Mechanical Analysis
37
STOP Analysis
• Deformation and Thermal Contours on the
Spacecraft
Unit Bulk Temperature Change
Greg Clarke - Mechanical Analysis
1.0 degC/m Gradient in X
38
STOP Analysis
• Deformation and Thermal Contours on the
Spacecraft
1.0 degC/m Gradient in Y
1.0 degC/m Gradient in Z
Greg Clarke - Mechanical Analysis
39
STOP Analysis
Relative Deformations between CG Locations
Relative Deformations between CG Locations
Unit Bulk 1 DegC 1 DegC 1 DegC
Case
Description
Temp.
/meter
/meter
/meter
Dir
Change
in X
in Y
in Z
R1
-0.50
-0.83
-0.50
0.19
Star Tracker +X
R2
-0.02
0.04
0.01
-0.09
wrt
R3
0.10
0.16
0.06
-0.05
Star Tracker -X
R1
-0.13
-0.16
-0.78
-1.03
Crater
R2
0.08
1.34
0.04
-3.99
wrt
R3
-0.15
0.58
4.57
-0.03
Star Tracker +X
R1
-0.63
-0.09
-2.62
0.44
Crater
R2
-0.06
3.19
-0.11
4.38
wrt
R3
-0.01
-0.93
-3.84
0.00
Star Tracker -X
R1
-0.13
-0.16
-2.68
-5.34
Diviner
R2
0.08
3.25
0.07
-4.35
wrt
R3
-0.15
4.73
4.78
-0.01
Star Tracker +X
R4
-0.63
-0.99
-3.18
-5.15
Diviner
R5
0.06
3.29
0.08
-4.44
wrt
R6
-0.05
4.89
4.85
-0.07
Star Tracker -X
R7
0.03
0.03
0.07
-0.04
Lamp
R8
0.10
0.08
0.07
0.08
wrt
R9
-0.35
-0.46
-0.35
-0.12
Star Tracker +X
R10
-0.47
-0.80
-0.43
0.15
Lamp
R11
0.08
0.11
0.07
-0.01
wrt
R12
-0.25
-0.30
-0.29
-0.18
Star Tracker -X
R13
-0.13
-0.16
-2.62
0.31
Lend
R14
0.08
3.23
-0.01
4.48
wrt
R15
-0.14
-1.15
-3.94
-0.03
Star Tracker +X
Lend
R16
-0.63
-0.99
-3.12
0.50
wrt
R17
0.06
3.26
0.00
4.39
Star Tracker -X
R18
-0.05
-0.99
-3.88
-0.08
R19
-0.09
-0.15
-0.03
-0.02
Lola
R20
0.17
0.19
0.14
0.13
wrt
R21
-0.06
-0.12
-0.03
-0.01
Star Tracker +X
R22
-0.58
-0.98
-0.53
0.17
Lola
R23
0.16
0.22
0.15
0.04
wrt
R24
0.03
0.04
0.03
-0.06
Star Tracker -X
R25
-0.15
-0.19
-0.08
-0.06
Lroc NAC +X
R26
-0.11
-0.25
-0.13
0.15
wrt
R27
-0.11
-0.21
-0.10
-0.03
Star Tracker +X
R28
-0.64
-1.02
-0.59
0.13
Lroc NAC +X
R29
-0.13
-0.22
-0.13
0.06
wrt
R30
-0.02
-0.05
-0.04
-0.08
Star Tracker -X
*Dimensions are Rx: arc-sec
Description
Max
0.83
0.09
0.16
1.03
3.99
4.57
2.62
4.38
3.84
5.34
4.35
4.78
5.15
4.44
4.89
0.07
0.10
0.46
0.80
0.11
0.30
2.62
4.48
3.94
3.12
4.39
3.88
0.15
0.19
0.12
0.98
0.22
0.06
0.19
0.25
0.21
1.02
0.22
0.08
Case
Dir
Lroc NAC -X
R1
wrt
R2
Star Tracker +X
R3
Lroc NAC -X
R1
wrt
R2
Star Tracker -X
R3
Lroc WAC
R1
wrt
R2
Star Tracker +X
R3
Lroc WAC
R1
wrt
R2
Star Tracker -X
R3
Diviner
R1
wrt
R2
LOLA
R3
Lamp
R1
wrt
R2
Lroc NAC +X
R3
Lamp
R1
wrt
R2
Lroc NAC -X
R3
Lola
R1
wrt
R2
Lroc NAC +X
R3
Lola
R1
wrt
R2
Lroc NAC -X
R3
Lroc NAC +X
R1
wrt
R2
Lroc NAC -X
R3
IMU
R1
wrt
R2
Star Tracker +X
R3
IMU
R1
wrt
R2
Star Tracker -X
R3
*Dimensions are Rx: arc-sec
Greg Clarke - Mechanical Analysis
Unit Bulk 1 DegC 1 DegC 1 DegC
Temp.
/meter
/meter
/meter
Change
in X
in Y
in Z
-0.08
-0.07
0.00
-0.13
0.14
0.03
0.09
0.10
-0.13
-0.21
-0.11
-0.03
-0.57
-0.90
-0.51
0.06
0.12
0.07
0.10
0.01
-0.03
-0.06
-0.04
-0.08
-0.39
-0.69
-0.30
-0.13
-0.29
-0.62
-0.31
-0.09
-0.07
-0.14
-0.01
0.00
-0.89
-1.52
-0.80
0.06
-0.31
-0.58
-0.31
-0.18
0.03
0.02
0.05
-0.06
-0.05
-0.01
-2.65
-5.32
-0.09
3.06
-0.07
-4.48
-0.08
4.86
4.81
-0.01
0.17
0.22
0.16
0.02
0.21
0.33
0.20
-0.07
-0.24
-0.25
-0.25
-0.09
0.10
0.10
0.07
0.09
-0.04
0.04
-0.03
-0.02
-0.22
-0.25
-0.25
-0.09
0.06
0.04
0.06
0.04
0.29
0.44
0.27
-0.02
0.05
0.09
0.07
0.02
-0.01
-0.08
-0.03
0.11
0.04
0.15
0.05
0.02
0.07
0.09
0.08
0.02
-0.07
-0.12
-0.08
0.07
-0.25
-0.29
-0.23
0.05
0.02
0.01
0.00
0.00
-0.13
-0.15
-1.02
-6.14
0.08
1.73
0.04
-1.54
-0.14
5.58
1.93
-0.07
-0.63
-0.98
-1.53
-5.95
0.06
1.77
0.04
-1.63
-0.05
5.74
1.99
-0.13
Max
0.13
0.14
0.21
0.90
0.12
0.08
0.69
0.62
0.14
1.52
0.58
0.06
5.32
4.48
4.86
0.22
0.33
0.25
0.10
0.04
0.25
0.06
0.44
0.09
0.11
0.15
0.09
0.12
0.29
0.02
6.14
1.73
5.58
5.95
1.77
5.74
40
2. OA – Vibroacoustics
• Orbiter SEA won’t be available by PDR although it
is planned for post-PDR.
• Performed preliminary SEA on IM and SAS.
• Random specifications for components are currently
based on GEVS.
• A post-PDR SEA and FEA is expected to confirm
levels for components and instruments. (We believe
GEVS is adequate for PDR)
Greg Clarke - Mechanical Analysis
41
2. OA – Shock
• Shock analysis at the orbiter level pending.
Greg Clarke - Mechanical Analysis
42
2. OA – Future Work
•
Review/Update Mechanical Systems Specification document
with results from CLA (model submitted for preliminary run).
•
Stress Analysis: Continue sizing Orbiter primary structure.
•
Develop specific shock levels for Orbiter zones, instruments, and
components (semi-empirical/ virtual modes methods).
•
Update Orbiter FEM’s.
•
Perform Orbiter Vibroacoustic Analysis.
•
Perform Orbiter STOP Analysis.
Greg Clarke - Mechanical Analysis
43
3. IM&DA – Instrument Module
• Frequency
– Goal: 50 Hz
– FEM: 70.8 Hz
(local mode)
• 1st Primary Mode: 80.9 Hz
– X Axis Translational Mode
• 30% mass participation in X
– Max Strain Energy
• Ribs supporting the lower
LAMP fastener location
Undeformed Model
1st Primary Mode
Greg Clarke - Mechanical Analysis
• 2nd Primary Mode: 101.4 Hz
– X Axis Bending Mode
• 29% mass participation in Z
• 37% mass participation about X
– Max Strain Energy
• Facesheet near the lower LOLA
Mounting Pad.
2nd Primary Mode
44
3. IM&DA – HGAS
Reflector
FEM completed
• Model scaled to current mass allocation of 44.7 kg
• Gimbals modeled as 6 DOF springs
– Gimbal stiffness values same as for solar array
gimbals
Gimbal
Gimbal
• Model constrained at 2 latch locations, 3 restraint
locations and 4 orbiter attach points
• Frequency Goals:
Stowed: > 35 Hz
Deployed: > 1 Hz
23
3
23
3
23
Boom
(Next Step: run the model)
123456
(x4)
Hinge
Greg Clarke - Mechanical Analysis
45
3. IM&DA – SAS Modal Frequencies (Stowed)
1st Mode (34.78Hz) , Goal 35 Hz
–
–
–
2nd Mode (37.04Hz)
Outside Panel bending
7% effective modal mass, Y-axis
20% effective modal inertia, Rz
–
–
–
X
Y
Inner Panels bending
3% effective modal mass, Y-axis
10% effective modal inertia, Rz
X
Y
Greg Clarke - Mechanical Analysis
46
3. IM&DA – SAS Modal Frequencies (Deployed)
1st Mode (0.61Hz), Goal 1 Hz
–
–
–
2nd Mode (0.72Hz)
Az gimbal bending, Rx
40% effective modal mass, Z-axis
89% effective modal inertia, Rx
–
–
–
Y
Az gimbal bending, Rz
61% effective modal mass, X-axis
91% effective modal inertia, Rz
X
Z
Y
Greg Clarke - Mechanical Analysis
47
3. IM&DA – IM STOP
•
•
•
•
•
Optical Bench temperatures were provided for 14 thermal load steps. (Hot
Steady State, Cold Steady State, 10 Cold Transient & 2 Hot Transient)
Run on the fixed base FEM of the optical bench
Preliminary local distortions determined for LAMP, LOLA & LROC
Uncertainty Factor of 2.0 applied to the Delta Rotations
Results:
– All margins are positive for the steady state and hot transient cases
– Negative margins for rotations about X during the cold transient cases
• LOLA wrt side mounted star tracker (-0.10)
• LROC NAC +X wrt side mounted star tracker (-0.12)
• LROC NAC –X wrt side mounted star tracker (-0.05)
– Margins are expected to improve with the optimization of the structure as
well as thermal designs.
Greg Clarke - Mechanical Analysis
48
3. IM&DA – Vibroacoustics
•
Preliminary Analysis levels for instruments mounted on the IM are higher ( up to
10 dB in the mid range) than currently proposed (GEVS).
•
These levels are believed to be conservative and have not yet been flown down to
the instruments.
•
Plan to perform FEA pressure analysis on the IM and assess the situation.
Greg Clarke - Mechanical Analysis
49
3. IM&DA – Shock
• Addressed in a previous section (OA).
Greg Clarke - Mechanical Analysis
50
3. IM&DA – Future Work
• IM
–
Trades to optimize the structure
− Detailed bonded joint analysis
– Detailed flexure analysis
– Complete Vibroacoustic Analysis
•
HGAS
– Refine FEM to include flexibility of latch flex mounts and restraint
towers
– Determine stowed and deployed HGAS fundamental frequencies
– Determine gimbal loading due to launch loads
– Perform trade studies to size and optimize design
– Perform detailed stress analysis
– Perform thermal distortion analysis
Greg Clarke - Mechanical Analysis
51
3. IM&DA – Future Work cont
SAS
•
•
•
•
•
Static Loads Analysis
– Quasi-static Launch Loads
– Transportation and Test loads
– Thermal Distortion (stowed gradients)
Dynamic Loads Analysis
– Launch - Base Drive, Acoustics, Coupled Loads
– On-orbit – Deployment to stop, main engine cut-off, ACS maneuvers
– Shock
Detailed Stress analysis of SAS hardware:
– Gimbal Brackets
– Interface hardware
– Panel Hinges
– Deployment Latches
o Need to determine design preloads for launch
– Inserts
Detailed SA panel stress analysis
Complete Vibroacoustic Analysis
Greg Clarke - Mechanical Analysis
52
4. CLA – Coupled Loads Analysis Plan
Model
Delivery
Date
Preliminary
Design Load
Cycles
(PDLC)
Final Design
Load Cycle
(FDLC)
Verification
Load Cycle
(VLC)
Feb.’06
Through
April ‘06
April ’06
May‘08
Date
Results
Expected
Feb.’06
Through
April ‘06
June ‘06
July ‘08
Use of Results
Identify problems and preliminary structural
design. Preparation for Peer Review.
Confirm post-PDLC changes are good; facilitate
the definition of verification parameters for
Modal, Sine Vibration, and Random Vibration
Testing ( e.g identify target modes, notching
limits). Preparation for CDR.
Verification of math models.
Verify Orbiter has positive margins.
Greg Clarke - Mechanical Analysis
53
4. CLA – Coupled Loads Analysis Plan cont’
• Load updates will be :
– Placed in Mechanical Systems Specification document.
– Forwarded to subsystem, instrument and component
designers/providers.
Greg Clarke - Mechanical Analysis
54
Backup Slides
IM Backup Slides
Greg Clarke - Mechanical Analysis
56
Instrument Module
• FEM
– Mass Breakdown: Total Mass: 151.38 lbs (68.66 kg)
• Instruments & Components:
–
–
–
–
–
–
–
LAMP:
11.68 lbs (5.30 kg)
LOLA & LOLA Radiator: 25.77 lbs (11.69 kg)
LOLA Electronics
7.98 lbs (3.62 kg)
LROC NACs (2)
16.29 lbs (7.39 kg) *includes radiator mass
LROC WAC
1.46 lbs (0.66 kg)
LROC Electronics
2.34 lbs (1.06 kg)
Star Tracker (2)
8.25 lbs (3.74 kg)
» TRA Values Given
• Structure:
–
–
–
–
98.31 lbs (44.59 kg)
Subtotal: 53.07 lbs (24.07 kg)
Laminate Elements:
Flexure Beam Elements:
Ti Insert Solid Elements:
Non-Structural Mass:
30.94 lbs (14.04 kg)
0.58 lbs (0.26 kg)
7.63 lbs (3.46 kg)
13.92 lbs (6.31 kg)
Greg Clarke - Mechanical Analysis
57
Instrument Module
•
Optical Bench Structure
– Baseline Material: Composite M55J fibers, 954-2 matrix, Q-I layup is
[0,45,90,135]SN
– Baseline Adhesive: 9309NA
– Baseline Insert Material: Ti-6Al-4V
– 16 gs in any direction, from the LRO MSS
– Minimum Margin: TBD
•
Flexure Preliminary Analysis
–
–
–
–
•
Cruciform & 2 Blades for a three point interface
Baseline Material: Ti-6Al-4V
30 degC (need to verify) thermal difference between interfaces
Minimum Margin: TBD
Future Work
– Trades to optimize the structure
– Detailed bonded joint analysis
– Detailed flexure analysis
Greg Clarke - Mechanical Analysis
58
Instrument Module
• Modal Effective Mass Summary Table
MODE NO.
FREQUENCY
(Hz)
1
70.78
2
75.37
3
80.90
4
98.74
5
101.42
6
105.06
7
107.55
8
112.69
9
117.40
10
122.94
TOTAL
Rigid Body
T1
T2
T3
R1
R2
R3
0.3734
1.7960
2467.0 16587.2
8979.3
10.3039
0.2%
1.2%
1.1%
3.3%
1.3%
6.8%
4.3843
0.3458 10.8112 15073.0 54509.6
2636.7
2.9%
0.2%
0.4%
7.2%
6.8%
10.8%
7.3795
1.6457
1255.6
2483.1 20821.0
45.2603
4.9%
1.1%
0.6%
0.5%
3.0%
30.0%
2.7001 11.6650 22302.5 28881.0 59226.2
17.6713
1.8%
11.7%
7.7%
10.1%
5.7%
8.6%
0.2235
2.5364 43.3622 81121.2 94178.5
7063.7
0.1%
1.7%
1.0%
28.7%
36.8%
18.7%
0.7328
2.9655
0.0078
33.7
13.1 11521.5
0.5%
2.0%
0.0%
0.0%
0.0%
1.7%
2.3316
0.9624
0.0509
211.4
3459.2
9716.1
1.5%
0.6%
0.0%
0.1%
0.7%
1.4%
3.1556
0.2348
0.0068
0.0
294.0 12382.3
2.1%
0.2%
0.0%
0.0%
0.1%
1.8%
0.8252
3228.9
2315.8
6629.0
19.0659 20.0099
0.5%
1.5%
0.5%
1.0%
12.6%
13.2%
2.7319
2.3878 32.5338 57943.5 113228.3 20083.2
1.8%
1.6%
2.9%
21.5%
26.3%
22.5%
105.8610 39.8956 102.7046 183636.7 315949.8 159058.9
151.0908 151.0908 151.0908 220159.1 503256.6 686386.3
Greg Clarke - Mechanical Analysis
59
Instrument Module
• FEM
29.2 in
47.1 in
14.5 in
Greg Clarke - Mechanical Analysis
60
HGAS Backup Slides
Greg Clarke - Mechanical Analysis
61
SA Backup Slides
Greg Clarke - Mechanical Analysis
62
Modal Effective Mass (Stowed)
Mode
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
Freq (Hz)
34.78
37.05
37.76
44.30
48.61
50.81
59.25
60.68
63.17
64.15
65.31
76.78
83.76
85.31
87.70
Totals
X-Dir
wt (lb)
%
1.0
0.6
1.3
0.7
0.0
0.0
0.1
0.1
30.0
16.7
71.4
39.8
0.6
0.3
0.6
0.4
18.8
10.5
0.6
0.3
25.1
14.0
0.0
0.0
0.1
0.0
0.0
0.0
0.7
0.4
150.2
83.8
Y-Dir
wt (lb)
12.1
6.0
0.0
0.0
0.0
1.6
2.7
0.0
0.2
0.0
0.5
0.0
0.0
0.0
4.2
27.5
%
6.8
3.3
0.0
0.0
0.0
0.9
1.5
0.0
0.1
0.0
0.3
0.0
0.0
0.0
2.3
Z-Dir
wt (lb)
%
0.0
0.0
0.0
0.0
0.0
0.0
0.6
0.3
22.7
12.7
12.7
7.1
0.5
0.3
0.4
0.2
0.0
0.0
0.0
0.0
0.1
0.1
28.3
15.8
0.1
0.1
43.4
24.2
22.0
12.3
15.3
131.0
73.1
Rx
Ix (lb-in2)
3
5
836
5,625
42,419
24,331
1,011
23
150
471
209
72,517
192
89,573
41,114
%
0.0
0.0
0.2
1.5
11.3
6.5
0.3
0.0
0.0
0.1
0.1
19.3
0.1
23.8
10.9
Ry
Iy (lb-in2)
1
0
31
4,653
85
60
103
73
809
260
131
208,115
1,172
340,638
166,204
%
0.0
0.0
0.0
0.5
0.0
0.0
0.0
0.0
0.1
0.0
0.0
23.5
0.1
38.4
18.8
Rz
Iz (lb-in2)
237,566
115,129
371
111
51,200
117,054
38
1,105
45,617
1,764
30,534
173
1,544
129
17,455
%
20.3
9.8
0.0
0.0
4.4
10.0
0.0
0.1
3.9
0.2
2.6
0.0
0.1
0.0
1.5
278,478
74.1
722,336
81.5
619,788
52.8
Total Mass = 179.28 lbs
Greg Clarke - Mechanical Analysis
63
Modal Effective Mass (Deployed)
Mode
1
2
3
4
5
Freq (Hz)
0.61
0.72
0.76
3.64
5.96
Totals
X-Dir
wt (lb)
%
0.0
0.0
105.7
60.5
0.0
0.0
0.0
0.0
17.8
10.2
123.4
70.6
Y-Dir
wt (lb)
0.0
0.1
0.0
0.0
0.2
0.4
%
0.0
0.1
0.0
0.0
0.1
0.2
Z-Dir
wt (lb)
%
70.0
40.1
0.0
0.0
1.5
0.9
0.5
0.3
0.0
0.0
72.1
41.3
Rx
Ix (lb-in2)
1,679,909
1
29,623
1,155
0
1,710,688
%
89.1
0.0
1.6
0.1
0.0
Ry
Iy (lb-in2)
23,636
1
274,374
1,994
2
90.7
300,006
%
7.5
0.0
86.8
0.6
0.0
Rz
Iz (lb-in2)
1
1,463,857
6
4
31,354
%
0.0
90.8
0.0
0.0
1.9
94.9
1,495,221
92.8
Total Mass = 174.76 lbs (does not include latch flexure weight)
Greg Clarke - Mechanical Analysis
64
Modal Effective Mass (Stowed)
Rigid gimbal brackets
Mode
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
Freq (Hz)
34.87
37.11
37.77
44.52
56.04
60.18
62.10
62.51
64.14
65.07
77.20
84.05
86.06
91.35
97.28
Totals
X-Dir
wt (lb)
%
0.4
0.2
0.6
0.3
0.0
0.0
0.0
0.0
65.0
36.3
0.2
0.1
17.4
9.7
37.1
20.7
0.6
0.3
26.7
14.9
0.0
0.0
0.3
0.2
0.1
0.0
0.5
0.3
0.0
0.0
148.9
83.1
Y-Dir
wt (lb)
%
11.9
6.6
6.5
3.6
0.0
0.0
0.0
0.0
2.2
1.2
0.0
0.0
1.0
0.6
0.6
0.3
0.0
0.0
0.7
0.4
0.0
0.0
0.3
0.2
0.0
0.0
72.6
40.5
0.1
0.0
95.9
53.5
Z-Dir
wt (lb)
%
0.0
0.0
0.0
0.0
0.0
0.0
0.1
0.0
0.0
0.0
6.3
3.5
15.4
8.6
11.1
6.2
0.0
0.0
0.0
0.0
25.2
14.1
0.6
0.3
69.1
38.5
0.0
0.0
1.2
0.7
128.9
71.9
Rx
Ix (lb-in2)
1
6
545
975
0
18,405
28,435
20,707
1,084
21
63,887
1,117
138,710
1
2,093
%
0.0
0.0
0.1
0.3
0.0
4.9
7.6
5.5
0.3
0.0
17.0
0.3
36.9
0.0
0.6
Ry
Iy (lb-in2)
0
0
27
4,972
94
167
265
123
197
2
209,884
3,638
512,394
2
9,832
%
0.0
0.0
0.0
0.6
0.0
0.0
0.0
0.0
0.0
0.0
23.7
0.4
57.8
0.0
1.1
Rz
Iz (lb-in2)
216,528
111,267
420
0
136,247
645
35,820
73,907
1,164
31,290
8
4,108
335
292,605
80
%
18.5
9.5
0.0
0.0
11.6
0.1
3.1
6.3
0.1
2.7
0.0
0.4
0.0
24.9
0.0
275,989
73.4
741,596
83.7
904,424
77.1
Total Mass = 179.28 lbs
Greg Clarke - Mechanical Analysis
65
Modal Effective Mass (Deployed)
Rigid gimbal brackets
Mode
1
2
3
4
5
Freq (Hz)
1.40
1.57
1.99
4.78
6.10
Totals
X-Dir
wt (lb)
%
0.0
0.0
103.2
59.1
0.0
0.0
0.0
0.0
16.4
9.4
119.6
68.4
Y-Dir
wt (lb)
0.0
0.1
0.0
0.0
0.1
0.2
%
0.0
0.0
0.0
0.0
0.0
0.1
Z-Dir
wt (lb)
%
0.0
0.0
0.0
0.0
76.6
43.8
0.0
0.0
0.0
0.0
76.7
43.9
Rx
Ix (lb-in2)
1,010
3
1,740,032
18
1
1,741,064
%
0.1
0.0
92.2
0.0
0.0
Ry
Iy (lb-in2)
272,639
8
5,595
22,935
158
92.3
301,335
%
86.2
0.0
1.8
7.3
0.0
Rz
Iz (lb-in2)
49
1,450,834
2
31
30,451
%
0.0
90.0
0.0
0.0
1.9
95.3
1,481,367
91.9
Total Mass = 174.76 lbs (does not include latch flexure weight)
Greg Clarke - Mechanical Analysis
66
SAS FEM Weights
• Total FEM Mass: *179.28 lb (81.31 kg)
–
–
–
–
–
–
SA Panels:
Gimbal Actuators:
Cable Wrap Mechanisms:
Latch Hdwr:
Gimbal brackets:
Wiring, Thermal, misc:
129.07 lb (58.54 kg)
19.8 lb (8.98 kg)
8.80 lb (3.99 kg)
10.0 lb (4.53 kg)
4.80 lb (2.18 kg)
6.82 lb (3.09 kg)
• Notes:
– 56.26 lb (25.51 kg) of total weight is made up of NSM
– Total does not include approx. 15.4 lb (7.0 kg) Gimbal Controller
mass on SC
Greg Clarke - Mechanical Analysis
67
Gimbal Stiffness
Bearing Spring Stiffness
Low Stiffness
High Stiffness
Direction (lb/in or in-lb/rad) (lb/in or in-lb/rad)
Axial
3.60E+05
4.45E+05
Lateral
2.83E+05
4.30E+05
Bending
1.30E+06
2.00E+06
Torsion
1.00E+05
4.25E+05
Note: For Launch, torsion is released
Greg Clarke - Mechanical Analysis
68
IM&DA–Vibroacoustics Models
•
•
•
Purpose: Analyze the interaction of the
launch vehicle acoustic environment with
the LRO structure to determine random
vibration specifications for components.
Input: Enveloped Atlas V-Delta IV acoustic
environment
Modeling Assumptions
–
Damping:
•
•
Instrument Module
–
–
2% with 250 Hz pivot frequency for metal
parts
0.5% for composite parts
Absorption: 2% in all acoustic cavities
Model simplified to provide largest possible
subsystems. (Large subsystems have more
accurate results, due to statistical nature of
the analysis.)
Solar Array
Greg Clarke - Mechanical Analysis
69
IM&DA-IM SEA Results
LRO IM Component I/F Qual Levels from Acoustics
LROC Electronics
LOLA 1
LROC NAS +X
LAMP 2
LOLA 2
GEVS Qual
Startracker +X
Startracker -X
LAMP and LOLA Electronics
LROC NAS -X
100
10
ASD(g^2/Hz)
1
0.1
0.01
0.001
0.0001
10
100
1000
10000
Frequency(Hz)
Greg Clarke - Mechanical Analysis
70
IM&DA-Solar Array SEA Results
Solar Array Wing Peak Reponses to Atlas V-Delta IV Qual Level Input
Inner Wing
Middle Wing
Outer Wing
100
ASD(g^2/Hz)
10
1
0.1
0.01
10
100
1000
10000
Frequency(Hz)
Greg Clarke - Mechanical Analysis
71
2. OA – Backup Slides
Greg Clarke - Mechanical Analysis
72
Orbiter Finite Element Model
•
•
•
•
•
•
•
•
•
Finite element model (FEM) created in FEMAP and analyzed using MSC NASTRAN
FEM usage
– Determine loads for detailed stress analysis (basedrive, CLA)
– Analytical verification of fundamental frequency requirements
– Preliminary sizing of structural members
– Dynamic and static displacement checks
– Future derivation of instrument/component sine and random vibration levels
Total mass: 2000 kg (4409.24 lbm)
The (XYZ)orb center-of-mass location for the Orbiter with respect to the center of the interface (separation
plane) is: [38.12495, -3.532522, 0.459009] in
For all loads and normal modes runs the observatory dry mass was scaled up to obtain the observatory
2000 kg (4409.24 lbm) maximum allocation
– Scaling was done in NASTRAN using the PARAM, WTMASS card
– Propulsion tanks, fuel, IM, SAS, HGAS, instruments, components were reduced such that scaling
brought those masses back to total resource allocation
– The scale factor was 1.29
Components and propulsion tanks modeled using lumped masses (CONM2) and rigid elements (RBE3)
Instrument models
– Smeared mass: Mini-RF
– Lumped mass: Crater, LAMP, LEND, LOLA LROC NAC, LROC WAC
– Physical model: Diviner
51365 Elements
48029 Grids
Greg Clarke - Mechanical Analysis
73
Orbiter Finite Element Model
•
•
•
•
Assumptions
– Instruments/components with fundamental modes greater than 75 Hz assumed rigid (RBE3)
– Smeared non-structural mass captures items not physically modeled (blankets, cable harness,
fillets, inserts, etc.)
Interfaces:
– FEM constrained in all DOF at launch vehicle interface
– Base of IM flexures connected to spacecraft using rigid elements (RBE2) connected in all
DOF
– SAS & HGAS connected to spacecraft using rigid elements (RBE2) connected in all DOF
with appropriate releases modeled on SAS and HGAS side of interface
– Panel-to-panel connections made with common grid in all degrees of freedom (DOF)
– PM deck and stub skirt connections modeled using rigid elements (RBE2) at coincident grids
at fastener locations. RBE connections are in all DOF
Materials
– All spacecraft structure modeled using Al 6061-T651 and Al 3/16-5056-.001 3.1 pcf core
where applicable
– Honeycomb panels modeled with laminate element type (PCOMP)
– PM deck modeled using plate (PSHELL) and solid elements (PSOLID) to reflect load path
Validity Checks
– Unit gravity loading - passed
– Free-free - passed
– Ground check – passed
Greg Clarke - Mechanical Analysis
74
Orbiter Finite Element Model
Crater
IM
SAS
HGAS
Diviner
LEND
Greg Clarke - Mechanical Analysis
75
Orbiter Finite Element Model
Propulsion
Tank
Propulsion
Tank
Avionics
Module
Greg Clarke - Mechanical Analysis
76
Orbiter Normal Modes
•
Modal effective weights – relative to the launch vehicle interface plane as
percentage of total mass
MODE FREQUENCY
(Hz)
X-WT. (%) Y-WT. (%) Z-WT. (%) I-XX (%)
I-YY (%)
#
0.81%
0.20%
0.08%
0.52%
1
21.172
44.77%
2.41%
0.00%
0.28%
92.92%
2
22.303
50.47%
3
29.173
0.01%
0.11%
0.04%
0.32%
0.03%
0.04%
0.00%
0.00%
1.16%
4
30.698
5.30%
5
31.107
3.06%
0.24%
1.20%
0.02%
0.52%
6
31.537
0.00%
0.00%
0.76%
0.00%
0.27%
4.38%
0.00%
0.03%
0.01%
7
32.958
5.23%
8
33.658
0.23%
0.49%
0.01%
1.47%
0.00%
9
34.678
0.00%
1.52%
0.00%
0.08%
0.00%
10
36.775
0.18%
0.71%
0.00%
2.23%
0.00%
11
37.026
0.00%
0.83%
0.00%
0.36%
0.00%
12
37.736
0.00%
0.00%
0.00%
0.12%
0.00%
0.14%
0.38%
0.33%
0.01%
13
40.591
20.04%
14
43.551
0.09%
0.03%
0.00%
0.13%
0.01%
15
43.915
0.49%
0.01%
0.00%
2.64%
0.09%
16
44.332
0.00%
0.00%
0.00%
0.01%
0.00%
17
45.621
1.41%
0.54%
0.17%
1.65%
0.11%
18
46.258
2.12%
0.05%
0.35%
0.03%
0.10%
19
49.820
2.33%
1.14%
1.17%
0.29%
0.28%
0.01%
0.48%
1.33%
0.02%
20
50.373
6.03%
21
52.727
1.30%
0.41%
0.00%
0.01%
0.00%
22
53.411
0.36%
0.05%
0.03%
0.08%
0.00%
23
53.717
0.00%
0.03%
0.00%
0.03%
0.00%
24
54.583
0.04%
0.21%
0.00%
0.17%
0.03%
25
56.842
2.80%
0.02%
0.04%
0.06%
0.00%
26
57.632
2.26%
0.44%
0.02%
0.02%
0.00%
0.00%
0.02%
0.14%
0.00%
27
58.190
14.56%
0.07%
0.03%
0.84%
0.00%
28
58.438
8.41%
I-ZZ (%)
91.64%
0.42%
0.11%
0.00%
0.05%
0.00%
1.03%
0.23%
0.61%
0.25%
0.30%
0.00%
0.00%
0.00%
0.03%
0.00%
0.15%
0.00%
0.04%
0.01%
0.01%
0.00%
0.00%
0.00%
0.00%
0.01%
0.05%
0.01%
Greg Clarke - Mechanical Analysis
• EELV 10 Hz lateral and 27 Hz
axial requirements have been
met with some margin.
77
Orbiter Normal Modes
•
•
Mode 1 – 21.17 Hz
Lateral Bending: RZ
•
•
Mode 2 – 22.30 Hz
Lateral Bending: RY
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
78
Orbiter Normal Modes
•
•
Mode 4 – 30.70 Hz
Torsion: RX
•
•
Mode 7 – 32.96 Hz
Thrust: X
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
79
Orbiter Normal Modes
•
•
Mode 13 – 40.59 Hz
Torsion: RX
•
•
Mode 20 – 50.37 Hz
Torsion: RX
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
80
Orbiter Normal Modes
•
•
•
•
Mode 27 – 58.19 Hz
Thrust: X
Mode 28 – 58.44 Hz
Thrust: X
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
81
Orbiter FEM – SAS Deployed
•
•
•
•
The purpose of the deployed FEM is to verify the on-orbit frequency requirements
The SAS is expected to exhibit the lowest deployed frequency.
Deployed FEM identical to stowed FEM except the SAS is deployed
Free-free boundary conditions
MODE FREQUENCY
#
Rigid Body Modes
Exceeds on-orbit requirement of 0.5 Hz
Greg Clarke - Mechanical Analysis
1
2
3
4
5
6
7
8
9
10
(Hz)
0.000
0.000
0.000
0.000
0.000
0.001
0.782
0.844
0.894
3.615
82
Orbiter Normal Modes – SAS Deployed
•
•
Mode 1 – 0.782 Hz
SAS Torsion
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
83
Orbiter Normal Modes – SAS Deployed
•
•
Mode 2 – 0.844 Hz
SAS Bending RZ
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
84
Orbiter Normal Modes – SAS Deployed
•
•
Mode 3 – 0.894 Hz
SAS Bending RX
Deformed model. Total translation contour displayed for both plots
Greg Clarke - Mechanical Analysis
85
Orbiter Basedrive Analysis
•
•
•
•
•
The purpose is to investigate dynamic interactions of orbiter, assess quasistatic limit loads, and estimate interface loadings (SAS, HGAS, etc.) prior
to coupled loads analysis
FEM constrained at launch vehicle interface
Analysis was performed to 60 Hz to account for dynamic uncertainty
(DUF = 1.2)
Envelope of Delta IV and Atlas V PPG sine input specifications
Force limited – envelope of PPG net CG limit accelerations and orbiter
mass and CG used to develop limits for interface forces
Greg Clarke - Mechanical Analysis
86
Orbiter Basedrive Analysis
Input limited at modes 1 and 2 in lateral directions
LRO Basedrive Limited Input Levels
1.2
1
Input Acceleration, g
•
0.8
X
Y
Z
0.6
0.4
0.2
0
0
10
20
30
40
Frequency, Hz
Greg Clarke - Mechanical Analysis
50
60
70
87
Orbiter Basedrive Analysis
Results show that all instrument and component responses enveloped by MAC
Only flexible, massive subsystems and components have exceedences
Reductions to current levels (MAC) likely
LRO Basedrive Results Summary: Net CG Limit Loads
Instruments, & Components < 100 kg
100
Limit Load (g)
•
•
•
10
MAC
1
0.1
1
10
100
Mass (kg)
Greg Clarke - Mechanical Analysis
88
Orbiter Basedrive Analysis
•
•
•
•
•
Summary table shows current
estimate of limit level loads
Basedrive responses above MAC
levels ignored
MAC limit load estimated using
CBE mass
All structure must survive the
maximum PPG static load (6.6 g)
at a minimum
Results valid for current
configuration – modifications to
the current configuration may
cause changes in limit loads
Item
IM Net CG
HGAS Net CG
SAS Net CG
Diviner Net CG
Propulsion Tank (2)
Pressurant Tank
Mini-RF Ebox
Mini-RF Xpander
Battery Box
Reaction Wheel (2)
C&DH
PSE
S-Band Transponder
PDE
TWTA
Ka Modulator
EPC
SA & HGA Gimbal Controller
IMU
Crater
Lamp
LEND
LOLA
LOLA Electronics Box
LROC NAC -X
LROC WAC
LROC Electronics Box
Star Tracker +X
SAS El-Gimbal Motor
Greg Clarke - Mechanical Analysis
Summary
CBE Mass
Expected Limit
lbm
kg
Load (g)
ID
Rationale
NET 151.51 68.71
14.9 Basedrive
NET
68.69 31.15
19.7
MAC
NET 151.70 68.80
15.1
MAC
NET
22.05 10.00
19.9 Basedrive
1002 1057.96 479.80
13.4
MAC
1005
33.96 15.40
23.9 Basedrive
1015
8.60
3.90
6.6
Static
1016
1.94
0.88
6.9 Basedrive
2015
66.15 30.00
9.9 Basedrive
2017
26.46 12.00
15.7 Basedrive
2020
46.75 21.20
8.9 Basedrive
2023
32.85 14.90
12.5 Basedrive
2024
11.03
5.00
10.0 Basedrive
2026
30.87 14.00
9.8 Basedrive
2027
2.21
1.00
6.9 Basedrive
2028
8.82
4.00
6.6
Static
2029
2.21
1.00
11.2 Basedrive
2030
15.44
7.00
10.7 Basedrive
2031
9.92
4.50
10.0 Basedrive
3027
12.35
5.60
7.5 Basedrive
3029
11.03
5.00
17.7 Basedrive
3030
52.26 23.70
11.0 Basedrive
3031
16.76
7.60
15.7 Basedrive
3032
6.62
3.00
14.4 Basedrive
3034
14.33
6.50
17.2 Basedrive
3035
1.54
0.70
22.4 Basedrive
3036
1.54
0.70
12.4 Basedrive
3037
6.62
3.00
17.6 Basedrive
162001
11.91
5.4
25.0 Basedrive
89
LRO PDR
Mechanical Programmatics/Summary
Giulio Rosanova
February 8, 2006
LRO Mechanical Summary Agenda
•
Programmatics (G. Rosanova)
– Development Flow
– Make/Buy
– Risks
– Safety
– Schedule
– Summary
Giulio Rosanova - Mechanical Programmatics/Summary
2
PM Structure (ETU & FLT) Development
PM
Design/Analysis
PM Drawing
Package
PM (FLT)
Structure Fabrication
PM Contract
Award
PM (FLT)
Assembly
TO LRO FLT
ASSEMBLY /
BUILD-UP
PM Material
Procurement
PM (ETU)
Structure Fabrication
TO LRO ETU
ASSEMBLY /
BUILD-UP
PM (ETU)
Assembly
PM Insert
Proof Tests
Giulio Rosanova - Mechanical Programmatics/Summary
3
AM Structure (ETU & FLT) Development
AM
Design/Analysis
AM Contract
Award
AM Drawing
Package
AM (FLT)
Structure Fabrication
AM (FLT)
Assembly
TO LRO FLT
ASSEMBLY /
BUILD-UP
AM Material
Procurement
AM (ETU)
Structure Fabrication
TO LRO ETU
ASSEMBLY /
BUILD-UP
AM (ETU)
Assembly
AM Insert
Proof Tests
Giulio Rosanova - Mechanical Programmatics/Summary
4
IM Structure (FLT Only) Development/Proof
IM Drawing
Package
IM
Design/Analysis
IM Contract Award
IM (FLT)
Structure Fabrication
IM (FLT)
Assembly
IM Material
Procurement
IM MASS
SIMULATORS
IM Pre-Test
Alignment
IM Proof
Test
IM Post-Test
Alignment
TO LRO FLT
ASSEMBLY /
BUILD-UP
TO LRO ETU
ASSEMBLY /
BUILD-UP
IM Insert
Proof Tests
Giulio Rosanova - Mechanical Programmatics/Summary
5
HGADS (ETU & FLT) Development
HGADS Design &
Analysis
HGADS Drawing
Package
HGADS FAB
Contract
AWARD
HGADS (ETU)
Mechanisms
Fabrication
HGADS (ETU)
Assembly
TO LRO ETU
ASSEMBLY /
BUILD-UP
HGADS (FLT)
Mechanisms
FabricationT
HGADS (FLT)
Assembly
TO LRO FLT
ASSEMBLY /
BUILD-UP
Giulio Rosanova - Mechanical Programmatics/Summary
TO LRO ETU
ASSEMBLY /
BUILD-UP
6
SADS (ETU & FLT) Development
SADS Design &
Analysis
SADS (ETU)
Mechanisms
Fabrication
SADS Drawing
Package
SADS FAB
Contract
AWARD
SADS (ETU)
Assembly
TO LRO FLT
ASSEMBLY /
BUILD-UP
TO LRO ETU
ASSEMBLY /
BUILD-UP
SA
SUBSTRATE
(ETU)
SADS (FLT)
Mechanisms
FabricationT
SADS (FLT)
Assembly
Giulio Rosanova - Mechanical Programmatics/Summary
TO SAS
DEVELOPMENT
7
SA Substrate (ETU & FLT) Development
SA SUBSTRATE
Design & Analysis
SA SUBSTRATE
Drawing Package
SA SUBSTRATE FAB
Contract AWARD
MODULE FRAME
(ETU) Fabrication
MODULE SUBSTRATE
(ETU & FLT) Fabrication
DELIVER (6) MODULE
SUBSTRATES & FRAME
(QUAL /TEST) to POWER
GROUP
POPULATE ETU
MODULE SUBSTRATE
w/ S/A CELL
SIMULATORS
BOND ETU MODULE
SUBSTRATE to ETU
FRAMES
TO SADS
ASSY (ETU)
DELIVER (85) MODULE
SUBSTRATES (FLT) to
POWER GROUP
MODULE FRAME
(FLT) Fabrication
TO FLT SAS
DEVELOPMENT
Giulio Rosanova - Mechanical Programmatics/Summary
8
Common Deployables Components
Dampers,
Bearings, &
Potentiometer
Procurement
HGAS Tooling
(Design & Fab.)
TO ETU HGADS
& ETU SADS
ASSEMBLY
SAS Tooling
(Design & Fab.)
Pyro Actuator
Contract Award
Pyro
Actuator
Procurement
Giulio Rosanova - Mechanical Programmatics/Summary
9
FLT Primary Structure Assembly / Build-Up
FLT IM
FLT PM
FLT AM
DELIVER FLT
PM to I&T
LRO FLT
PRIMARY
STRUCTURE
ASSEMBLY
DELIVER FLT
AM to I&T
ETU HGADS
ETU SADS
DEPLOYABLES
ASSEMBLY to
FLT PRIMARY
STRUCTURE
FLT IM to
ETU
ASSEMBLY/
BUILD-UP
Giulio Rosanova - Mechanical Programmatics/Summary
ETU HGADS
to ETU
ASSEMBLY/
BUILD-UP
ETU SADS to
ETU
ASSEMBLY/
BUILD-UP
10
ETU Primary Structure Assembly / Build-Up
FLT IM
ETU PM
ETU AM
INSTALL MASS
SIMULATORS, SG, &
ACCELERAOMETERS
ETU HGADS
ETU SADS
DEPLOYABLES
ASSEMBLY to
ETU PRIMARY
STRUCTURE
LRO ETU
PRIMARY
STRUCTURE
ASSEMBLY
DEPLOYABLES
FUNCTIONAL TEST
(POP & CATCH)
TO MECH
QUAL.
TESTING
INSTRUMENT & S/C
COMPONENT MASS
SIMULATORS
Giulio Rosanova - Mechanical Programmatics/Summary
11
Mechanical Primary Structure Qualification Testing
MASS
PROPERTIES
(1 Axis)
MODAL
SURVEY
(On Shaker)
DELIVER IM
to I&T
STRENGTH
TEST
(Centrifuge)
DIS-ASSEMBLY
ETU PM+AM & ETU
HGADS & ETU SADS
AVAILABLE for
OFF-LINE TESTING
Giulio Rosanova - Mechanical Programmatics/Summary
12
GIMBAL Development/Qualification
Develop
GIMBAL
Spec & SOW
GIMBAL
Design &
Analysis
GIMBAL
Development
(Vendor) LIFE
UNIT-(1 SET)
GIMBAL
Mechanical
Qual. Test
GIMBAL
Development
(Vendor) FLT
UNITS-(2 SETS)
GIMBAL Life
Test
GIMBAL
Mechanical
Accept. Test
GIMBAL
Development
(Vendor) SPARE
UNIT-(1SET)
Contract
Award
TO FLT SAS &
FLT HGAS
DEVELOPMENT
GIMBAL
Mechanical
Accept. Test
Giulio Rosanova - Mechanical Programmatics/Summary
13
HGAS FLT Development/Qualification
ASSEMBLE FLT
HGADS, GIMBALS,
WG, HGA
HGAS Modal Survey
(Basedrive on Shaker)
HGAS Sine Vib
Testing
HGAS RF END to
END Testing
(Ambient)
HGAS MECH
FUNCTIONAL
Testing
HGAS Strength
Testing
HGAS Random
Testing
Thermal Vac.
Deploy/Functional
END-to-END Testing
Deliver HGAS to
I&T
Giulio Rosanova - Mechanical Programmatics/Summary
14
SAS (ETU & FLT) Development/Qualification
ASSEMBLE ETU
SADS, GIMBALS, ETU
Substrate
SAS Modal Survey
(Basedrive on Shaker)
ETU SAS Mechanical
Functional/Deployment
Testing Ambient
SAS Strength
Testing
SAS Thermal Vac.
Deployment/Functional
Testing
Receive FLT Frames
from Vendor
Receive FLT Cell
Modules from
POWER GROUP
Bond FLT Cell
Module to FLT
Frames
Panels to POWER
GROUP for Wiring &
Electrical Testing
Thermal Cycle
Panels
Panels to POWER
GROUP for
Electrical Testing
ASSEMBLE S/A WING
(FLT-SADS, GIMBALS,
FLT-SA Panels)
SAS Mechanical
Funct./Deployment
Testing Ambient
SAS Strength
Testing
Deliver SAS to I&T
Giulio Rosanova - Mechanical Programmatics/Summary
15
Make / Buy Decisions
•
Primary Structure & Secondary Structure
– Build to Print - Code 547 Planning Office
•
Gimbal System (Actuators & GCE)
– Actuator Commonality for HGAS & SAS
• Must Meet Requirements for Both Systems
– Plan - Competitive RFP
• Possibility of Leveraging SDO Contract
•
Solar Array Substrate (Modules & Frames)
– Plan - Sole Source Procurement
•
Pyro-Actuators
– Purchase Order
Giulio Rosanova - Mechanical Programmatics/Summary
16
LRO Mechanical System Risk Matrix
Rank &
Trend
5
L
I 4
K
E
L 3
I
H
O 2
O
D
2
Risk Title
Risk ID
1
↓
M
Mechanical Interface Control
Documents/Drawings
MECH-103
2
→
M
Gimbal Procurement
MECH-105
3
→
M
Primary Structure ETU
MECH-106
4
↓
M
Mechanical Mass Allocation
MECH-108
5
↓
R
EELV Selection
MECH-107
1
3
Approach
4
5
1
2
1
3
4
5
CONSEQUENCES
Criticality
L&C Trend
Approach
High
↓
Decreasing (Improving)
M – Mitigate
Med
↑
Increasing (Worsening)
W – Watch
Low
→
Unchanged
A – Accept
New since last month
R – Research
*
Giulio Rosanova - Mechanical Programmatics/Summary
17
Safety
•
S/C Structure Sub-System Hazard
– Hazard Descriptions
• Structural failure of spacecraft during launch sequence
– Worst Case Severity
• Catastrophic
– Control and Mitigation Strategy
• Stress analysis to verify margins of safety
• Structural qualification testing
• Materials selection for resistance to stress corrosion and materials compatibility
• Fracture control plan
• Exclude counterfeit fasteners
•
Deployables Sub-System Hazards
– Hazard Descriptions
• Inadvertent deployment of solar array and high gain antenna
– Worst Case Severity
• Critical
– Control and Mitigation Strategy
• Mechanical release/detention mechanisms
• Tethering during ground processing
• Three independent inhibit scheme controlled through PDE Inhibit Unit
Giulio Rosanova - Mechanical Programmatics/Summary
18
Schedule - Key Deliverables
UID
Date
5289, 5290
3/15/2006
2005
4/17/2006
4105, 4120, 4133
7/3/2006
5664,5683
5/1/2007
Primary Structure Qualification Complete
5678
9/21/2007
Deliver IM to I&T
5682
9/28/2007
Gimbal System Available
5358
7/9/2007
Solar Cell Modules Available
5702
10/1/2007
Deliver HGAS to I&T
5370
1/8/2008
Deliver SAS to I&T
5527
2/1/2008
Receive Final MICD’s (Inst., Av. Components)
Mechanical Peer Review
Primary Structure Fabrication Start
Deliver PM, AM to I&T
Giulio Rosanova - Mechanical Programmatics/Summary
19
Summary
• Preliminary Design and Analyses Indicate the Mechanical
System will Meet LRO Mission Requirements
• Our Challenges are Clearly Understood
• We are Ready to Proceed to the Detailed Design Effort
Giulio Rosanova - Mechanical Programmatics/Summary
20
Summary
•
Preliminary Design and Analyses Indicate the Mechanical System will Meet LRO
Mission Requirements
•
Our Challenges are Clearly Understood
•
We are Ready to Proceed to the Detailed Design Effort
Giulio Rosanova - Mechanical Programmatics/Summary
21
LRO PDR
Thermal
Charles Baker
February 8, 2006
Thermal Team
Team Member
Role
Charles Baker/GSFC
Thermal Lead
Bill Chang/ESS
Instrument Accommodations
Christine Cottingham/LM
AJ Mastropietro/GSFC
Deployable Thermal Analyst
Cynthia Simmons/ESS
Instrument Thermal Analyst
Spacecraft Thermal Analyst
Charles Baker - Thermal
2
Agenda
•
•
•
•
•
•
•
•
•
•
•
•
Level II to Level III Requirements
Temperature Requirements
Instrument Accommodation
Bounding Thermal Cases
Thermal Design
Thermal Analysis
Trade Studies
Mass & Power
Verification
Safety
Risk & Mitigation
Wrap-up
Charles Baker - Thermal
3
Requirements
Higher Level Rqmt
Identifier
Object Text
MRD-81 Monitoring
MRD-79 Survival Ranges
MRD-78 Operational Ranges
MRD-70 Structural Stability
MRD-67 Thermal Environments
MRD MRD-31 Power Allocations
MRD MRD-20 Spacecraft Safing
MRD-52 Sun Avoidance
MRD MRD-50 Mission Phases
MRD-4 Launch Window
MRD-21 Mission Duration
THM-1
THM-2
THM-3
THM-4
THM-5
THM-6
THM-7
THM-8
THM-9
MRD-81 Monitoring
MRD-79 Survival Ranges
MRD MRD-78 Op Ranges
MRD-70 Structural Stability
MRD-67 Thermal Environments
MRD-31 Power Allocations
MRD-28 Vehicle Interfaces
MRD-20 Spacecraft Safing
MRD-52 Sun Avoidance
MRD-50 Mission Phases
MRD-4 Launch Window
MRD-21 Mission Duration
I
A D
T
Responsi
ble
Instrument couplings are controlled in the Thermal ICDs unique to
each instrument. S/C interface temperatures are controlled in the
431-SPEC-000091. Control Heater Powers are Allocated in the 431SPEC-000091 if external to instrument power leads. Environments
and Mission Modes are specified in 431-SPEC-00091
3
2
Instrumen
t Pis, 545
THM-10
THM-11
The propulsion module shall protect the spacecraft avionics and LRO
payloads from LRO thruster fires. : The propulsion module shall be
thermally isolated from the avionics module.
2
545
MRD-79 Survival Ranges
MRD-78 Operational Ranges
THM-12
All mission critical LRO heater circuits shall be one fault
tolerant
2
545
MRD-79 Survival Ranges
MRD-78 Operational Ranges
THM-13
LRO heaters shall not exceed recommend power densities for heater
construction and application method
2
MRD-81 Monitoring
THM-14
The thermistor/PRT shall be capable of being read over all specified
temperature ranges. The sensors shall be accurate within 0.5°C from
-45 to +75°C
2
Charles Baker - Thermal
4
Baseline Thermal Design Concept
• Avionics Module coupled to
remote radiators
• Mostly isolated instruments with their own
dedicated radiators
• Coupled Instruments have dedicated S/C
radiators or coupled to avionics module
Charles Baker - Thermal
5
Instrument I/F Temperature Limits
TEMPERATURE RANGE (°C)
COMPONENT
Op I/F Limit
Surv I/F Limit
Cosmic Ray Telescope of the
Effects of Radiation
(CRaTER)
S/C at I/F to CRaTER
-30 to +25
-40 to +35
Diviner (DRLE)
S/C at I/F to Diviner Instr
-20 to +45
-35 to +60
S/C at I/F to remote electronics box
-20 to +45
-35 to +65
Lyman-Alpha Mapping Project
(LAMP)
S/C I/F at base of LAMP’s feet
-30 to +30
-40 to +40
Lunar Exploration Neutron
Detector (LEND)
S/C at I/F to LEND
-20 to +40
-30 to +50
S/C at I/F to Optics Package
-30 to +30
-60 to +40
S/C at I/F to Instrument Electronics
-30 to +40
-60 to +50
S/C at base of NAC
-30 to +30
-40 to +40
S/C I/F at base of WAC
-30 to +30
-40 to +40
S/C I/F at base of SCS
-30 to +30
-40 to +40
S/C I/F at base of antennae’s feet
-50 to +50
-60 to +60
S/C I/F at base of electronics feet
-30 to +50
-40 to +60
Lunar Orbiter Laser Altimeter
(LOLA)
Lunar Reconnaissance Orbiter
Camera (LROC)
Mini RF
Charles Baker - Thermal
6
Temperature Limits
TEMPERATURE RANGE (°C)
SUBSYSTEM
COMPONENT
Op I/F Limit
Surv I/F Limit
Power Subsystem Electronics
-10 to +40
-20 to +50
Battery
+10 to +30
+0 to +40
Star Trackers
-30 to +50
-35 to +60
Inertial Measurement Unit
-30 to +65
-35 to +75
Reaction Wheels
0 to +50
-30 to +60
Coarse Sun Sensors
-140 to +135
-140 to +145
Propulsion and Deployables
Electronics (PDE)
PDE
-10 to +40
-20 to +50
C&DH
C&DH
-10 to +40
-20 to +50
S Comm
S-band components
-10 to +50
-20 to +60
Ka Comm
Ka Baseband Modulator
-10 to +50
-20 to +60
Ka TWT
-10 to +70
-20 to +80
EPC
-10 to +50
-20 to +60
Power
Attitude Control System (ACS)
Charles Baker - Thermal
7
Temperature Limits (cont.)
TEMPERATURE RANGE (°C)
SUBSYSTEM
COMPONENT
Operational
Survival
Mechanical
Structure Propulsion Module
-50 to +70
-60 to +80
Structure -Avionics Module
-50 to +50
-60 to +60
Structure –Avionics to Propulsion
-50 to +50
-60 to +60
Structure. Instrument Module
-50 to +50
-60 to +60
High Gain Antenna (HGA) Gimbals
-10 to +50
-20 to +60
HGA Boom
-75 to +75
-85 to +85
Mechanisms
Propulsion System
HGA Release and Deploy
-10 to +50
-20 to +60
Solar Array (S/A) Gimbals
-10 to +50
-20 to +60
S/A Boom
-10 to +50
-20 to +60
S/A Release and Deploy
-10 to +50
-20 to +60
Hydrazine Tank 1
+10 to +40
N/A
Pressure Tanks
+0 to +50
N/A
High and Low Press Transducers
+10 to +40
N/A
Gas and Liquid Latch Valve
+10 to +40
N/A
Fill and Drain
+10 to +40
N/A
Gas System Filters
+0 to +50
N/A
Liquid Filters
+10 to +40
N/A
Pressure Regulators
+0 to +50
N/A
Plumbing Lines
+10 to +40
N/A
NC Pyro Valves
-40 to +40
-60 to +50
HGA
-140 to +145
-140 to +145
Wave Guide
-50 to +50
-60 to +60
S-Band Comm
TT&C Omni Antenna
-120 to +80
-130 to +90
Power
Solar Array
-125 to +135
-160 to +145
K-Band Comm
Charles Baker - Thermal
8
Instrument Model (from Instrument Teams)
Coupling
Radiator
Op Pwr
Surv Pwr
W/K
in2
W
W
Instrument
3.
N/A
N/A
N/A
Diviner
(DRLE)
Instrument
0.05
111.
6.
13.7
DREB
1.4
N/A
N/A
N/A
LAMP
Instrument
0.054
22.3
5.
8.
LEND
Instrument
4.
N/A
N/A
N/A
Optics Bench
0.05
175.
20.4
29.4
MEB
0.05
120.
2.
16.
NAC
0.033
90.
0.
19.3 x 2
WAC
0.033
39.
5.8
6.7
SCS
0.033
30.
4.75
4.75
Antennae
TBD
N/A
N/A
N/A
Transmitter
21.4
N/A
N/A
5
Control Processor
16.4
N/A
N/A
5
N/A
N/A
33
28
CRaTER
LOLA
LROC
Mini RF
Optical Bench
Instrument Requirements are known, working with them for CDR Designs
Charles Baker - Thermal
9
Lunar Thermal Design
Lunar Orbit Environment Parameters
Lunar IR Emission as a function of Beta Angle
Hot
q”IR = [(C1-C2)*cos(β)*cos(θ)] + C2
Solar
1420
W/m2
Cold
1280
W/m2
Albedo
Factor
0.13
0.06
IR (at
subsolar C1
below)
1420 x (10.06) = 1335
W/m2
1280 x (10.13) = 1114
W/m2
IR (Cold
side)
5 W/m2
5 W/m2
where:
q”IR= IR flux from Lunar surface
C1 = Peak flux at subsolar point
C2 = Minimum flux emitted from
shaded Lunar surface
β = Beta angle
θ
= Angle from subsolar point
Beta 90o
Charles Baker - Thermal
10
Design Cases Operational
LRO Orbit Parameters:
• Beta 0° is the Hot Op Case
• Most severe IR loading
• Zenith facing radiators flip through
sun
• Instrument apertures “see” sun near
dawn and dusk 74°off bore site at 70
km
• Beta 90° is the Cold Op Case
• Zenith facing radiators look at deep
space
• Minimal IR loading
Views from Sun
Beta 0o
LRO Bounding Thermal Cases:
Type
Lunar Circular
Altitude
50±20km
Inclination
90° (polar orbit)
Orbit Period
113 minutes
Full Sun Orbits Beta 90.0° to 76.4° (55 days/yr)
Eclipsed Orbits Beta 76.4° to 0.0° (310 days/yr)
Max. Eclipse
Beta 90o
48 minutes (Beta 0°)
Charles Baker - Thermal
11
Design Cases Survival
-Y Sun Pointing (Solar Inertial)
Zenith facing radiators can be edge
on to the moon with Beta 0° Hot, and
Beta 90° Cold
On Orbit Safe, Viewed from Sun
Lunar Cruise (Solar Inertial)
Same orientation as survival – no IR
loading
Lunar Eclipse (Solar Inertial)
Worst case Beta 90° start with 160
minute eclipse duration
Charles Baker - Thermal
12
Thermal Design Changes Since SRR
• Project SRR – August, 2005
• Thermal Peer Review ‘J’ Design – October, 2005
– 30 RFAs – all closed
– Add Variable Conductance
– Isothermalize Instrument Optical Bench
• SRM Thermal Design – November, 2005
– Fully Blanketed Isolated Instrument Optical Bench
– Isothermal Avionics Panel
• Baseline Configuration – December, 2005
– Maintained Positive Features from SRM design
• MLI covered Optical Bench
• Isothermal Avionics Panel
– Improved thermal by having more space to locate radiators
– Increased heater power during cruise by allowing deployment of
solar arrays
Charles Baker - Thermal
13
Passive Design versus Active (VCHP) Design
• VCHPs on the avionics and battery realized a net
saving of control heater power in the 100-300 W
range
• Isothermal heat pipe embedded panel to support the
electronics boxes
– Decoupled the radiator design is not primary structure
(later need date)
– Allows later design optimization and accommodation
• The VCHPs and Radiator will be ordered in the
Mission CDR timeframe accommodating the LRO
design evolution
Charles Baker - Thermal
14
Baseline Design
•
Isothermal Avionics Panel
– Most Avionics thermally coupled into embedded CCHP panel
– Dual bore Header Heat Pipes couple Isothermal Panel to VCHP Radiator that
is separately mounted on Zenith Surface
– Radiators may be re-sized at CDR if necessary
– Battery is maintained on a separate VCHP network to a separate radiator
•
De-Coupled Instrument Optical Bench (Low thermal distortion)
– Uses low-CTE Composite M55J
– Fully Blanketed
– Heated with low density heaters to maintain cold limit temperatures
•
CRaTER, LEND
– S/C provided radiators are thermostatically controlled
•
Prop Module
– 2 Prop Tanks and Pressurant Tank are Aluminum Taped and thermostatically
controlled
•
Deployable Gimbals – Thermostatically controlled
Charles Baker - Thermal
15
Baseline Instrument Module
¾CRaTER
¾ LEND
Isolated
Coupled with
dedicated
S/C Radiators
Isolated
S/C Panels
¾ Mini RF Antenna
¾ Diviner Instrument
Optical Bench
¾ LOLA OTA
¾ LOLA MEB
¾ LAMP
¾ LROC WAC
Isolated with ¾ Star Cameras
dedicated
S/C Radiators
Isolated
¾ LROC (NAC1, NAC2, SCS)
Isolated via flexures
S/C +Y Side (Anti-sun) of Prop Module
S/C Isothermal Heat Pipe Panel
Coupled
Charles Baker - Thermal
¾ Gyro
¾ Diviner DREB
¾ Mini-RF
Transmitter&Elec
+Y (Anti-Sun side)
-Z (Zenith)
16
Avionics & Prop Module Heat Dissipation Paths
Zenith Radiator Panel WITH Embedded HPs
Coupled with Bolt on Headers and VCHPs
AS
RW
MLI
LINEAR COUPLING
VARIABLE
CONDUCTANCE
PRESS
TANK
Avionics
Boxes
RW
AS
Instrument Module
- Y (Sunside)
SOLAR ARRAY (BETA 0 NOON)
PROP TANK
PROP TANK
Z (Nadir)
ISOTHERMAL PANEL WITH
EMBEDDED HPs
Charles Baker - Thermal
17
Heat Pipe Network – Avionics Panel
25.4W-KaTWT
17.7W-KaMod
38W-Battery
12.5W-SAcntrl
4.1W-KaEPC
31.7WS-Band
2.8W-DDA
12.5W-HGcntrl
114W-CDH
7.W-DREB
82.5W-PSE 37.5W-PDE
35W-IRU
4.4W-OSC
2.2W-PA
8.7W-MAIN
• Embedded heatpipes inside aluminum honeycomb panel draw heat from avionics
• Red externally mounted panel header heatpipe from radiator assembly brings heat
to radiator
• Entire Heat Pipe Network is 1-g testable and 1 fault tolerant
• All heatpipes have >50% QL margin
Charles Baker - Thermal
18
Heat Pipe Network - Radiator
• CCHP heatpipes embedded in aluminum
honeycomb radiator
• Radiator has Spreaders for high Fin
efficiency
• VCHPs allow -65°C Radiator in cold case
• Radiator can be sized and ordered after CDR
• Entire Heat Pipe Network is 1-g testable and
1 fault tolerant
• Externally mounted VCHP,
Header, and Avionics Transport
Pipe Assembly are attached to
the isothermal panel late in I&T
PANEL
HEADERS
(4X)
FURTHEST
VAPOR FRONT
VCHP (4X)
RADIATOR HEADER (4X)
EMBEDDED SPREADER (~10X)
AVIONICS
BOXES
Charles Baker - Thermal
19
Thermal Hardware
•
Standard Thermal Hardware
–
VCHP (Variable Conductance Heat Pipes) are standard Ammonia with NCG
(Nitrogen or Helium) – 50% QL margin, work in 1-g
– CCHP (Constant Conductance Heat Pipes) - 50% QL margin, work in 1-g
(reflux)
– 5 Software Controlled Heater Circuits will be allocated to: 1 circuit each for
the 3 Avionics Panel VCHP (remaining 3 are T-Stat controlled), 1 circuit for
LEND, 1 circuit for the Instrument Optical Bench
– Radiator Thermal Coatings are Indium Tin Oxide covered Optical Solar
Reflectors or NS43C conductive white paint
–
–
–
–
–
Heaters are Kapton Film Heaters or Vishay-Dale Ohm Heaters
MLI is 15 layer, 3 mil VDA backed Kapton outer layer
Thermistors are YSI 2.252 k ohm or 2k ohm PRT (172 read by C&DH)
Interface Materials are Chotherm – 1671 or NuSil CV-2946 or CV2948
All other Heater Services will have 2 mechanical thermostats in series
(Honeywell mechanical thermostats 701 Series)
Charles Baker - Thermal
20
Thermal Mass
EQUIPMENT LIST
Subsystem
Thermal
Components
Sensors
Heaters
Misc H/W
MLI
•
TRA
Allocation
Kg
26.8
2.4
0.4
5.2
18.9
CURRENT
BEST
ESTIMATE (kg)
22.8
2.3
0.3
4.2
16.0
Mass (kg)
17.8%
5.3%
17.6%
25.0%
17.6%
CCHPs, VCHP, Radiators, and OSR coating mass are all contained in the
mechanical allocation
Charles Baker - Thermal
21
Analysis Summary
• LRO Environmental Constants and Basic Design
Principles were reviewed at the ‘J’ Thermal Peer
Review
• A TSS and SINDA model were built of the EELV
configuration to demonstrate adequate power
margins for LRO thermal
• All assumptions are biased hot and cold
– Environment
– Power Dissipations
– Estar
Charles Baker - Thermal
22
Thermal Analysis Status
• Baseline Design Reduced 890 Nodes Geometric and Thermal
Model completed and run for PDR in bounding cases
– Heat Pipe Layout is included
– Simplified VCHP model is included
– LEND and CRaTER are full models
• Baseline Design Reduced 1900 Nodes Geometric and
Thermal Model completed and run to verify previous results
– Detailed Heat Pipe Layout, Avionics Panel and Radiator
• Detailed Optical Bench Model run and results provided to
Mechanical for Optical Bench STOP
• Full Detailed Model in progress
– Detailed Instruments and Instrument Module
Charles Baker - Thermal
23
Temperature Results - Hot
•
•
•
•
All component temperatures are showing
at least 5 C of temperature margin (versus
VCHP setpoint of +15 C)
Avionics Radiator has 20 C of margin
Battery Radiator has 30 C of margin
Allocation Dissipations
Charles Baker - Thermal
24
Temperature Results - Cold
•
•
•
All component temperatures are showing at
least 5 C of temperature margin (versus
VCHP setpoint of -12 C)
Cold CBE Dissipations Assumed
All Heaters have GEVS margin
Charles Baker - Thermal
25
Temperature Results – Survival (Beta 90)
•
•
•
All component temperatures are showing at
least 5 C of temperature margin (versus
VCHP setpoint of -12 C)
Minimum CBE Dissipations Assumed
All Heaters have GEVS margin
Cold Case
Charles Baker - Thermal
26
Instrument Transients
Hot
Cold
Charles Baker - Thermal
27
Avionics and Radiator Transients
Hot Op
Cold Op
Charles Baker - Thermal
28
Hot Dissipation with Cold Environment Raw Htr Pwr
Orbiter Power
Beta 0
Beta 50
Beta 60
Beta 70
Beta 90
Electrical Dissipation (W)
634
634
634
634
634
Thermal Dissipation (W)
684
684
684
684
684
Avionics EHB (W)
10
18
19
19
20
15
50
58
66
83
5
11
17
19
25
Propulsion Op/Surv (W)
42
80
90
96
108
Instrument Op* (W)
0
14
28
35
61
Instrument Survival* (W)
N/A
N/A
N/A
N/A
N/A
Gimbal Heaters (W)
32
32
32
32
32
Deployment Heaters (W)
N/A
N/A
N/A
N/A
N/A
Total CBE Heater Power (W)
104
204
244
267
329
Allowable CBE (W)
157
216
261
379
746
Max Heater Power Available
Energy Balance (W)
189
259
313
455
895
C&DH Heaters (W)
VCHP, LEND, OB Htrs
SC Op (W)
Star Trackers, LR Htrs
*Does not include Diviner Operational Heater
Charles Baker - Thermal
29
Cold Dissipation with Cold Environment Raw Htr Pwr
Orbiter Power
Beta 0
Beta 50
Beta 60
Beta 70
Beta 90
B90
Safehold
Electrical Dissipation (W)
382
382
382
382
382
212
Thermal Dissipation (W)
407
407
407
407
407
237
Avionics EHB (W)
10
23
25
26
27
77
15
54
62
69
88
30
5
15
17
19
25
N/A
Propulsion Op/Surv (W)
48
111
119
124
134
141
Instrument Op* (W)
0
23
29
35
66
N/A
Instrument Survival* (W)
0
0
0
0
0
200
Gimbal Heaters (W)
37
37
37
37
37
48
Deployment Heaters (W)
N/A
N/A
N/A
N/A
N/A
32
Total CBE Heater Power (W)
115
263
288
312
376
528
Allowable CBE (W)
367
426
471
589
956
1098
Max Heater Power Available
Energy Balance (W)
441
511
565
707
1147
1317
C&DH Heaters (W)
VCHP, LEND, OB Htrs
SC Op (W)
Star Trackers, LR Htrs
*Does not include Diviner Operational Heater
Charles Baker - Thermal
30
Ongoing Trade Studies
Trade Study
Options
Criteria
Radiator Coating
Selection
• Handling
• OSRs
• Cost
• NS43C white paint • Performance
Isothermal Avionics • VCHP
and Battery Radiator • Passive CCHP
Optimization
• Propylene LHP
VCHP Evaporator
Orientation
• In-Plane VCHPs
• Vertical Leg
Outcome
Preliminary
studies show
OSR produce
lower transients
• Complexity of
hardware
• Control heater power
savings
• Radiator Size Savings
• Cost
VCHP is
currently
baselined but
encapsulates
other designs
• Eliminate header heat
pipes
• Moderate savings
In plane VCHPs
currently
baselined
Charles Baker - Thermal
31
Active versus Passive Trade
• In the interest of minimizing thermal complexity, the
simplest thermal design possible is a desirable
attribute
• By baselining a VCHP design, the passive design
(replacing the VCHP with a CCHP) is encapsulated
• Between PDR and CDR a trade on thermal power
versus allocation will be performed on both the
avionics and battery thermal control systems
Charles Baker - Thermal
32
Preliminary Reliability Predictions/Analysis
Results
There are approximately 100 heater circuits in thermal subsystem. Assume 40 of those are mission critical with redunda
R(Non-Critical 60-Circuits) does not contribute to the total mission loss then R(Thermal) = 0.9999999
Mission Time = 10,220 hours (14 months)
Failure Rates and Percentages
Components
Thermostat
Thermostat
Kapton Heater
Kapton Heater
Thermistor
Thermistor
Total Failure Rate =
R(Thermal, 1-Mission Critical Circuit) =
R(Thermal, 40-Mission Critical Circuits) =
Card/Component
Reliability
0.999847
0.999847
0.999982
0.999982
0.999993
0.999993
Block
Redundancy
0.9999999
1.0000000
1.0000000
Failure Rates
1.5000E-08
1.5000E-08
1.8000E-09
1.8000E-09
7.2000E-10
7.2000E-10
9.2336E-12
Percent of Total
99.5840%
0.3586%
0.0574%
100.00%
0.9999999
0.9999962
Charles Baker - Thermal
33
Orbiter Thermal Verification
• Component Level Qualification
– 8 Cycles on avionics (min) prior to delivery
– 4 Cycles (instrument component) + 4 Cycles (instrument) prior to delivery
– Isothermal Panel and Radiator/VCHP Assemblies qualified in separate
Component Level T-Vac Tests
• Performance Testing
• Verification of Heat Pipe Conductances
• Orbiter Level T-Vac Qualification Objectives
–
–
–
–
Qualify orbiter assembly and thermal joints
Verify radiator and heater sizing (Orbiter)
Demonstrate performance during orbital transient simulation
Correlate thermal model
Charles Baker - Thermal
34
Orbiter Thermal Verification (cont.)
Test will include:
Hot Case Thermal Balance
Hot Case Transient Simulation
Cold Case Thermal Balance
Hot and Cold Starts
Survival Case Heater Verification (Thermal Balance levels)
4 Thermal Cycles
Lunar Reconnaissance Orbiter Thermal Vacuum Test Concept
Bakeout
8
Surv Hot
7
6
Qual Hot
Control Condition
5
Hot Op
4
Ambient
3
Cold Op
2
Qual Cold
1
Survival
ot
C
yc
l
Tr e #
C an 1
si
ol
t
d
C ion
yc
le
Tr
#1
a
H ns
iti
ot
C on
yc
l
Tr e #
C an 2
ol
si
t
d
C ion
yc
le
Tr
#2
a
H ns
iti
ot
C on
yc
l
T e#
C ran 3
si
ol
t
d
C ion
yc
le
Tr
#3
a
H ns
itio
ot
C
yc n
l
Tr e #
C an 4
si
ol
t
d
Cy ion
cl
e
Tr
#
an 4
si
ti
Am on
bi
en
Ba t
ck
fil
l
H
ou
t
ot ot
Tr TB
an
si
en
t
C
ol
d
TB
Su
rv
iv
al
TB
H
H
ke
Ba
D
ow
n
0
p
–
–
–
–
–
–
–
Pu
m
•
Test Segment
Charles Baker - Thermal
35
Thermal Safety
•
Hazard Descriptions
– Release of ammonia from heat pipes
– Mission critical heater circuits in Propulsion System
•
Worst Case Severity
– Critical
•
Control and Mitigation Strategy (Mission Critical Heater Circuit)
– Thermal analysis
– 2 fault tolerant heaters with ground monitoring and limit checking
•
Control and Mitigation Strategy (Release of Ammonia)
–
–
–
–
–
–
Over Temperature Heaters as required by analysis
2 fault tolerant heaters with ground monitoring and limit checking
Proof testing of lines and components >2 x MDP
Burst Pressure by Similarity >4 x MDP
NDE as appropriate (X-Ray)
Leak before burst Analysis
Charles Baker - Thermal
36
Schedule
Item
UID
Date
Status
Delta Thermal Peer Review EELV
5869
4/14/2006
Currently comfortably on schedule
Order Avionics Panel Embedded Heat Pipes
4121
6/30/2006
On schedule
Order early needed heaters, thermistor,
thermostats
5930
8/1/2006
On schedule
Receive Heat Pipes and provide to Honeycomb
Panel Assembler
Completed
4121 first
step
10/2006
Order VCHPs, Radiators, remaining Heaters,
Thermistors, thermostats after CDR
5919, 5944
10/2006
Receive early thermal hardware
5930
completed
3/21/2007
One month schedule margin
Receive remainder of the thermal hardware
5919
completed
6/4//2007
One month schedule margin
Vibe + T-Vac test of VCHP and Radiator
Assembly
5974
6/18/2007
Test Avionics Panel in T-Vac
5954
6/29/2007
Occurs right before I&T Delivery
VCHP and Radiator due
5789
12/14/2007
Five months schedule margin
Start Orbiter Thermal Vacuum Test
5813
5/22/2008
Charles Baker - Thermal
37
LRO Thermal Risk Matrix
Rank &
Trend
5
L
I 4
K
E
L 3
I
H
O 2
O
D
3
1
2
1
3
1
4
5
CONSEQUENCES
Criticality
L&C Trend
Approach
High
↓
Decreasing (Improving)
M – Mitigate
Med
↑
Increasing (Worsening)
W – Watch
Low
→
Unchanged
A – Accept
New since last month
R – Research
*
Risk Title
Risk ID
1
*
W
Design Configuration Changes on
the Avionics Panel
THERM-85
2
*
W
Instrument Thermal Design
Development Lagging
THERM-84
3
↓
M
Instrument Thermal Model Delivery
is lagging
THERM-9
•
2
Approach
Design Configuration Changes on the
avionics panel - Yellow
– Avionics Panel is being mechanically
laid out and designed in the March
/April timeframe
– If major changes in layout or
configuration occur later than mid May
then the heat pipe design may not meet
the thermal design requirements
– This will impact the thermal and
structural delivery schedule
– Currently design is on-track to meet all
delivery dates and project understands
impact to thermal and mechanical over
late changing
Charles Baker - Thermal
38
Thermal Road Forward
• Baseline Delta Thermal Peer Review – April 14, 2006
• Drawings in for Heat Pipe Procurement – June 30,
2006
• CDR Avionics Radiator, Instrument, Propulsion
Thermal Design – October, 2006
Charles Baker - Thermal
39
Status
• Orbiter Reduced Model Complete
• Relative Temperature gradients provided to
Mechanical Analysis for Instrument Module
• Detailed Model Analysis completed by peer review
• Thermal STOP Analysis inputs to Mechanical
March 1, 2006
• Control Heater power is within allocation
• All temperature predicts have Gold Rules margin (5
C or better)
• In process of a detailed analysis of Isothermal Panel
with a detailed model
Charles Baker - Thermal
40
Conclusions
• LRO Thermal has adequate resources to control
Instruments, Avionics, and Propulsion
• LRO Thermal has adequate temperature margins
• LRO Thermal is ready to proceed to CDR following
Thermal Peer Review in April
Charles Baker - Thermal
41
LRO PDR
Guidance Navigation and Control
Systems Overview
Eric Holmes
February 8, 2006
Overview
•
•
•
•
GN&C Team Leads
Top Level Mission Constraints & Related Functions
ACS Modes
Pointing and Alignment
Eric B. Holmes – GN&C Overview
2
GN&C Team Leads
• Code 591 – GN&C Systems Engineering Lead
– Eric Holmes
• Code 595 – Flight Dynamics Lead
– Mark Beckman
• Code 591 – ACS Lead Engineer
– Jim Simpson
• Code 595 - ACS Analysis Lead
– Joseph Garrick
• Code 596 – GN&C Hardware Lead
– Ken McCaughey
• Code 582 – ACS Flight Software Lead
– Ji-Wei Wu
• Code 597 – Propulsion Lead Engineer
– Chuck Zakrzwski
Eric B. Holmes – GN&C Overview
3
Top Level GN&C Mission Constraints & Related
Functions
•
14 Month Mission
– Consumables For 14 Months and a 4 Year Extended Mission (Low Maintenance Orbit)
– Met By Propellant Budget
•
Accommodate 7 Nadir Pointing Instruments With Moderate to Fine Pointing And
Knowledge Requirements
– Three Axis Stabilized Platform, Zero Momentum ACS
– Star Tracker/Gyro/Kalman Filter ACS and Attitude Determination
•
Class C+ Spacecraft Bus
– Single String Architecture With Selective Redundancy
•
Calendar Year 2008 Launch
– Aggressive but Achievable Schedule
•
•
Enhanced Expendable Launch Vehicle
50 km (+/- 20 km) Mission Orbit
– Regulated Hydrazine Propulsion System
– Sufficient Propellant for
• MCC
• LOI
• Mission Orbit Maintenance
Eric B. Holmes – GN&C Overview
4
ACS Mode Diagram
Power-On/Reset
Cmd
Sun-Safe
Sun-Safe
• Manage sun relative to S/C
• Wheels, CSSs, IMU-optional
• SA in predefined position
• HGA in predefined position
Cmd,
Cmd
Safing
Observing
• Nadir, Inertial, Offset pointing
• Wheels, IMU, STs
• SA tracking Sun
• HGA tracking Earth
Auto, Cmd,
Safing
Delta-H
Delta-H
• Hold attitude, unload mom.
• Thrusters, IMU
• SA in predefined position
• HGA in predefined position
Cmd
Auto, Cmd,
Safing
Delta-V
• Hold attitude, adjust velocity
• Thrusters, IMU, STs
• SA in predefined position
• HGA in predefined position
Cmd
Observing
Delta-V
Auto, Cmd,
Safing
Eric B. Holmes – GN&C Overview
5
GNC FSW Control Modes
•
Sun-Safe
– Maintain Sun on the solar array and keep the spacecraft in a safe condition until ground
takes action
– Use the CSSs and the IRU (if available) to measure sun direction and body rate
– Command the SA to its index position and then use the RWs to point the SA normal
vector in the sun direction, with body rate subsequently nulled
•
Observing
– A desired target attitude quaternion for the spacecraft is computed as a function of
time. Use the STs and the IRU for attitude and rate measurement and the RWs for
control
– Provide Lunar nadir-pointing during Lunar science with other targeting options of
offset-pointing and absolute inertial pointing
•
Delta-H
– Use thrusters for momentum dump as commanded from Sun-Safe or Observing.
Transition back to original mode when done
•
Delta-V
– Entered from Observing with the proper attitude
– Use thrusters to perform orbit maneuvers (MCC, Lunar Orbit and maintaining attitude.
Zero-torque commands are supplied to the RWs
Eric B. Holmes – GN&C Overview
6
Pointing & Alignment Budget
• Some Data Products and Instruments Drive Knowledge,
Some Drive Accuracy
• Instruments on Spacecraft Bus Have Different Allocations
Than Instruments on Instrument Module
• Pointing & Alignment Budgets are Controlled in a Document
– Pointing & Alignment Specification (431-SPEC-000113)
• Pointing Allocations Are Derived For:
– Each of the LRO Instruments
– The High Gain Antenna System
– The Solar Array System
• Allocations Are Derived Such That if They Are Met, Then
The Instrument Requirements are Met
Eric B. Holmes – GN&C Overview
7
Pointing & Alignment Budgets Development
Process
• Spacecraft Error Sources Are Identified
–
–
–
–
ACS
Mechanical
Thermal
Jitter
• Allocations Are Assigned
– Spacecraft Errors
– Instrument Errors
• Error Sources Are Classified
– Biases
– Time Varying
• Error Allocations Are Appropriately Combined
– RSS for Time Varying Sources (grouped by frequency)
– RSS’d Sub-totals Are Summed
Eric B. Holmes – GN&C Overview
8
Instrument Pointing Requirements
Instrument
Crater
Knowledge (arc-sec)
[requirement/allocation]
Accuracy (arc-sec)
[requirement/allocation]
Roll
Roll
Pitch
Yaw
Pitch
Yaw
36000 18898
36000
18898
126000
18913
126000
18913
126000
18913
36000
18898
Diviner
619
616
619
616
619
616
1238
631
1238
631
1238
631
LAMP
216
215
216
215
216
215
540
230
540
230
540
230
LEND
360
331
360
331
360
331
360
346
360
346
360
346
LOLA
206
143
206
143
206
143
3600
158
3600
158
3600
158
LROC-NAC
206
116
206
116
206
116
516
131
516
131
516
131
LROC-WAC
1031
501
1031
501
1031
501
2063
516
2063
516
2063
516
Mini RF
2160
2091
2160
2091
2160
2091
2160
2106
2160
2106
2160
2106
Eric B. Holmes – GN&C Overview
9
Example Allocations Chart (LOLA Data)
ACS
ACS Knowledge
Jitter
Jitter
Thermal
Thermal Distortion
Static Bias
Calibration Error
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
20
105
30
Spacecraft Total
Instrument
ACS
97 Mech. allocation
509 Mech./Thermal allocation
145 Instrument
Knowledge Total
Knowledge Reqt.
185
30
215
216
897
145
1042
1047
ACS Control
Accuracy Total
Accuracy Reqt.
15
230
540
73 ACS analysis
1115 Knowledge + Control
2618 Instrument Team
Eric B. Holmes – GN&C Overview
Add
Instrument
Add
Instrument Team
10
End to End – Pointing Performance
• End to End Pointing Performance is a Function of
Both Attitude Determination and Orbit
Determination
• Orbit Determination Will Be Covered In The Flight
Dynamics Presentation
Eric B. Holmes – GN&C Overview
11
GN&C Status
• Reviews Held
– Propulsion Subsystem Specification Peer Review on 6/15/05
• Requirements for Out-of-House Propulsion Procurement
• Majority of Requirements are Still Applicable for Current Configuration
– GN&C PDR Peer Review on 9/29/05
• Requirements for Flight Dynamics, ACS and ACS Hardware
• Complete PDR Level Analysis for Flight Dynamics, ACS and ACS
Hardware
• Future Design Reviews
– GN&C/Propulsion PDR Peer Review on 4/26/05
• Complete PDR Level Requirements/Analysis Peer Review for Flight
Dynamics, ACS, ACS Hardware and Propulsion
• Review Reflects Current EELV Baseline
– GN&C/Propulsion CDR Peer Review on 9/11/06
• Complete CDR Level Requirements/Analysis Peer Review for Flight
Dynamics, ACS, ACS Hardware and Propulsion
• Review Reflects Current EELV Baseline
Eric B. Holmes – GN&C Overview
12
Backup Charts
• Detailed Pointing Allocations
Eric B. Holmes – GN&C Overview
13
CRaTER Pointing Allocation
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
ACS
ACS Knowledge
Jitter
Jitter
Thermal
Thermal Distortion
500
2424 Mech./Thermal allocation
Static Bias
1-g Release
Launch Shift
Alignment
100
30
300
318
485
145
1454
1540
898
18000
18898
36000
4352
87266
91618
174533
50
Static Bias RSS
Spacecraft Total
Instrument
Knowledge Total
Knowledge Reqt.
ACS
ACS Control
Accuracy Total
Accuracy Reqt.
15
18913
126000
Eric B. Holmes – GN&C Overview
242 Mech. allocation
Mech. analysis
Mech. analysis
Mech. analysis, Instrument
RSS
Add
Instrument
Add
Instrument Team
73 ACS analysis
91691 Knowledge + Control
610865 Instrument Team
14
Diviner Pointing Allocation
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
ACS
ACS Knowledge
Jitter
Jitter
Thermal
Thermal Distortion
Static Bias
1-g Release
Launch Shift
Alignment
20
75
10
10
76
Static Bias RSS
Spacecraft Total
Instrument
616
Knowledge Total
Knowledge Reqt.
ACS
490
ACS Control
Accuracy Total
Accuracy Reqt.
616
619
15
631
1238
Eric B. Holmes – GN&C Overview
97 Mech. allocation
2376 Mech./Thermal allocation
364
48
48
370
Mech. analysis
Mech. analysis
Mech. analysis, Instrument
RSS
2988 Add
Instrument
2988 Add
3000 Instrument Team
73 ACS analysis
3061 Knowledge + Control
6000 Instrument Team
15
LAMP Pointing Allocation
ACS
ACS Knowledge
Jitter
Jitter
Thermal
Thermal Distortion
Static Bias
Calibration Error
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
20
105
30
Spacecraft Total
Instrument
ACS
97 Mech. allocation
509 Mech./Thermal allocation
145 Instrument
Knowledge Total
Knowledge Reqt.
185
30
215
216
897
145
1042
1047
ACS Control
Accuracy Total
Accuracy Reqt.
15
230
540
73 ACS analysis
1115 Knowledge + Control
2618 Instrument Team
Eric B. Holmes – GN&C Overview
Add
Instrument
Add
Instrument Team
16
LEND Pointing Allocation
ACS
ACS Knowledge
Jitter
Jitter
Thermal
Thermal Distortion
Static Bias
1-g Release
Launch Shift
Alignment
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
20
100
485 Mech./Thermal allocation
75
10
30
81
364
48
145
395
Knowledge Total
Knowledge Reqt.
231
100
331
360
1122
485
1607
1745
ACS Control
Accuracy Total
Accuracy Reqt.
15
346
360
73 ACS analysis
1679 Knowledge + Control
1745 Instrument Team
Static Bias RSS
Spacecraft Total
Instrument
ACS
97 Mech. allocation
Eric B. Holmes – GN&C Overview
Mech. analysis
Mech. analysis
Mech. analysis, Instrument
RSS
Add
Instrument
Add
Instrument Team
17
LOLA Pointing Allocation
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
ACS
ACS Knowledge
Jitter
Jitter
10
Thermal
Thermal Distortion
50
242 Mech./Thermal allocation
Static Bias
Calibration Error
22
105 Instrument
Spacecraft Total
Instrument
Knowledge Total
Knowledge Reqt.
ACS
ACS Control
Accuracy Total
Accuracy Reqt.
112
31
143
206
15
158
3600
Eric B. Holmes – GN&C Overview
48 Mech. allocation
541
150
691
1000
Add
Instrument
Add
Instrument Team
73 ACS analysis
764 Knowledge + Control
17453 Instrument Team
18
LROC WAC Pointing Allocation
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
ACS
ACS Knowledge
Jitter
Jitter
Thermal
Thermal Distortion
100
Static Bias
1-g Release
Launch Shift
Alignment
75
10
300
309
364
48
1454
1500
Mech. analysis
Mech. analysis
Mech. analysis, Instrument
RSS
Knowledge Total
Knowledge Reqt.
449
52
501
1031
2179
250
2429
5000
Add
Instrument
Add
Instrument Team
ACS Control
Accuracy Total
Accuracy Reqt.
15
516
2063
10
Static Bias RSS
Spacecraft Total
Instrument
ACS
Eric B. Holmes – GN&C Overview
48 Mech. allocation
485 Mech./Thermal allocation
73 ACS analysis
2501 Knowledge + Control
10000 Instrument Team
19
LROC NAC Pointing Allocation
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
ACS
ACS Knowledge
Jitter
Jitter
10
Thermal
Thermal Distortion
50
Static Bias
Calibration Error
5
Spacecraft Total
Instrument
ACS
48 Mech. allocation
242 Mech./Thermal allocation
24 Instrument
Knowledge Total
Knowledge Reqt.
95
21
116
206
461
100
561
1000
ACS Control
Accuracy Total
Accuracy Reqt.
15
131
516
73 ACS analysis
633 Knowledge + Control
2500 Instrument Team
Eric B. Holmes – GN&C Overview
Add
Instrument
Add
Instrument Team
20
Mini-RF Pointing Allocation
arcsec microrad
Source
30
145 LRO-431-SPEC-000113
ACS
ACS Knowledge
Jitter
Jitter
Thermal
Thermal Distortion
500
2424 Mech./Thermal allocation
Static Bias
1-g Release
Launch Shift
Alignment
100
30
500
511
485
145
2424
2476
Knowledge Total
Knowledge Reqt.
1091
1000
2091
2160
5288
4848
10136
10472
ACS Control
Accuracy Total
Accuracy Reqt.
15
2106
2160
73 ACS analysis
10209 Knowledge + Control
10472 Instrument Team
50
Static Bias RSS
Spacecraft Total
Instrument
ACS
Eric B. Holmes – GN&C Overview
242 Mech. allocation
Mech. analysis
Mech. analysis
Mech. analysis, Instrument
RSS
Add
Instrument
Add
Instrument Team
21
HGA Pointing Accuracy Allocation
Uncompensated
asec
deg
Post-calibration
asec
deg
Source
Bias
Known
Hardware Alignment Errors
Thermal Distortion
Sum Subtotal
Random/Gnd-to-Orbit
Hardware Alignment Errors
Launch/Deployment/Gravity Release Errors
Thermal Distortion
RSS Subtotal
Random
Very Low Freq.
ACS/GN&C Knowledge/Command Errors
Thermal Distortion
RSS Subtotal
Low Freq.
Thermal Distortion
High Freq.
ACS/GN&C Knowledge/Command Errors
Dynamic Pointing Errors
RSS Subtotal
Total
Requirement
6084
540
6624
1.69
0.15
1.84
50
100
150
0.01 SDO Mech & Comm, Allocation
0.03 SDO Mech, Allocation
0.04 Add
2693
0.75
173
352
392
0.05 SDO Mech & Comm, Allocation
SDO Mech & Comm, Allocation
0.10 LRO Mech/Thermal Allocation
0.11 RSS
2693
0.75
12
108
109
0.00
0.03
0.03
12
108
109
0.00 LRO ACS Analysis
0.03 1.5x SDO Thermal Allocation
0.03 RSS
163
0.05
163
0.05 1.6x SDO Thermal Allocation
50
250
255
9843
0.01
0.07
0.07
2.73
50
250
255
1069
1080
Eric B. Holmes – GN&C Overview
0.01
0.07
0.07
0.30
0.30
LRO ACS Analysis
0.75x SDO ACS & Mech, Allocation
RSS
Add
Comm
22
SA Pointing Accuracy Allocation
Uncompensated
asec
deg
Source
Bias
Known
Hardware Alignment Errors
Thermal Distortion
Sum Subtotal
Random/Gnd-to-Orbit
Hardware Alignment Errors
Launch/Deployment/Gravity Release Errors
Thermal Distortion
RSS Subtotal
Random
Very Low Freq.
ACS/GN&C Knowledge/Command Errors
Thermal Distortion
RSS Subtotal
Low Freq.
Thermal Distortion
High Freq.
ACS/GN&C Knowledge/Command Errors
Dynamic Pointing Errors
RSS Subtotal
Total
Requirement
6084
540
6624
1.69 SDO Mech & Comm, Allocation
0.15 SDO Mech, Allocation
1.84 Add
398
2693
501
2768
0.11
0.75
0.14
0.77
SDO Mech & Comm, Allocation
SDO Mech & Comm, Allocation
LRO Thermal Analysis
RSS
0.02
144
144
0.00 LRO ACS Analysis
0.04 2x SDO Thermal Allocation
0.04 RSS
204
0.06 2x SDO Thermal Allocation
119
338
358
10097
18000
Eric B. Holmes – GN&C Overview
0.03
0.09
0.10
2.80
5.00
LRO ACS Analysis
SDO ACS & Mech, Allocation
RSS
Add
Power
23
LRO PDR
Flight Dynamics
Mark Beckman
February 7, 2006
Key Flight Dynamics Requirements
Higher Level
Rqmt
Identifier
Object Text
I A D T
MRD-24
FDS-1
LRO shall be launched on a EELV
1
MRD-3
MRD-3
FDS-2
1
MRD-24
MRD-3
FDS-3
LRO shall have one daily launch opportunity that consists
of a short or long coast minimum energy direct transfer to
the moon.
LRO shall perform an MCC1 maneuver prior to L+24
(TBD) hours to correct for LV dispersions.
1
MRD-10
MRD-25
MRD-4
FDS-4
MRD-11
FDS-5
MRD-146
FDS-6
MRD-146
FDS-7
The LRO launch window shall be constrained such that
the resulting lunar orbit has a sun beta angle of less than
20 deg at lunar solstice.
LRO shall insert into lunar orbit over the lunar southern
pole and have an initial southern argument of periselene.
(TBR)
LRO shall maintain a minimum altitude of 100 km during
LOI phase.
LRO shall maintain a periselene of greater than 200 km
after LOI1.
Mark Beckman – Flight Dynamics
1
1
1
1
2
Key Flight Dynamics Requirements
MRD-146
FDS-8
LRO shall perform up to six LOI maneuvers to insert the
spacecraft into the commissioning orbit.
1
MRD-11
MRD-41
FDS-9
1
MRD-138
FDS-10
MRD-12
FDS-11
LRO shall have a minimum Orbit Determination
Accuracy of 500/18m (Total Position RMS/Radial RMS),
1-sigma, post-processed.
LRO maneuvers shall be performed in view of S-band
tracking stations.
LRO shall perform a stationkeeping cycle approximately
every 27.4 days.
MRD-146
MRD-12
FDS-12
1
MRD-12
FDS-13
The primary mission shall be conducted in a circular
mapping orbit with a nominal altitude of 50 +/- 20 km
(altitude is measured to mean lunar surface).
LRO's stationkeeping shall target periselene in the
southern hemisphere 50% of the time.
MRD-146
MRD-11
FDS-14
The commissioning orbit shall be a frozen orbit at 30x216
km altitude and a 270 deg argument of periapsis.
1
MRD-10
Mark Beckman – Flight Dynamics
1
1
1
3
Trade Studies
Trade Study
Options
Criteria
Transfer Orbit
• Transfer orbit ΔV budget
• Direct Transfer
• Weak Stability Boundary • Operations cost
Direct Transfer
(WSB)
• Spacecraft design
complexity
• Phasing Loops
Thrusters
• Two 80-N thrusters
• Four 80-N thrusters
• One 440-N thruster
• Two 440-N thrusters
• LOI-1 minimum required
thrust
Four 80-N thrusters
• LOI-1 contingencies
• ACS controllability
Mission Orbit
• 100 km circular orbit
• 50 km circular orbit
• 30 km circular orbit
• 30x216 frozen orbit
• SK ΔV budget
• Accuracy of science
measurements
• Instrument data set
quality
Stationkeeping
• Optimize time between
SK maneuvers
• Minimize altitude
variation
• Bias periselene
• Repeatable pattern
Repeatable pattern with
• Ground station view
equal north/south
• SK ΔV budget
periselene distribution
• Science team preference
Mark Beckman – Flight Dynamics
Outcome
50 km circular orbit
4
Alternative Transfer Options to Moon
•
Weak Stability Boundary (WSB)
–
Used by Japanese mission HITEN to reach lunar orbit without sufficient fuel to directly insert
•
–
–
–
Applicable to weak captures about the moon
ΔV savings to low lunar orbit are very small (~3%)
Higher C3 required from launch vehicle
•
–
–
–
–
•
Slightly less launch mass capability
Initial inclination control is difficult and could add costly ΔV
Transfer time of about 100 days
Limited launch window of several days per month
Not a robust transfer option when designing a transfer to a very specific low lunar orbit
Phasing Loops
–
–
–
–
–
–
•
Caused by launch vehicle failure
One or more phasing loops about the Earth prior to direct lunar transfer
Adds 10+ days to transfer time
Passes through radiation belts multiple times
Creates an added environment (elliptical Earth orbit) in which spacecraft must operate (i.e.
additional eclipses, spacecraft torques, etc.)
Initial launch vehicle correction maneuver may be reduced and performed at apogee
Initial launch into less than -2 km2/sec2 C3 requires additional spacecraft fuel mass
Trade on alternative transfers complete – direct transfer is baseline
–
Project determined that advantages and simplicity of Direct Transfer outweighed any possible
benefits associated with other options
Mark Beckman – Flight Dynamics
5
GN&C Peer Review
• GN&C Peer Review held September 29, 2005
• Total of 29 RFAs received
• Only two Flight Dynamics related RFAs
– HGAS Calibration on FD Product List
– Propellant Budget for Insertion Contingency
• Both have already been addressed
Mark Beckman – Flight Dynamics
6
Lunar Transfer
• Minimum energy transfer to Moon
– Transfer duration chosen to minimize orbital energy off launch
vehicle
– Maximizes launch vehicle capability
– Two opportunities per day – long and short coast
Mark Beckman – Flight Dynamics
7
Short & Long Coast
•
•
•
In order to minimize time on battery power and improve ground station
coverage, only one of each short & long coast solution per day will be
considered
Trans-Lunar Insertion (TLI) burn is fixed at –(Moon Latitude)
Short coast solution achieves TLI latitude with a shorter coast time and
later launch time
Mark Beckman – Flight Dynamics
8
TLI Quadrants
• In coast orbit, to optimize ground station coverage,
Argument of Latitude must be 1st or 4th quadrant (white
segments in plot)
• Maximum coast time up to 68 min (~ ¾ orbit)
Launch
SECO-1
Short coast solutions
for northern latitude TLIs
Long coast solutions
for northern latitude TLIs
Short coast solutions
for southern latitude TLIs
Mark Beckman – Flight Dynamics
Long coast solutions
for southern latitude TLIs
9
Lunar Orbit Insertion
•
Cis-lunar transfer achieves a periselene distance of ~200 km after LOI-1
– Closest approach to moon is only ~ 300 km
•
•
Critical initial lunar insertion maneuver into 5-hr eccentric lunar orbit
Four additional LOI maneuvers to place the spacecraft into the 30x216
km quasi-frozen orbit at about L+6 days
Mark Beckman – Flight Dynamics
10
Lunar Orbit Insertion Contingencies
• Entire Lunar Orbit Insertion maneuver is in view of the earth
• No onboard failure detection or correction is required/desired
• Robust against thruster failures and restart delays
• 1st burn (4x80N thrusters) lasts 45 min; results in a 5-hr orbit
• Need about half the burn (time or thrust) to capture into orbit
5-hr Orbit
• Contingency A: loss of single 80 N thruster (start of burn)
– Continue with nominal 45 min burn (3/4 thrust)
– Capture into somewhat higher orbit
– Pay small ΔV penalty (< unallocated ΔV margin)
• Contingency B: loss of diagonal set of 80 N thrusters
– Continue with nominal 45 min burn (1/2 thrust)
– Capture into two day lunar orbit (worst case)
– Pay large ΔV penalty (~1/2 of the extended mission)
2-day Orbit
• Contingency C: loss of thruster(s) prior to LOI
– Re-plan the burn to account for thruster loss
– Capture into somewhat higher orbit
– Pay small ΔV penalty (< unallocated ΔV margin)
• Contingency D: processor reset during LOI
– Immediately execute canned restart sequence (ground)
– 10 min interruption if processor was only failure
– Single reset would give higher capture orbit (small penalty)
– Two resets would give two day capture orbit (large penalty)
– An autonomous restart was deemed to be too unpredictable
Mark Beckman – Flight Dynamics
11
Commissioning Orbit
• Frozen orbit at 30x216
km altitude
• Periselene fixed over
lunar south pole
• No secular growth in
eccentricity or
argument of periapsis
Ecc
AoP
– No stationkeeping
required
• Remain in
commissioning orbit
for up to 60 days
Mark Beckman – Flight Dynamics
12
Mission Orbit
•
50 km mean altitude
– Lunar non-spherical gravity introduces secular eccentricity growth
– Altitude controlled to ± 20 km via stationkeeping
•
•
113 min period
Approximately 90° lunar equatorial inclination
– Lunar non-spherical gravity introduces periodic and secular variations on
inclination
•
27.4 day lunar sidereal period
Mark Beckman – Flight Dynamics
13
Stationkeeping Strategy
•
Stationkeeping strategy meets
all requirements
Point every ascending node
Lunar longitude labeled
– Altitude controlled to within
±15 km
– SK maneuvers at lunar
longitude of ascending node of
270 deg
– 12-month SK ΔV cost of 150
m/sec
•
Repeatable SK cycle
SK ΔV 2
SK ΔV 1
– Phase plot is same every lunar
sidereal period
– SK maneuver pairs exactly
27.4 days apart
• 66 min between maneuvers
within pair
– SK ΔV costs are the same each
period
Mark Beckman – Flight Dynamics
14
ΔV Budget
MISSION PHASE
ΔV
(m/sec)
Fuel
Mass
(kg)
MCC
70
61.8
3σ LV errors
MCC @ 24 hours, sooner uses less fuel, consistent with 9 m/sec
LV velocity error at separation
Lunar Orbit Insertion - 1st burn
567
422.1
Deterministic
1st insertion burn, finite ΔV, any launch date, 320 N thrust
LOI-2
185
115.1
Deterministic
320 N thrust
LOI-3
133
76.7
Deterministic
320 N thrust
LOI-4
41
22.7
Deterministic
Mission Orbit Insertion (3 burns)
50
27.1
Deterministic
Stationkeeping
150
77.5
Deterministic
Extended Mission
100
47.2
Placeholder
Sufficient for 8 additional months in mission orbit OR boost to
frozen orbit for 4+ years
Momentum Unloading
–
17
Conservative
Calculated based on worst case mass properties and
perturbations
Other
–
12
Conservative
Residuals
Unallocated Margin
30
14.9
Conservative
Additional breathing room in conservative budget
1326
894
Total
COMMENTS
Mark Beckman – Flight Dynamics
50 +/- 20 km altitude
15
Orbit Determination Requirements
• LRO shall have a minimum 1-σ post-processed Orbit
Determination Accuracy of 500/18m (Total Position
RMS/Radial RMS)
• Derived requirement from actual FDF Lunar Prospector
overlap compares
– LP OD accuracy was conservative; significant improvements since
– Assumes tracking schedule, measurement accuracy and gravity
modeling similar to LP
• OD Process
– Weighted least squares batch processing: 55-hr batch arcs
– Daily OD
– 31 hr overlap period
Mark Beckman – Flight Dynamics
16
OD Accuracy
Lunar Prospector Orbit Determination Accuracy
Component
100-km orbit w/ LP75D
30-km orbit w/ LP100J
Radial Root Mean Square (RMS)
13 m
28 m
Crosstrack RMS
155 m
673 m
Alongtrack RMS
189 m
976 m
Position RMS
270 m
1249 m
Mapping LP OD Accuracy to LRO
1400
30
1200
25
LP
Extended
20
LP
Nominal
800
15
600
Pos RSS
Radial
Power (Pos RSS)
Power (Radial)
10
LRO
Nominal
400
Radial Accuracy (m)
Total Position Accuracy (m)
1000
5
200
0
0
20
40
60
80
100
0
120
Mean Altitude (km)
Mark Beckman – Flight Dynamics
17
Orbit Determination Improvements
• Recently completed study using actual Lunar Prospector
tracking data during the extended mission (30 km mean
altitude)
• Incorporated the latest models and process improvements
• Real-time LP OD definitive overlap compares were 1250 m
• Improved process LP OD definitive overlap compares are 60
m
– Factor of 20 improvement in accuracy!
• Expect LRO OD real-time accuracy to be ~ 50-100 m (1sigma)
– 5 m radial (1-sigma)
Mark Beckman – Flight Dynamics
18
Tracking Schedule & Measurements
• 113 min orbital period
• Lunar occultation of up to 48 min/orbit
• Tracking schedule requirement of 30 min/orbit
– Goal of up to 58 min/orbit
– Not required to be consecutive 30 min of tracking
– Limited by constraints on HGA
• HGA can only point above local horizontal
– Range and Doppler measurements required for entire pass
• Two-way coherent measurements
– Doppler
• 1 mm/sec (3-sigma) at one ground station
• 8 mm/sec (3-sigma) at others
• 5 sec integration time
– Range
• 15 m (3-sigma)
Mark Beckman – Flight Dynamics
19
Tracking Stations
• Primary Stations
– NASA White Sands Complex (WS1) 18-Meter S/KaBand Antenna
– Universal Space Network (USN) 15m S-Band at Redu,
Belgium
– USN 13m S-Band at Dongara, Australia
– USN 11m S-Band at Kiruna, Sweden
– USN 13m S-Band at South Point, Hawaii
• All stations have/will have ranging systems that
meet LRO requirements
– WS1 designed to meet 1 mm/sec (3-sigma) Doppler
accuracy
– All other stations meet 8 mm/sec (3-sigma) requirement
Mark Beckman – Flight Dynamics
20
Event Prediction
• Predicted OD accuracy is important in event prediction
• Along-track error converts directly into event timing
• Several instruments have need for accurate event predictions:
– LROC – time of flyover of various targets to ensure capture within
limited exposure time
– Diviner – ascending node crossing time keys start of measurement
phase
– LAMP – instrument failure when open during sunlight, terminator
crossing times are backup to onboard sensor
• LP re-processing study shows:
– 36-hr prediction → 200 m (1-sigma) → 0.13 sec
– 84-hr prediction → 330 m (1-sigma) → 0.21 sec
Mark Beckman – Flight Dynamics
21
Flight Dynamics Facility (FDF)
•
•
•
•
•
•
FDF is an institutional GSFC facility
Secured fully integrated operational control
center and remote lab
Provided primary navigation and mission
design for past two US lunar missions:
Clementine and Lunar Prospector
Support LRO as part of a Project Support
Level Agreement (PSLA)
FDF responsible for routine orbit
determination, maneuver planning and
product generation for LRO
FDF Tools:
–
Orbit Determination
•
–
Supporting Software
•
–
Testing Readiness and Maintenance Program
(TRAMP)
Mission Design/Maneuver Planning
•
–
Goddard Trajectory Determination System
(GTDS)
Satellite Tool Kit™ (STK™)/Astrogator™
Mission Products
•
Software is TBD
Mark Beckman – Flight Dynamics
22
Station Certification
• FDF certification of ground station tracking
measurement accuracy
• Requires 6-12 passes of tracking data from existing
spacecraft
• Likely use of LEO spacecraft with extrapolation to
lunar orbit
• Dongara and South Point previously certified
Mark Beckman – Flight Dynamics
23
Product List
• Current list of 34 Flight Dynamics products
generated by FDF or ADS:
–
–
–
–
–
–
–
7 SPICE files
3 Orbit determination products
3 Acquisition data products
4 Maneuver products
2 Calibration products
6 Attitude products
9 Miscellaneous products
Mark Beckman – Flight Dynamics
24
Key Deliverables
Deliverable
UID #
Due Date
Status
Finalize Flight Dynamics
Specifications Document
4046
July 31, 2006
Draft on 2/1/06
Tracking Station
Certification Report
1996
August 1, 2007
Not begun
Flight Dynamics Attitude
Determination System
(ADS)
4045
October 1, 2007
Generic ADS system can be configured in two
months
Mark Beckman – Flight Dynamics
25
Summary
• FDAB (GSFC Code 595) and FDF have extensive
experience in mission design and orbit determination
for lunar spacecraft mission
• Dynamics and models for lunar orbit mission design
are well understood
• All requirements are met
Mark Beckman – Flight Dynamics
26
LRO PDR
Propulsion
Charles Zakrzwski
February 8, 2006
Requirements - Key Level 3
Higher Level Rqmt Identifier
MRD-10
MRD-93
4.2.1
MRD-12
4.2.2
MRD-13
MRD-95
MRD - 10
4.2.4.1
MRD-97
4.2.4.2
MRD-94
4.2.7.1
MRD-1
4.2.8
MRD-69
Object Text
I A D T
The propulsion system shall be capable of providing the
delta-V requirements specified in the LRO Technical
Resource Allocations Specification (431-SPEC-000112).
The propulsion subsystem shall be capable of performing all
delta-V maneuvers with an impulse accuracy of better
than 5%.
The propulsion system shall be capable of producing 160 N
of thrust, in such a way that allows for lunar capture, with
backup, as determined by Flight Dynamics.
The propulsion subsystem shall have the capability of
adjusting the spacecraft momentum within 1 N-m-s of a
desired set-point over the entire mission lifetime.
2
The maximum amount of propellant shall not exceed the
propellant budget documented in the LRO Technical
Resource Allocations Specification (431-SPEC-000112).
This propellant allocation shall include all propellant
required for all delta-V and attitude control maneuvers as
well as subsystem residual propellants and uncertainties in
propellant loading.
The propulsion subsystem shall be designed to use HighPurity grade hydrazine per “MIL-PRF-26536E with
Amendment 1” as the propellant.
3
Charles Zakrzwski - Propulsion
3
4
3
3
2
Requirements - Key Level 3
GSFC-STD-1000 1.27
4.3.3.3
4.3.10.2
All heaters shall be single fault tolerant to overheating on
orbit. The propulsion subsystem shall be two fault tolerant
to overheating the fueled SC in order to meet range
safety requirements.
The propulsion subsystem shall be designed to meet all
requirements of this document in the event of a failure of
any single thruster.
All valve seats shall be independently tested for leakage at
subsystem MEOP to verify the requirements of Section
4.2.11.3.
Flow control orifices shall be incorporated into the
propellant manifold as necessary to prevent the overpressurization of components or ignition of propellant
due to transient flow (waterhammer) events.
Enable plugs shall be used for all thruster and isolation
valves, catalyst bed heater circuits, and all pyro valves.
3
2
3
GSFC-STD 1000 1.26
4.6.1.2.7.1
GSFC -STD 1000 1.21
GSFC -STD 1000 1.28
4.3.18
GSFC -STD 1000 1.24
4.4.4.5
GSFC - STD 1000 1.22
4.8.4
All test gasses shall be dried to a dew point of -60 F. If
water is used for testing purposes, the subsystem shall then
be dried to -60 F dew point.
3
MRD-58
4.8.9
The propulsion subsystem shall be an all welded subsystem.
Use of threaded mechanical assemblies shall be limited to
the fill and drain vale (e.g., poppet, AN cap, external cap)
and the pyro valve initiators.
3
GSFC - STD 1000 1.20
Charles Zakrzwski - Propulsion
3
3
3
Architecture - Design Summary
•
Monopropellant Hydrazine System
– 894 kg of hydrazine (1326 m/sec delta-V capability)
– Two 28,144 in3 titanium diaphragm propellant tanks (40 in OD oblate
spheriod)
•
Twelve Dual Coil Catalytic Hydrazine Thrusters
– Four on axis 80 N class thrusters for lunar orbit insertion
– Eight canted 20 N class thrusters for attitude control and lunar orbit
maintenance maneuvers
– Isolation valves with redundant coils used to isolate thruster banks in the
event of a thruster failure
– Flow control orifices to prevent waterhammer surges
•
Helium Pressure Regulated System
– 4200 psi COPV Helium pressurant tank (17 in OD x 29.6 in L )
– Two stage regulator (single fault tolerant) 300 psi nominal set point
– Redundant normally closed pyro valve for high pressure isolation for ground
and launch operations
– High pressure latch valve to isolate high pressure source during mission
operations
Charles Zakrzwski - Propulsion
4
Architecture - Schematic
P
Pressure Transducer
Pressurant Tank
(Helium)
P
Fill and Drain Valve
High Pressure
Panel
Pyrotechnic Valve
(Normally Closed)
Fill/Drain
Valve
Panel
R
R
Filter
P
Latch Valve (Dual coil)
R
R
Regulator (2 stage)
Propellant Tank
20 N Class Thruster
Propellant Tank
P
80 N Class Thruster
Flow Control Orifice
Low Pressure
Panel
Dual
Thruster
Module #1
Dual
Thruster
Module #2
Dual
Thruster
Module #3
Charles Zakrzwski - Propulsion
Dual
Thruster
Module #4
Insertion Thrusters
5
Architecture - Component Locations
Pressurant Tank
Propellant Tanks (2)
Low and High
Pressure Panels
Fill and Drain
Valve Panel
Attitude Control
Thrusters (8)
Insertion Thrusters (4)
Charles Zakrzwski - Propulsion
6
Architecture - Operation Modes
•
Launch Configuration
–
–
–
–
–
•
High pressure isolated by pyro valves
Pad pressure in propellant tanks
Tank and high pressure isolation valves closed
Thruster isolation valves open on pad (Optional)
Pressure transducer and latch valve status telemetry
Activation - Post Separation /Deployment
– Open all latch valves
– Fire pyro valves
– Verify pressure ranges
•
Thruster Calibration
– Preheat catalyst beds
– Pulse ACS thrusters - verify temperature & torques
– Pulse insertion thrusters - verify temperatures
•
Mid- Course Correction
– Preheat catalyst beds
– Insertion thrusters steady-state with ACS Thrusters on-pulsed
or
– No insertion thrusters, off-pulse ACS thrusters
Charles Zakrzwski - Propulsion
7
Architecture - Operation Modes
•
Lunar Orbit Insertion
–
–
–
–
•
Lunar Station-keeping
–
–
–
–
–
•
Preheat catalyst beds
Insertion thrusters (4) steady-state with ACS thrusters on-pulsed
1st LOI burn 45 minutes (nominal)
Open tank isolation valves
Preheat catalyst beds
Off - pulse ACS thrusters
One set of two orbit adjust burn each month
Delta-V 5.5 m/sec per burn
Close high pressure latch valve only if regulated pressure increases
Momentum Unloading
– Preheat catalyst beds
– Use ACS thrusters in Delta-H mode
•
Extended Mission
– Use ACS thrusters for maneuvers
– High pressure side will go to blow-down mode
– Improves mass gauging accuracy
Charles Zakrzwski - Propulsion
8
Architecture - Components
Insertion Thruster
Model: Aerojet MR107P
Valve: Dual Seat
Total Impulse: 1,000,800 N-sec
Total Pulses: 4,329
Isp (@250 psi): 227.1 sec
Thrust (@250 psi): 81.8 N (18.4 lbf)
HRV Hardware from X-38
ACS Thruster
Model: Aerojet MR106L
Valve: Dual Seat
Total Impulse: 551,552 N-sec
Total Pulses: 125,000
Isp (@250 psi): 234.5 sec
Thrust (@250 psi): 24.0 N (5.4 lbf )
HRV Hardware new procurement
Charles Zakrzwski - Propulsion
9
Architecture - Components
Propellant Tanks (2)
Model: PSI
6Al-4V Titanium
AF-E-322 Diaphragm
400 psi MEOP
28,144 in3 Internal Volume
99% Expulsion Efficiency
HRV Hardware from X-38
Pressurant Tank
Model: PSI A2100AXL
16.7 ID X 29.6 Cylindrical COPV
4800 psi MEOP
4967 in3 Internal Volume
Polar boss mounted
HRV Hardware new procurement
Charles Zakrzwski - Propulsion
10
Architecture - Components
Regulator
Model: Stanford Mu High Flow
Series Redundant
300 psi primary, 310 psi secondary set points
Inlet pressure range up to 4500 psi
HRV H/W New Procurement
High Pressure
Latch Valve
Model: Vacco
3/8 Torque Motor LV
4500 psi MEOP
24 - 32 Vdc operating voltage
HRV H/W New Procurement
Gas Filter
Vacco P/N F0D10647-02
Etched Disk
12 μm absolute
0.250 g AC Coarse Test Dust Dirt Capacity
4800 psi MEOP
HRV H/W New Procurement
Charles Zakrzwski - Propulsion
11
Architecture - Components
Pressure
Transducers
Taber P/N 2911
High Pressure MEOP 4200
Low Pressure MEOP 350
19-34 Vdc operating voltage
LRO Option from SDO Contract with modifications
High Pressure
Fill/Drain Valve
Moog model 50E889
4500 psi MEOP
3 seals
HRV H/W New Procurement
Low Pressure
Latch Valve
Vacco P/N V1E10948
3/8” Torque Motor LV
300 psi MEOP
19-35 Vdc operating range
LRO Option from SDO Contract (2)
New Procurement [with 350 psi MEOP](4)
Charles Zakrzwski - Propulsion
12
Architecture - Components
Liquid Filter
Vacco P/N F0D10691-01
Etched Disk
15 μm absolute
1.4 g AC Coarse Test Dust Dirt Capacity
500 psi MEOP
HRV H/W New Procurement
Low Pressure
Fill/Drain Valve
Moog model 50-787
400 psi MEOP
3 seals
HRV H/W New Procurement
Pyro Valve
Conax P/N 1832-207
4200 psi MEOP
Titanium body - Stainless Steel initiator housing
Redundant Initiators (NSI’s)
New Procurement
Charles Zakrzwski - Propulsion
13
Architecture - Command/Power/ Telemetry
Propulsion Heater Power (Prime/Redundant) (2x1)
Thruster Cat-bed Htrs (Prime/Redundant) (2 x 12)
PSE
Thruster EP
Propulsion
Deployment
Electronics
C & DH
Propulsion EGSE
Cat-Bed EP
Valve EP
Test
Connector
Thruster Fire Upstream (12)
Thruster Fire Downstream (12)
Pyro Valve Fire (2 x 2)
Isolation Valve Open (Prime Redundant) (2 x 7)
Isolation Valve Close (Prime/Redundant) (2 x 7)
Isolation Valve Status (7)
Pressure Transducer Power (3)
Pressure Transducer Telemetry (3)
Thruster Cat-bed Temperatures PRT’s (12)
Propulsion System Thermisters (32)
Propellant Tanks (8)
Pressurant Tank (2)
Thruster Valves (12)
Lines (7)
Fill/Drain Valve Panel (1)
High Pressure Plate (1)
Low Pressure Plate (1)
LV Umbilical
Charles Zakrzwski - Propulsion
14
Analyses - Fuel Budget
S/C Dry Mass
Total Launch Mass
Max Fill fraction
Max Temperature
Regulated Tank Pressure
949
1846.5
98.50%
50
300
kg
kg
• 3 sigma worst case Isp (at 250 psi inlet)
• Worst case delta-V
C
psi
delta - V
Delta-V Inputs
Calculations
ACS
ACS
fuel
(% DV
fuel)
%
Propellant
Mid course correction
Lunar Orbit Insertion -1
Lunar Orbit Insertion -2
Lunar Orbit Insertion -3
Lunar Orbit Insertion -4
Mission Orbit Insertion
Station Keeping
Margin
Momentum Unloading
Extended Mission
Total Requirement (Usable)
Residual Propellant
Total Propellant
Total Ghe Pressurant
Delta V
m/sec
Isp
sec
Isp
sec
220
220
220
220
220
220
224
224
70
567
185
133
41
50
150
30
na
100
1326
220
220
220
220
220
220
220
220
na
224
220
Charles Zakrzwski - Propulsion
5
3
3
3
3
3
5
5
0
5
Total
Propellant
kg
61.8
422.1
115.1
76.7
22.7
27.1
77.5
14.9
17.0
47.2
882.0
12.0
894.0
3.5
15
Analyses - Operating Pressure
305
300
295
Blowdown Pressure
290
285
320
315
310
Orbit
Maintenance/
Momentum
Unloading
Tipoff/
MCC
LOI 1-6
305
280
275
EOL
270
780
800
820
840
860
880
900
300
295
290
285
280
275
270
0.00
250.00
500.00
750.00
1000.00
Propellant Used (kg)
Charles Zakrzwski - Propulsion
16
Analyses - Component Pressure Requirements
Component
MEOP
Proof/Exposure
Burst
Req’t (psi)
Capability (psi)
Req’t (psi)
Capability (psi)
Req’t (psi)
Capability (psi)
COPV
4200
4500
5250
6000
6300
7200
Pyrovalves
4200
4200
6300
7500
10500
12500
Gas Filter
4200
4800
6300
7200
10500
12000
HPLV
4200
4500
6300
6750
10500
12000
Regulator
4200
4500
6300
6750
10500
11250
HP F&D Valves
4200
4500
6300
6750
10500
11250
HP P-ducers
4200
4200
6300
6300
10500
10500
LP P-ducers
350
350
525
1200
875
1500
LP F&D Valves
350
650
525
975
875
1625
Propellant Tank
350
400
437.5
600
525
800
LPLV (SDO
option)
300
300
525
900
875
1600
LPLV (New)
350
350
525
900
875
1600
Liquid Filter
350
500
525
750
875
1250
5-lb Thrusters
350
400
625
1500
875
2500
20-lb Thrusters
350
400
625
1050
875
2630
Charles Zakrzwski - Propulsion
17
Analyses: Cycle and Throughput Requirements
Cycle Requirements
Requirement1
Capability
MEOP Cycles
10
50
Proof Cycles
5
8
MEOP Cycles
50
84 cycles to 545 psi
Proof Cycles
10
10
F&D Valves
Cycles
50
50
HP Latch Valve
Cycles
100
2000
LP Latch Valves
Cycles
2000
SDO: 3000
20-lb Thruster Valve
Cycles
1000
25,000
20-lb Thruster Full
Thermal
Cycles
14
14, analyzed to 50 for
HRV
5-lb Thruster Valve
Cycles
100,000
100,000
5-lb Thruster Full
Thermal
Cycles
124
144
Requirement1
Capability
Component
COPV
Propellant Tank
Throughput Requirements
Component
20-lb thruster
Nominal, 100% margin
315 kg
445 kg
5-lb thruster
Nominal, 100% margin
225 kg
246 kg
1. Preliminary estimates, detailed ACS modeling required to determine refine requirements
Charles Zakrzwski - Propulsion
18
Reliability
Failure Rates and Percentages
Component
Tanks (3)
Pres. Pyro Valves
Propellant Filter
Latch Valve
Regulator
Pressure Transducer
Fill and Drain Valve
Propellant Filter
Latch Valve
Thrusters, 20 N (4)
Latch Valve
Thrusters, 20 N (4)
Latch Valve
Thrusters, 80 N (2)
Latch Valve
Thrusters, 80 N (2)
Tank Heater (8 Redunda
Tank Thermistor (8 Redu
Total Failure Rate =
HRSDM EM Propulsion
Component
Reliability
0.999806
0.997897
0.998968
0.996429
0.994903
0.999789
0.997897
0.998968
0.961673
0.998968
0.961673
0.998968
0.990277
0.998968
0.990277
0.999972
0.999998
Block
Redundancy
1.00000
1.00000
0.999999
0.998454
0.999885
FR Source
HRVDM-DM
HRVDM-EM
Landsat-7
Landsat-7
Triana
Triana
Landsat-7
Landsat-7
Landsat-7
HRVDM-DM
Landsat-7
HRVDM-DM
Landsat-7
HRVDM-DM
Landsat-7
HRVDM-DM
HRVDM-EM
Failure Rates
1.1000E-11
1.9000E-08
2.0600E-07
1.0100E-07
3.5000E-07
5.0000E-07
2.0648E-08
2.0600E-07
1.0100E-07
9.5600E-07
1.0100E-07
9.5600E-07
1.0100E-07
9.5600E-07
1.0100E-07
9.5600E-07
1.8300E-07
4.6800E-08
1.4484E-06
Percent of Total
0.002%
0.000%
14.223%
0.007%
24.165%
34.521%
1.426%
14.223%
10.452%
0.780%
0.189%
0.012%
100.00%
0.98531
Charles Zakrzwski - Propulsion
19
Trade Studies - Major Trades Since SRR
Trade
Options Considered
Option Chosen
Rationale
Previous Configurations
Propellant Tank
3 tanks/Single Cylindrical
tank
Single Cylindrical Best mechanical packaging
Tank
Lowest mass solution
Nutation Time
Constant (NTC)
Monoprop with single
tank/SolidMonoprop/Bipropellant
SolidMonopropellant
Most deterministic solution to
solve NTC issue
Launch Vehicle
Delta II class with solidmonoprop/EELV with
existing tanks
EELV with
existing tanks
Lowest Risk :
Eliminated NTC issue
Lowest propulsion subsystem
cost and schedule
Current Configuration (EELV with Two TDRSS type tanks)
Insertion
Thrusters
Two 80 N/ Four 80 N
Thrusters
Four 80 N
Thrusters
Lower propellant mass
Shorter burn duration
Increase operational flexibility
Pyro-Valve
SDO type pyros/high
pressure latch valve/ Other
SDO type pyros
Only option which meets
requirements
Charles Zakrzwski - Propulsion
20
Mass and Power
SUBSYSTEM
Propulsion (Dry Mass)
COMPONENTS
Prop Tanks
Pressurant Tanks
20N Thrusters
88N Thrusters
Prop Hardware
Lines and Fittings
Thermal / Secondary Structure
ALLOCATION
(kg)
147.0
72.1
13.1
5.8
2.9
12.7
8.1
32.4
CURRENT
BEST ESTIMATE
(kg)
136.1
70.0
12.7
5.6
2.8
11.0
7.0
27.0
Charles Zakrzwski - Propulsion
MARGIN
(%)
8.0%
3.0%
3.0%
3.0%
3.0%
15.0%
15.0%
20.0%
21
Integration and Test
Components
Build Component
Modules
(HPM, LPM,
Thrusters, F&D)
Tanks
Module Level
Tests (Thermal
Cycling, Vibe)
Integrate
Tanks with
Structure
Integrate HPM
with Structure
Integrate F&D
Module with
Structure
Fabricate and
Test Liquid
Manifold
Close-Out
Welds
Thruster Modules
Fabricate and
Test Gas
Manifold
Integrate LPM
with Structure
Integrate
Thruster
Modules
Thermal H/W
Subsystem
Verification
(Proof, Leak,
Functional)
Install Thermal
H/W
Deliver for S/C
I&T
Charles Zakrzwski - Propulsion
22
Verification - Components
• Component Level Acceptance Tests
– All components will be/have been Acceptance Tested at Vendor
– Additional testing at GSFC as necessary
• Additional thermal cycles to bring component total to 8
• LRO vibration levels higher than HRV
• Sub-Module level Acceptance Tests (if necessary)
– Low and High Pressure Panels
• May test components at sub-module level to reduce I&T time
– Thermal Cycle
– Vibration
Charles Zakrzwski - Propulsion
23
Verification - Components
Component Acceptance Testing Matrix
Physical
Examination
Mass
Volume
Proof Pressure
Pressure Drop
Propellant Tanks
Thrusters
Pressure
Regulator
Pressurant Tank
Filters
Fill/Drain Valves
Pyro Vales
Pressure Xducers
Latch Valves
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
Voltage
X
X
X
Power
X
X
X
Insulation
Resistance
EMI/EMC
X
X
X
Response Time
X
X
X
External Leakage
Internal Leakage
X
X
X
X
X
X
X
X
X
Cleanliness
NDE (Welds, etc.)
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
X
Shock Test
Static Loads Test
Min Response
Frequency
(Sine Sweep)
Random
Vibration
Sine Vibration
Thermal
Cycling
Performance Test
X
X
X
Special Tests
Valve Acc. Tests,
Life Cycle,
Alignment
Pressure overshoot
Shock Out put, Lot
Acc. Tests
Charles Zakrzwski - Propulsion
Pull - In
VoltagBack Relief
24
Verification - Subsystem
•
Engineering Model Test
– Transient Flow (Waterhammer) tests to size flow control orifices
•
Subsystem Acceptance Tests
–
–
–
–
•
Weld Leak Tests
Proof Pressure Test
Component Function Tests
Thruster Alignment
Special Propulsion Tests during S/C I&T
– System Leak Test during Thermal Vacuum
•
Comprehensive Performance Tests (CPT’s) at S/C I&T
• Note: not all tests performed at all CPT’s
–
–
–
–
External Leak Test
Thermal Circuit/Catbed Heater Tests
End-To-End PDE to Propulsion Test
Electrical functional
• Valve response
• Pressure transducer calibration
– Internal valve and regulator leakage
– Diaphragm leakage
Charles Zakrzwski - Propulsion
25
Propulsion Sub-System Hazards
•
Hazard Descriptions
– Rupture and/or leakage of Hydrazine lines or components
– Rupture and/or leakage of helium pressurant tank and associated
lines
– Inadvertent firing of thrusters
•
Worst Case Severity
– Catastrophic
•
Control and Mitigation Strategy
– Structural and Thermal analyses
– Proof and burst pressure testing
– Materials selection for resistance to stress corrosion, compatibility
– Three independent inhibit scheme
Charles Zakrzwski - Propulsion
26
Schedule - Key Deliverables
Date
Status
Peer Review (Post-S/C PDR)
4/26/2006
Propulsion Peer Review to consider
with GN&C Peer Review
Peer Review (Pre-S/C CDR)
9/11/2006
Complete Analysis
(Flow -Steady & Transient, Plume, Etc)
9/29/2006
Complete Integration Procedures
10/16/2006
Receive Pyro-Valve (longest lead component)
4/30/2007
Subsystem Integration
Pyro Valve Specification and SOW
released to CM
Start High Pressure Module
Integration 1/10/2007
Complete Component Modules
5/2/2007
Receive Prop Module Structure
5/1/2007
Complete Propulsion Subsystem Welding
8/7/2007
Start Subsystem Testing
8/10/2007
Deliver Integrated Prop Module to I&T
Flow Analysis, Filter Analysis,
Leakage Analysis Started.
8/31/2007
Charles Zakrzwski - Propulsion
27
Peer Review Summary
• Subsystem Specification Peer Review - 6/15/05
– Requirements for Out-of-House Propulsion Procurement
– Majority of requirements still applicable in current configuration
• Subsystem Peer Review - To be held 4/26/06
– Work to be completed for Subsystem Peer Review
•
•
•
•
•
•
•
Steady State Flow Analysis
Cycle/Life Analysis
Transient Pressure/Temperature Analysis of Pressurization System
Filter/Contamination Analysis
First order Plume Analysis
Revision of Subsystem Level 3 Requirements
Draft of Mechanical/Thermal/Electrical ICD’s
Charles Zakrzwski - Propulsion
28
Propulsion Risk Matrix
Rank &
Trend
5
L
I 4
K
E
L 3
I
H
O 2
O
D
Approach
Risk Title
Risk ID
1
→
M
Pyro valve misfire issues
PROPUL101
2
→
M
80 N Thruster Steady State Fire
Duration
PROPUL102
2
Notes:
1
1
2
1
3
4
5
Risk 1 - Failure of booster charge to ignite
after firing NSI’s during ground tests.
GSFC has working group to address this
issue for SDO.
CONSEQUENCES
Criticality
L&C Trend
Approach
High
↓
Decreasing (Improving)
M – Mitigate
Med
↑
Increasing (Worsening)
W – Watch
Low
→
Unchanged
A – Accept
New since last month
R – Research
*
Risk 2 - Steady state operation of thruster
requires significant amount of heat to be
dissipated from thruster valve to prevent
valve overheating. Currently investigating
thermal interface between thruster and
spacecraft to address issue
Charles Zakrzwski - Propulsion
29
Status - Documents
Title
Status
Lunar Reconnaissance Orbiter Propulsion Subsystem
Specification
Configured 07/21/05
Updated by 02/15/06
431-SPEC-000490
Lunar Reconnaissance Orbiter Performance
Specification for HRV Propulsion Component
Specifications
Draft to Lockheed Martin 12/20/05
To be configured 02/25/06
431-SPEC-000484
Lunar Reconnaissance Orbiter Pryo Valve Specification
Draft Completed 01/23/06
To be configured 02/10/06
431-SPEC-000488
Lunar Reconnaissance Orbiter Low Pressure Isolation
Valve Specification
Draft Completed 01/30/06
To be configured 02/15/06
431-ICD-(TBD)
Lunar Reconnaissance Orbiter Propulsion Subsystem
Mechanical Interface Control Document
Draft by 04/26/06
431-ICD-(TBD)
Lunar Reconnaissance Orbiter Propulsion Subsystem
Thermal Interface Control Document
Draft by 04/26/06
431-ICD-(TBD)
Lunar Reconnaissance Orbiter Propulsion Subsystem
Electrical Interface Control Document
Draft by 04/26/06
Level 3
431-SOW-00017
Level 4
Charles Zakrzwski - Propulsion
30
Wrap-up
•
Subsystem design meets all major requirements driver
– Provides Delta-V to meet maneuver requirements
• 1326 m/sec (894 kg hydrazine)
– Provides adequate thrust for critical lunar orbit insertion maneuver
• Four 80 N insertion thrusters
– Provides sufficient torque capability for attitude control
• Eight 20 N canted attitude control thrusters
– Maximizes use of existing hardware to minimize cost and schedule
• Monopropellant system
• Thrusters, Tanks, Regulators, Fill/Drain Valves, Filters, from HRV
• Propellant capacity of tanks provides for adequate s/c dry mass capability
– Provides acceptable propulsion system volume
• Regulated system allows propellant tanks to be filled to maximum capacity with
with minimum volume
– Required thruster redundancy
• Capability to perform lunar insertion with an insertion thruster failure
• Redundant bank of attitude control thrusters
Charles Zakrzwski - Propulsion
31
LRO PDR
Attitude Control System
James Simpson
February 8, 2006
ACS Team
ACS Analysis
• Joe Garrick – ACS Analysis
Lead
• Oscar Hsu – ACS Analyst
• Phil Calhoun – ACS Analyst
• Roger Chen – ACS Analyst
• Neal Patel - ACS Analyst
GN&C Hardware
• Ken McCaughey – HW Lead
• Jason Badgley – PDE Lead
• Edward Davis – IRU COTR
• Noble Jones – ST COTR
• Kristen Brown – CSS COTR
• Russ Roder – RW Lead
• Nat Gill – RW & PDE Firmware
Engineer
• John Wilson – RW Electronics
Lead
• George Heisey – GDS Lead
• Joe Vitale – GDS Engineer
• Linh Nguyen – GDS Engineer
• Tom Correll – GDS Consultant
James Simpson - ACS
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ACS Requirement Flow Down
Higher Level Identifier
Rqmt
MRD-14
GNACS-1
MRD-62
GNACS-3
MRD-62
GNACS-4
MRD-62
GNACS-5
MRD-62
GNACS-6
MRD-62
GNACS-7
MRD-62
GNACS-8
MRD-63
GNACS-9
MRD-63
GNACS-10
Object Text
The LRO ACS shall provide a 3-axis stabilization attitude pointing relative to the
orbit normal and nadir reference as defined in the Lunar Reconnaissance Orbiter
Attitude Control System Coordinate System and Alignment Document (431-TBDTBD).
The LRO ACS shall use an orthogonal, body-fixed coordinate system as defined
in the Lunar Reconnaissance Orbiter Attitude Control System Coordinate System
and Alignment Document (431-TBD-TBD).
The LRO ACS shall use for all reference data an Earth Centered Inertial (ECI),
J2000 reference frame.
The LRO ACS onboard algorithms shall make use of the ECI, J2000 inertially
fixed coordinate system for all reference data.
The ECI reference frame shall be defined based on a J2000 inertially fixed
coordinate system. The ECI coordinate frame is defined in the Lunar
Reconnaissance Orbiter Attitude Control System Coordinate System and
Alignment Document (431-TBD-TBD).
The LRO SC shall normally be in a low Lunar, nadir pointing orbit. The coordinate
system, which establishes the desired normal mission attitude, is the Orbit
Coordinate System (OCS). The OCS coordinate frame is defined in the Lunar
Reconnaissance Orbiter Attitude Control System Coordinate System and
Alignment Document (431-TBD-TBD).
The onboard coordinate system shall be the Body Coordinate System (BCS) and
is fixed in the SC. The BCS coordinate frame is defined in the Lunar
Reconnaissance Orbiter Attitude Control System Coordinate System and
Alignment Document (431-TBD-TBD).
The LRO ACS subsystem analysis shall be conducted using the meter-kilogramsecond (MKS) measurement system.
The LRO ACS subsystem shall make use of quaternions for all attitude and
coordinate system transformation representations whenever possible.
James Simpson - ACS
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ACS Requirement Flow Down
MRD-75
GNACS-11
All ACS control modes shall meet stability margin in accordance with established
Guidance, Navigation and Control (GN&C) design practices with a single set of
gains for the entire range of inertias over the mission lifetime, from Beginning of
Life (BOL) to End of Life (EOL).
3
MRD-84
GNACS-12
3
MRD-84
GNACS-13
MRD-84
GNACS-14
MRD-84
GNACS-15
MRD-84
GNACS-16
The ACS control loops shall be designed such that the phase margin is greater
than or equal to 30 degrees.
The gain margin and modal suppression shall be greater than 6 decibels (db),
with a goal of 12 db.
The controller bandwidth shall be selected to minimize interaction with known
structural modes.
The first structural mode shall be no less that 1 Hertz (Hz) and at least one
decade above the controller bandwidth.
Control system stability margins shall be met throughout the mission life.
MRD-84
GNACS-17
All linear stability analyses, the flexible mode damping ratio shall be assumed to
be 0.001 (0.1%), unless a different number is provided by the structures group.
3
MRD-84
GNACS-18
A one cycle delay shall be assumed for all digital systems when determining
linear stability margins.
The LRO ACS shall support the science objectives of the LRO mission.
3
MRD-84
GNACS-19
MRD-14
GNACS-20
MRD-49
GNACS-21
MRD-84
MRD-74
GNACS-22
MRD-17
GNACS-23
MRD-17
GNACS-24
MRD-10
GNACS-25
MRD-92
GNACS-26
MRD-49
MRD-20
GNACS-27
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The LRO SC shall be in a Nadir/Lunar Pointing orientation during science
operations.
The LRO ACS shall meet the pointing, knowledge and stability requirements at
the instrument interface associated with science taking modes.
3
The LRO science instruments shall not induce disturbances to impact the SC
pointing requirements.
The LRO ACS shall support the non-science objectives of the LRO mission using
thrusters.
The LRO ACS shall maintain orbit performance and unload excess momentum.
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The LRO SC shall nominally be in a Nadir/Lunar pointing orientation for all
thruster modes.
The LRO ACS shall meet the pointing, knowledge and stability requirements
associated with all thruster modes.
In the event of the detection of an anomalous condition that threatens the health
and safety of the SC, the LRO SC shall automatically enter a SC contingency
state commensurate with the anomaly.
James Simpson - ACS
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ACS Requirement Flow Down
MRD-20
GNACS-28
Low power contingency configurations that are able to maintain fine pointing (i.e.,
STs) shall remain in science taking attitude.
All other low power contingencies shall transfer to the coarse attitude safe
pointing.
In the event that the ACS cannot maintain the fine pointing attitude requirements
for science taking for low power or attitude related anomalies, then the SC shall
autonomously transfer to the coarse attitude safe pointing mode.
3
MRD-20
GNACS-29
MRD-20
GNACS-30
MRD-20
GNACS-31
The ACS subsystem shall provide an onboard means of Failure, Detection and
Correction (FDC) for anomalous conditions.
The LRO ACS shall meet the pointing, knowledge and stability requirements
associated with the coarse, non-science taking, sun pointing.
3
MRD-90
GNACS-32
MRD-102
GNACS-33
MRD-102
GNACS-34
MRD-102
GNACS-35
Following separation of the payload fairing, the ACS subsystem shall begin
processing available sensor and actuator data in preparation for initial acquisition.
2
MRD-102
GNACS-36
2
MRD-102
GNACS-37
The transition from Launch state to initial sun pointing shall occur autonomously
after despin and post separation.
The ACS shall establish a power positive attitude with TBD minutes after the
Solar Arrays have reached index position when commanded to sun pointing.
MRD-102
GNACS-38
2
MRD-27
GNACS-39
Power positive shall be defined as 45 degrees between Sun and targeted Sun
line.
The ACS subsystem shall employ the use of wheels and thrusters for nulling tipoff rates and residual despin rates, depending on system momentum level.
The launch state shall be the initial state of operation for the LRO SC from launch
to separation from the Delta II ELV.
The LRO SC shall be in a stowed position with minimal SC components powered.
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MRD-26
MRD-6
MRD-5
MRD-27
GNACS-40
MRD-27
GNACS-41
The ACS shall be capable of absorbing momentum, associated with ELV induced
tip-off rates and residual despin rates, equal to no more than 2.0 degrees per
second (deg/sec), per axis.
The ACS shall employ thrusters to null rates above 2.0 deg/sec to a maximum of
30 deg/sec.
James Simpson - ACS
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ACS Requirement Flow Down
MRD-17
GNACS-42
MRD-62
GNACS-43
MRD-147
GNACS-44
MRD-147
MRD-147
GNACS-45
GNACS-46
MRD-147
GNACS-47
MRD-86
GNACS-48
GNACS-263
MRD-52
GNACS-49
MRD-16
GNACS-50
MRD-16
GNACS-51
MRD-16
GNACS-52
MRD-16
MRD-16
GNACS-53
GNACS-54
MRD-15
GNACS-55
MRD-15
GNACS-56
GNACS-57
MRD-15
GNACS-58
MRD-15
GNACS-59
The ACS shall implement hardware and software capable of managing
momentum for periods of no less than 2 weeks (4 weeks goal) over the life of the
mission.
The ACS shall use SC ephemeris in the ECI, J2000 inertially fixed reference
frame.
The ACS shall provide for a means to utilize the onboard ephemeris supplied by
the Flight Dynamics Facility (FDF).
The ACS shall meet all the orbit determination accuracies.
The LRO onboard SC ephemeris shall be accurate to the following requirements
after 3 days.
The FDF shall uplink a SC ephemeris state at a frequency of 3 days.
The ACS shall model onboard the solar and lunar ephemeris to within needed
accuracies for target attitude and interference purposes, HGA pointing and for
Solar Array pointing.
The FDF shall uplink any ground station ephemeris needed for HGA pointing.
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The ACS algorithms shall not intentionally place the Sun into the field of view
(FOV) of the science instruments.
The LRO SC ACS subsystem shall generate HGA pointing solutions and pointing
commands during the Observing Mode.
HGA commands shall be inhibited during launch phase and initial orbit acquisition
while the HGA system is not deployed or not validated.
3
The ACS shall ensure a hemispherical coverage on the -Z side of the SC, with an
accuracy of TBD degrees.
The half-angle beam width of the HGA shall be 3.0 degrees.
The HGA mechanism shall ensure that perturbations are not significant to
degrade ACS and SC pointing requirements.
The LRO SC ACS subsystem shall generate Solar Array pointing solutions and
pointing commands during all nominal on-orbit modes except contingency
modes.
Solar Array commands shall be inhibited during launch mode, initial acquisition,
while the Solar Array is not deployed.
The ACS shall be responsible for commanding the Solar Array to the index
position during Sun Acquisition Mode.
The ACS shall ensure that the solar arrays track the sun from Beta angles of 0 to
90 degrees, with an accuracy of 5 degrees.
The Solar Array mechanism shall ensure that perturbations are not significant to
degrade ACS and SC pointing requirements.
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James Simpson - ACS
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ACS Requirement Flow Down
MRD-18
GNACS-60
Pointing accuracy is not applicable during the slew maneuvers, but pointing
accuracy shall be enforced at the beginning and end of each slew maneuver
corresponding to the control mode in which the maneuver takes place.
3
MRD-18
GNACS-61
3
MRD-18
GNACS-62
MRD-18
GNACS-63
The slews shall nominally be accomplished on wheels alone, but in the event of
wheel failures, the propulsion system may be used.
The propulsion system shall have the ability to handle what is nominally wheel
slews in the event of wheel failure, with the note that this will shorten the mission
lifetime.
No science taking functions shall be done during the slewing periods.
MRD-18
GNACS-64
3
MRD-18
GNACS-65
MRD-18
GNACS-66
Slew timing shall commence upon successful receipt of a ground command to
the ACS.
The LRO ACS shall provide a 180 degree yaw maneuver sub-mode during the
nominal science taking mode.
The 180 degree yaw maneuver shall take no more than 30 minutes.
GNACS-67
GNACS-68
The LRO ACS shall provide for infrequent 20 degree roll, science related slews
during the nominal science mode.
The 20 degree roll, science slews shall take no more than 10 minutes.
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MRD-12
GNACS-69
The LRO SC shall be placed in an initial transfer orbit of TBD km ± TBD km x
TBD km ± TBD km with an inclination of 90 degrees ± TBD degrees.
3
MRD-12
GNACS-70
3
MRD-12
GNACS-71
The science mission orbit shall be a circular 50 km ± 20 km with an inclination of
approximately 90 degrees.
All modes in the ACS shall be capable of operating in all orbits defined.
MRD-102
GNACS-72
MRD-102
GNACS-73
GNACS-73
MRD-102
GNACS-74
MRD-102
GNACS-75
MRD-84
GNACS-76
3
The LRO SC shall turn on specified ACS hardware at launch needed to supply
telemetry until separation.
The LRO SC shall have the IRU on during the launch phase.
The LRO SC shall turn on all remaining sensors and actuators after separation
and prior to any control mode being entered.
The ACS shall inhibit commanding the HGA and Solar Array until deployed and
the SC is in Sun Acquisition mode.
Omnis shall be used for communication during the launch to separation period,
so the ACS shall not generate HGA commands during this period.
The LRO ACS shall perform Attitude Determination to estimate the 3-axis attitude
of the SC with respect to the ECI, J2000 inertially fixed reference frame.
James Simpson - ACS
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ACS Requirement Flow Down
MRD-84
GNACS-77
The Attitude Determination function shall be capable of being executed in any
flight software (FSW) main processor control mode.
2
MRD-84
GNACS-78
The ACS attitude determination function shall estimate onboard attitude using
filtering techniques, propagation or direct sensor measurement.
3
MRD-27
GNACS-79
The ACS shall be able to handle body rates (and its associated momentum) after
despinning to within 12 deg/sec.
3
MRD-5
MRD-27
GNACS-80
Nulling of post-separation rates shall be initiated autonomously after separation
using wheels for rates at or below 2 deg/sec and thruster above 2 deg/sec.
3
MRD-27
GNACS-81
3
MRD-27
GNACS-82
MRD-27
GNACS-83
The despin/rate nulling phase shall have no unique knowledge, pointing or jitter
requirements.
The timing requirement for completing nulling tip-off and residual despin rates
shall be TBD seconds.
Omnis shall be used for communications during despin and rate nulling, thus
ACS shall not generate HGA commands for this period.
MRD-27
GNACS-84
The Solar Array, once successfully deployed, shall be set to the index position.
2
MRD-27
GNACS-85
2
MRD-8
GNACS-86
The HGA, once successfully deployed, shall be set to its initial home position
corresponding to zero degrees in each gimbal.
The LRO ACS subsystem shall provide a Sun Acquisition mode for whenever
conditions shall not allow control to remain within the other control modes.
MRD-8
GNACS-87
The Sun Acquisition Mode shall meet all pointing, knowledge and stability
requirements associated with the non-science taking, coarse attitude pointing
contingency mode. See GNACS-32, Table 3-3 for numerical specifications.
MRD-8
GNACS-88
The ACS shall employ the following hardware for the Sun Acquisition Mode:
2
MRD-8
GNACS-89
The Sun Acquisition Mode shall have a gyro and gyroless sub-modes.
2
MRD-8
GNACS-90
The Sun Acquisition Mode shall be entered autonomously from any mode or via
ground command.
MRD-26
MRD-6
James Simpson - ACS
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ACS Requirement Flow Down
MRD-8
MRD-8
GNACS-91
GNACS-92
MRD-8
GNACS-93
MRD-8
GNACS-94
MRD-8
MRD-8
GNACS-95
GNACS-96
MRD-8
MRD-8
MRD-14
GNACS-97
GNACS-98
GNACS-99
MRD-14
GNACS-100
MRD-14
GNACS-101
MRD-14
MRD-14
GNACS-102
GNACS-103
MRD-14
MRD-14
GNACS-104
GNACS-105
MRD-14
MRD-14
GNACS-106
GNACS-107
MRD-14
MRD-14
GNACS-108
GNACS-109
MRD-14
GNACS-111
Exit from Sun Acquisition Mode shall be via ground command.
Sun Acquisition shall acquire a power positive orientation, to within 15 degrees of
desired Sun line.
The Sun Acquisition Mode shall be able to maintain the nominal attitude pointing
for an indefinite period up to the manageable system momentum level.
The Omni antennas shall be used for communications during Sun Acquisition
Mode.
The ACS shall not generate HGA commands during this phase.
The Solar Arrays shall be rotated to and maintained at the index position during
Sun Acquisition Mode.
The Sun Acquisition Mode shall have no slewing requirements.
The Sun Acquisition Mode shall operate at a 5 Hz cycle time.
The LRO ACS shall provide an Observing Mode which will meet all requirements
to allow for science taking and calibrations.
Observing Mode shall have a Lunar Nadir, Lunar Nadir Offset, Inertial, or Sun
pointing target attitude.
Observing Mode shall meet all pointing, knowledge and stability requirements as
stated for science taking.
The ACS shall employ the following hardware for the Observing Mode:
The attitude determination function for Observing Mode shall employ a standard
six state Kalman Filter to estimate attitude error and gyro drift rate biases.
The HGA shall be used for communications during Observing Mode.
The ACS shall be required to generate HGA commands as defined in the Lunar
Reconnaissance Orbiter Guidance Navigation and Control Attitude Control
System Hardware Interface Control Document (431-ICD-TBD).
The Solar Arrays shall nominally track the Sun during Observing Mode.
The ACS shall generate the appropriate Solar Array commands as defined in the
Lunar Reconnaissance Orbiter Guidance Navigation and Control Attitude Control
System Hardware Interface Control Document (431-ICD-TBD).
The Observing Mode shall operate at a 5 Hz frequency.
The ACS shall make use of SC, solar and lunar ephemerides to compute the
commanded target attitude orientation during Observing Mode.
The nominal Nadir Pointing target orientation shall align the BCS axis with the
OCS axis.
James Simpson - ACS
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ACS Requirement Flow Down
MRD-14
GNACS-112
MRD-9
GNACS-113
GNACS-141
GNACS-143
MRD-10
GNACS-114
MRD-87
MRD-87
GNACS-115
MRD-87
MRD-87
MRD-87
GNACS-116
GNACS-117
GNACS-118
MRD-87
GNACS-119
MRD-87
GNACS-120
MRD-87
GNACS-121
MRD-87
GNACS-122
MRD-87
MRD-87
MRD-87
MRD-87
GNACS-123
GNACS-124
GNACS-125
GNACS-126
MRD-87
GNACS-127
MRD-87
MRD-97
GNACS-128
GNACS-129
MRD-97
GNACS-130
MRD-97
MRD-97
MRD-97
GNACS-131
GNACS-132
GNACS-133
The maximum slew rate allowed by the ACS Inertial Pointing shall be 5.0
deg/sec, per axis.
The maximum slew rate for any Nadir Offset Pointing shall be 5 deg/sec., per
axis.
The Delta-H Mode shall operate at a 5 Hz cycle time.
Uncompensated system momentum from the science instrumentation shall not
exceed 1 Newton meter seconds (Nms), per body axis.
The ACS subsystem shall provide a Delta-V mode to allow for orbit insertion,
orbit adjusts and stationkeeping maneuvers.
3
The Delta-V Mode shall meet the pointing, knowledge and stability requirements
for non-science taking thruster modes. See GNACS-26, Table 3-2 for pointing
requirements.r
The LRO ACS shall employ the following hardware for the Delta-V mode:
Entry into Delta-V mode shall be via command.
The exit from Delta-V shall be autonomous back to same mode before entry.
3
The Delta-V mode shall meet the attitude pointing accuracies for any length burn
duration.
The Delta-V mode shall make use of the gyro propagated attitude state through
the burns.
The Delta-V mode shall provide thruster on-time information in telemetry for
computing fuel usage.
The Delta-V off-axis accuracy shall not exceed 5 percent, 3-sigma, of the
commanded value.
The Delta-V mode shall be capable of maintaining current wheel speeds.
The Delta-V Mode shall use Omnis for communications.
The ACS shall not generate HGA commands during this period.
The Solar Arrays shall be capable of tracking the Sun during Delta-V Mode.
3
The ACS shall be responsible for generating Solar Array commands during this
period.
The Delta-V Mode shall operate at a 5 Hz cycle time.
The ACS subsystem shall provide a Delta-H Mode to allow for momentum
management using the propulsion system.
The Delta-H Mode shall meet the pointing, knowledge and stability requirements
for non-science taking thruster modes. See GNACS-26, Table 3-2 for pointing
requirements.
The LRO ACS shall employ the following hardware for the Delta-H mode:
Entry into Delta-H shall be via command.
The Delta-H shall exit autonomously when momentum is below 1Nms or when a
time limit is reached.
James Simpson - ACS
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ACS Requirement Flow Down
MRD-97
GNACS-134
MRD-97
GNACS-135
MRD-97
MRD-97
MRD-97
GNACS-136
GNACS-137
GNACS-138
MRD-97
MRD-97
GNACS-139
GNACS-140
MRD-97
GNACS-142
MRD-19
GNACS-144
MRD-16
GNACS-145
MRD-16
GNACS-146
MRD-16
MRD-16
GNACS-147
GNACS-148
MRD-15
GNACS-149
MRD-15
MRD-15
GNACS-150
GNACS-151
MRD-15
GNACS-152
MRD-15
GNACS-153
MRD-15
GNACS-154
MRD-20
MRD-20
GNACS-155
GNACS-156
MRD-20
GNACS-157
The Delta-H mode shall propagate the attitude state through the burns using IRU
rate information.
The Delta-H mode shall provide thruster on-time information in telemetry for
computing fuel usage.
The ACS shall actively manage wheel speeds during Delta-H mode.
The Delta-H Mode shall use Omnis for communications
The ACS shall not be responsible for generating HGA commands during this
period.
The Solar Arrays shall capable of tracking the Sun during Delta-H mode.
The ACS shall be required to generate Solar Array commands during Delta-H as
defined in the Lunar Reconnaissance Orbiter Guidance Navigation and Control
Attitude Control System Hardware Interface Control Document (431-ICD-TBD).
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Algorithms for momentum redistribution of the reaction wheels shall be
implemented in the wheel based control modes.
The LRO ACS shall implement the necessary algorithms to ensure the accuracy
requirements are met for all eclipse periods when sun sensor measurements are
not available.
The ACS commands to the HGA shall be absolute pointing and rate commands.
3
The current gimbal positions provided by the HGA to the ACS shall be within TBD
degrees of accuracy.
Commands shall be inhibited when the HGA is stowed.
The LRO ACS shall rotate the HGA to the expected position for the next ground
contact.
The ACS commands to the Solar Array shall be absolute pointing and rate
commands.
Commands shall be inhibited when the Solar Array is stowed.
The index position for the Solar Array shall be put the panel normal vector
parallel with the -Y body axis.
The Solar Array shall provide an indication to the ACS subsystem that the panel
has reached its proper index position.
This indicator shall be independent of any array position encoder or
potentiometer.
The LRO ACS shall on entry to an eclipse period rotate the solar arrays to the
position expected for when the SC exits the eclipse.
The ACS subsystem shall provide an onboard means of FDC.
The FDC algorithms shall be capable of detecting non-nominal performance from
any of the ACS related hardware and the FSW.
The LRO SC in general is single string, however, the ACS shall employ selective
redundancy when needed.
2
James Simpson - ACS
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ACS Requirement Flow Down
MRD-20
MRD-20
GNACS-158
GNACS-159
MRD-20
GNACS-160
MRD-20
GNACS-161
MRD-85
GNACS-162
MRD-85
MRD-85
GNACS-163
GNACS-164
MRD-85
MRD-85
GNACS-165
GNACS-166
MRD-85
GNACS-167
MRD-85
GNACS-168
MRD-85
GNACS-169
MRD-85
The ACS shall employ a single fault recovery in the FDC.
The FDC shall take autonomous actions, when necessary, to maintain the health
and safety of the SC.
The FDC shall provide through telemetry indication of any non-nominal
performance detected in software or hardware.
There shall be no autonomous promotion of modes after FDC detects anomalies
that demotes current mode.
The IRU shall provide 3-axis angular rate and delta angle information to the ACS.
The IRUs shall nominally be aligned with the BCS reference frame.
Each IRU shall be aligned relative to the BCS reference frame to within 0.1 deg,
per axis, and form a right handed orthogonal system.
The IRU alignment knowledge shall be known to within 5 arcsec, per axis.
The IRU alignments shall be defined in the Lunar Reconnaissance Orbiter
Guidance Navigation and Control Attitude Control System Specification (431SPEC-TBD).r
The format for all IRU data to the ACS shall be defined in the Lunar
Reconnaissance Orbiter Guidance Navigation and Control Attitude Control
Hardware Interface Control Document (431-ICD-TBD).
The IRU shall have a saturation rate limit of no less than 18.0 deg/sec, per axis.
3
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2
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3
GNACS-170
The IRU shall provide information up to the survival rate level of 420.0 deg/sec
(70 revolutions per minute [RPM]).
The IRU shall provide a correct polarity indicator up to the survival rate level.
3
3
MRD-85
GNACS-171
The IRU shall have an accuracy of 1 arcsec/count for accumulated delta angles.
3
MRD-85
GNACS-172
3
MRD-85
GNACS-173
The IRU shall have a sigma-u value of no greater than 1.0E-03
arcsec./(sec^(3/2))
The IRU shall have a sigma-v value of no greater than 0.6 arcsec./(sec^(1/2))
MRD-85
MRD-85
GNACS-174
GNACS-175
MRD-85
MRD-85
GNACS-176
GNACS-177
MRD-85
GNACS-178
MRD-85
GNACS-179
3
The IRU shall have a sigma-e value of no greater than 1.0 arcsec.
The IRU shall have a Noise Equivalent Angle (NEA) of no greater than 0.2
arcsec, 3-sigma.
The IRU data shall be sampled from the IRU at a frequency of TBD Hz.
The IRU data shall be provided to the ACS at a frequency of no less than 5 Hz.
3
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The IRU shall provide to the ACS status flags as defined in the Lunar
Reconnaissance Orbiter Guidance Navigation and Control Attitude Control
Hardware Interface Control Document (431-ICD-TBD).
The CSS shall provide a signal proportional to the cosine of the angle of the Sun
from the CSS boresight.
3
James Simpson - ACS
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ACS Requirement Flow Down
MRD-85
MRD-85
GNACS-180
GNACS-181
MRD-85
GNACS-182
MRD-85
GNACS-183
MRD-85
GNACS-184
MRD-85
GNACS-185
MRD-85
GNACS-186
GNACS-261
MRD-85
GNACS-187
MRD-85
GNACS-188
MRD-84
MRD-85
GNACS-189
MRD-85
GNACS-190
MRD-85
GNACS-191
MRD-85
GNACS-192
MRD-85
GNACS-193
The accuracy of each CSS shall be to within 1 degree.
The CSS FOV shall nominally be an 85 degree half cone angle about the
boresight.
Each CSS boresight shall be aligned according to its specified FOV in the BCS
frame to within 1.0 degrees.
The knowledge of each CSS boresight alignment shall be known to within 1
degree.
The CSS alignments shall be defined in the Lunar Reconnaissance Orbiter
Attitude Control System Coordinate System and Alignment Document (431-TBDTBD).
The ACS subsystem shall employ eight CSS total to allow for as much spherical
coverage as possible and for redundancy.
The CSS data shall be provided to the ACS at a frequency of no less than 5 Hz.
The CSS output shall approximate the cosine curve from 0 to 85 degrees such
that the error between the CSS output and the true cosine of the angle from the
boresight is better than 0.5 degrees when within 10 degrees of the boresight,
better than 1.0 degrees when within 20 degrees of the boresight, better than 2.0
degres when within 40 degrees of the boresight and otherwise better than 8
degrees when within the FOV.
The STs shall output an ECI-to-ST quaternion in the J2000 inertial reference
frame.
There shall be two STs to ensure that the ACS onboard knowledge is maintained
during occultations.
3
3
3
3
3
2
2
3
3
3
The ST shall have an internally supplied and maintained star catalogue with the
ability to provide an attitude solution for any attitude orientation on the celestial
sphere.
The ST random errors shall not exceed TBD arcsec, 3-sigma, for the ST
boresight axis and TBD arcsec, 3-sigma, for each ST transverse axis at EOL.
3
The ST uncompensated bias (systemic) errors shall not exceed TBD arcsec, 3sigma, in the ST boresight axis and TBD arcsec, 3-sigma, for each ST transverse
axis at EOL.
Alignment of each ST shall be defined in the Lunar Reconnaissance Orbiter
Attitude Control System Coordinate System and Alignment Document (431-TBDTBD).
The alignment of each ST relative to the Master Reference Cube (MRC) and the
BCS reference frame shall be within 5 arcsec, per axis.
3
James Simpson - ACS
3
2
2
13
ACS Requirement Flow Down
MRD-85
GNACS-194
MRD-85
GNACS-195
MRD-85
GNACS-196
MRD-85
GNACS-197
MRD-85
GNACS-198
MRD-85
GNACS-199
GNACS-199
The alignment knowledge of each ST relative to the MRC shall be known to
within 5 arcsec, per axis.
Each ST alignment shall be calibrated during the trans-lunar, sun pointing phase.
The ST shall employ star observations from the Lunar Reconnaissance Orbiter
Camera (LROC) science instruments if possible, for calibration of the ST relative
the LROC boresight.
The STs shall provide to the ACS status flags as defined in the Lunar
Reconnaissance Orbiter Guidance Navigation and Control Attitude Control
System Hardware Interface Control Document (431-ICD-TBD).
The format for the quaternions and other ST data shall be defined in the Lunar
Reconnaissance Orbiter Guidance Navigation and Control Attitude Control
System Hardware Interface Control Document (431-ICD-TBD).
Each ST FOV shall be unobstructed by any SC structure, hardware or
instrumentation.
Each ST shall have, if needed, a sun shield that restricts light impingement on
the ST and have a radial FOV of TBD degrees.
The ST data shall be output data at a frequency of no less than 5 Hz.
The ST data shall be provided to the ACS at no less than 5 Hz.
The ST shall output a timetagged quaternion that has timing accuracy of no
greater than TBD msec.
The LRO ACS subsystem shall employ four reaction wheels for momentum
storage and attitude control.
2
1
3
3
3
2
3
MRD-85
GNACS-200
GNACS-201
GNACS-262
MRD-85
GNACS-202
MRD-84
MRD-85
MRD-85
GNACS-203
GNACS-204
The RWA shall provide tachometer information to the ACS.
The definition and format of the tachometer data for each wheel shall be defined
in the Lunar Reconnaissance Orbiter Guidance Navigation and Control Attitude
Control System Hardware Interface Control Document (431-ICD-TBD).
3
3
MRD-85
GNACS-205
3
MRD-85
GNACS-206
MRD-85
GNACS-207
The RWA tachometer data shall be provided to the ACS with an accuracy of
better than 0.01 rad/sec and a resolution better than 1E-04 rad/sec.
The alignment of the RWA shall be defined in the Lunar Reconnaissance Orbiter
Attitude Control System Coordinate System and Alignment Document (431-TBDTBD).
Each wheel spin axis shall be aligned to within an accuracy of 1.0 deg of the
desired spin axis in the BCS frame, with an alignment knowledge of 5 arcsec.
MRD-85
GNACS-208
MRD-85
GNACS-209
The ACS shall command the RWA by torque commands to each wheel. The
definition and format of the torque commands are in the Lunar Reconnaissance
Orbiter Guidance Navigation and Control Attitude Control System Hardware
Interface Control Document (431-ICD-TBD).
Each wheel shall deliver maximum values of no less than 0.16 Nm and 80 Nms
at minimum SC voltage.
James Simpson - ACS
3
3
3
2
2
3
3
3
14
ACS Requirement Flow Down
MRD-85
MRD-85
GNACS-210
GNACS-211
Each wheel shall have a static imbalance of no greater than TBD g-cm.
Each wheel shall have a dynamic imbalance of no greater than TBD g-cm^2.
3
3
MRD-85
GNACS-212
3
MRD-85
GNACS-213
The RWA manufacturer shall provide a torque-speed capability curve to the ACS
engineers.
The RWA manufacturer shall provide a drag torque curve to the ACS engineers.
3
MRD-85
GNACS-214
The RWA shall output tachometer data at a frequency of no less than 5 Hz.
3
MRD-85
GNACS-215
3
MRD-85
GNACS-216
MRD-87
GNACS-217
The RWA tachometer data shall be provided to the ACS at a frequency of no less
than 5 Hz.
The ACS shall provide RWA torque commands to each wheel at a frequency of 5
Hz.
The ACS shall employ 10 thrusters (8, 5 pound force [lbf] attitude/orbit and 2, 20
lbf) orbit insertion) for use in the orbit adjust/maintenance and momentum
unloading control modes.
MRD-84
MRD-85
MRD-85
GNACS-218
MRD-85
GNACS-219
MRD-85
GNACS-220
MRD-85
GNACS-221
MRD-85
GNACS-222
MRD-85
GNACS-223
MRD-85
GNACS-224
MRD-85
GNACS-225
MRD-16
GNACS-226
The Propulsion Subsystem shall employ enough thrusters to meet ACS
redundancy requirements, in case of any single thruster failure.
Each attitude thruster shall have a minimum force capability of 18 Newtons to
meet ACS attitude control requirements.
Each orbit maintenance thruster shall have a minimum force capability of 80
Newtons to meet the ACS orbit insertion burn requirements.
The ACS shall issue thruster on/off and pulse width commands to the Propulsion
PDE box, the format defined in the Lunar Reconnaissance Orbiter Guidance
Navigation and Control Attitude Control System Hardware Interface Control
Document (431-ICD-TBD).
The location and alignment of the thrusters shall be defined in the Lunar
Reconnaissance Orbiter Attitude Control System Coordinate System and
Alignment Document (431-TBD-TBD).
The minimum impulse bit for each thruster shall not be greater than TBD msec.
The Command and Data Handling (C&DH) shall be responsible for valve
openings and closings, heater turning on and off, etc., not the ACS, in
preparation for each thruster burn.
The ACS shall issue thruster on/off commands at a frequency of no less than 5
Hz.
The ACS shall provide gimbal, rate and operational commands to point the High
Gain Antenna System (HGAS) as defined in the Lunar Reconnaissance Orbiter
Guidance Navigation and Control Attitude Control System Hardware Interface
Control Document (431-ICD-TBD).
James Simpson - ACS
3
2
3
3
3
3
3
3
2
2
2
15
ACS Requirement Flow Down
MRD-16
GNACS-227
The HGA shall provide to the ACS current gimbal angles, a home direction
indicator, an index reached indicator, status flags, and other data if required as
defined in the Lunar Reconnaissance Orbiter Guidance Navigation and Control
Attitude Control System Hardware Interface Control Document (431-ICD-TBD).
MRD-16
GNACS-228
MRD-16
GNACS-229
The HGAS shall not introduce disturbances that will affect science pointing and
stability requirements.
The HGA data shall be provided to the ACS at a frequency of no less than 5 Hz.
MRD-16
GNACS-230
MRD-15
GNACS-231
MRD-15
GNACS-232
MRD-15
GNACS-233
MRD-15
GNACS-234
MRD-15
GNACS-235
MRD-86
GNACS-236
MRD-86
GNACS-237
MRD-86
GNACS-238
MRD-86
GNACS-239
MRD-86
GNACS-240
MRD-86
GNACS-241
The ACS shall issue HGAS gimbal commands at a frequency of no less than 5
Hz.
The ACS shall provide to the Solar Array angles and rate commands to point the
panel toward the sun as defined in the Lunar Reconnaissance Orbiter Guidance
Navigation and Control Attitude Control System Hardware Interface Control
Document (431-ICD-TBD).
The Solar Array shall provide to the ACS current gimbal angles, a home direction
indicator, an index reached indicator, status flags, and other data if required as
defined in the in the Lunar Reconnaissance Orbiter Guidance Navigation and
Control Attitude Control System Hardware Interface Control Document (431-ICDTBD).
The Solar Array data shall be provided to the ACS at a frequency of no less than
5 Hz.
The ACS shall issue Solar Array gimbal commands at a frequency of no less
than 5 Hz.
The Solar Array shall not introduce disturbances that will affect science pointing
and stability requirements.
The ACS subsystem shall deliver to the FSW subsystem an ACS Algorithms
Document sufficient in definition needed to implement all ACS attitude related
software.
The FSW shall provide for all commanding needed for the ACS modes as
defined in the Lunar Reconnaissance Orbiter Flight Software Specifications (431SPEC-TBD) document.
The ACS shall define all commands needed for the onboard control algorithms.
The FSW shall provide for all telemetry needs for ACS validation and ground
processing as defined in the Lunar Reconnaissance Orbiter Flight Software
Specifications (431-SPEC-TBD) document.
The ACS shall define telemetry packets for all ACS related hardware and control
algorithms.
The FSW shall output all sensor and actuator telemetry packets as defined in the
Lunar Reconnaissance Orbiter Flight Software Requirements (431-RQMT000139) document.
James Simpson - ACS
2
3
2
2
2
2
2
2
3
2
2
2
2
2
2
16
ACS Requirement Flow Down
MRD-86
GNACS-242
MRD-86
GNACS-243
MRD-86
GNACS-244
MRD-86
GNACS-245
MRD-86
GNACS-246
MRD-86
GNACS-247
MRD-86
GNACS-248
MRD-86
GNACS-249
MRD-86
GNACS-250
MRD-86
GNACS-251
MRD-86
MRD-86
GNACS-252
GNACS-253
MRD-86
GNACS-254
MRD-86
GNACS-255
MRD-62
GNACS-256
MRD-96
MRD-75
GNACS-257
The ACS shall define to FSW all necessary telemetry points and packets needed
to adequately access the performance of the ACS hardware.
The time used within the ACS shall be taken from the onboard clock in the C&DH
and be in the Universal Time Coordinate (UTC) format.
The C&DH shall provide with all data input to the ACS a time stamp equal to
when it was made available to the ACS.
The FSW shall provide for the onboard interpolation of the SC ephemeris state.
3
The FSW onboard SC ephemeris interpolator shall meet all accuracy
requirements as stated in previous sections.
The FSW shall provide for an onboard Solar and Lunar ephemeris model as
specified in the Lunar Reconnaissance Orbiter Flight Dynamics Specifications
(431-SPEC-TBD) document.
The FSW shall make the onboard ephemeris data available to the ACS at a
frequency of no less than 5 Hz.
The FSW shall implement the algorithms defined by ACS for pointing the HGA to
within required accuracies.
The FSW shall implement algorithms defined by ACS for the Solar Array to point
to within the required accuracies.
The FSW shall implement the control modes as defined by the Lunar
Reconnaissance Orbiter Attitude Control System Flight Software Algorithm
Document (431-SW-TBD).
The FSW shall implement control modes to operate at a 5 Hz cycle time.
The FSW shall implement the FDC algorithms as defined by the ACS in Lunar
Reconnaissance Orbiter Attitude Control System Flight Software Algorithm
Document (431-SW-TBD).
The ACS subsystem shall develop a high fidelity simulation to verify all attitude
related performance requirements for all configurations throughout the SC’s
mission lifetime.
The ACS subsystem shall deliver to the LRO Dynamics Simulator algorithms,
definitions, software or whatever is agreed to, necessary to implement a hybrid
simulation involving the FSW and/or SC ACS hardware.
2
The placement, alignment, FOVs and rotation matrices for each sensor, actuator
and hardware used by the ACS shall be defined in the Lunar Reconnaissance
Orbiter Attitude Control System Coordinate System and Alignment Document
(431-TBD-TBD).
The Mechanical subsystem shall provide the necessary mass properties, center
of pressure vectors, center of mass vectors, and other specified parameters, as
defined by the ACS subsystem at prescribed deliverable dates agreed to between
the subsystems.
James Simpson - ACS
2
2
2
2
2
2
2
2
2
2
3
3
2
2
17
ACS Requirement Flow Down
MRD-72
GNACS-258
MRD-96
GNACS-259
MRD-85
GNACS-260
The Mechanical subsystem shall keep the first mode of all SC structural and
hardware components to be no less than 0.5 Hz.
The Mechanical subsystem shall provide a Lunar Reconnaissance Orbiter
Mechanical Systems Specification (431-SPEC-000012) document that defines all
coordinate systems, units and definitions for all deliverables to ACS.
The ACS shall specify to the FDF the operational considerations needed to
sufficiently calibrate the onboard attitude sensors.
All requirements in this document shall be verified by one of the four methods
defined below.
James Simpson - ACS
2
2
2
18
ACS Subsystem Summary
•
Attitude Control System will determine spacecraft attitude, guidance to reach the
desired pointing vector, and use actuators to achieve the desired pointing vector
– Per Level 3 ACS requirements
•
•
Pointing support for High Gain Antenna and Solar Array
ACS consists of
– Sensors
• Star Trackers (2)
• Inertial Reference Unit (1)
• Coarse Sun Sensors (10)
– Actuators
• Reaction Wheels (4)
• 20 N ACS thrusters (8)
– Attitude Control Electronics
• Propulsion Deployment Electronics (PDE) component will provide thruster control as well as
inhibit during launch
•
•
GN&C Flight software (FSW) will be executed within the C&DH Single Board
Computer (SBC)
Testing of FSW will be supported by GSFC Dynamic Simulator
James Simpson - ACS
19
LRO ACS Configuration
Omnis
Mini-RF
ACS Responsibility
Low-Rate
Cmds & Tlm
LROC
S-Xpndr
SpaceWire Network
LAMP
ACS Support
Hi-Rate
Tlm
LAMP Sci. & HK
20MHz
LOLA
USO 9500
ATA
C&DH
HGA
Ka-Xmtr
DDA
Thermistors
Closed Loop
Htrs
ST(2)
HGA
Gimbals
LEND
IRW(4)
CSS(10)
Diviner
Unsw. + 28V
GIMBAL
CONROL
IMU
MIL-STD-1553 Network
CRaTER
Battery
Solar
Array
SA
Gimbals
Sw. and
Unsw.
+28V Pwr
Services
PSE
PDE
GIMBAL
CONTROL
Vehicle
Separation
Break Wires
James Simpson - ACS
Propulsion
SA & HG
Deploy
Actuation
20
Configuration
Star Trackers
+X
S/C Body Frame
+Y
+Z
Reaction Wheels
James Simpson - ACS
21
Configuration
Inertial Reference Unit
+Z
+X
+Y
S/C Body Frame
Propulsion Deployment Electronics
James Simpson - ACS
22
ACS Mode Diagram
Power-On/Reset
Cmd
Sun-Safe
Sun-Safe
• Manage sun relative to S/C
• Wheels, CSSs, IMU-optional
• SA in predefined position
• HGA in predefined position
Cmd,
Cmd
Safing
Observing
• Nadir, Inertial, Offset pointing
• Wheels, IMU, STs
• SA tracking Sun
• HGA tracking Earth
Auto, Cmd,
Safing
Delta-H
Delta-H
• Hold attitude, unload mom.
• Thrusters, IMU
• SA in predefined position
• HGA in predefined position
Cmd
Auto, Cmd,
Safing
Delta-V
• Hold attitude, adjust velocity
• Thrusters, IMU, STs
• SA in predefined position
• HGA in predefined position
Cmd
Observing
Delta-V
Auto, Cmd,
Safing
James Simpson - ACS
23
Sun Safe Spacecraft Configuration
+Y
+X
S/C Body Frame
+Z
Sun Line
2 body mounted CSS locations
on –Y side of spacecraft
4 CSS locations on array
Four body mounted CSS locations along +Y side of spacecraft
James Simpson - ACS
24
Star Trackers
• Spacecraft Attitude Determination
device
• Performance Specifications
Summary
– Boresight Error (3σ) at EOL
• 30 arcsec systematic and bias errors with
rates between 0.0 and 0.3 deg/sec
• 120 arcsec random errors with rates
between 0.0 and 0.3 deg/sec
– Transverse Error (3σ) at EOL
• 11 arcsec systematic and bias errors with
rates between 0.0 and 0.3 deg/sec
• 36 arcsec random errors with rates
between 0.0 and 0.3 deg/sec
James Simpson - ACS
25
Inertial Reference Unit
• Attitude rate measuring
device
• Used on other GSFC
missions
• AIDR
– Absolute Value of ≤ 5
arcsec/sec (3σ) over 10 mins
– Short Term Stability of ≤ 0.1
arcsec/sec (3σ) over 6 hours
– Long Term Stability of ≤ 1.0
arcsec/sec (3σ) over 30 days
• Scale Factor
– 300 ppm (Linear) & 150 ppm
(Asymmetry
• NEA
– 0.4 arcsec/sec/g (3σ)
James Simpson - ACS
26
Coarse Sun Sensor (CSS)
• Passive devices used to
determine spacecraft
attitude with respect to the
Sun
• Current output of each CSS
is equivalent to a sine
function where the peak of
the sine wave shows the
Sun on the CSS boresight.
• Accuracy to within one
degree
• Used only in Sun Safe
mode and FDC
James Simpson - ACS
27
ACS Hardware Procurement Status
• Star Tracker
– Request for Information (RFI) completed in December
– Statement of Work (SOW) ready for Contracting Officer
– Deliverables Item List and Schedule (DILS) ready for Contracting
Officer
– Specification is completing final CM review
• Inertial Reference Unit
– Sole Source Synopsis completed on Jan. 25, 2006 with no protests
– SOW, DILS, and Specification all under final CM review
• CSS
– RFI completed in January 2006
– RFP period started January 30, 2006
James Simpson - ACS
28
Reaction Wheels (RW)
•
•
•
•
•
•
RW bearing design based on SWAS, TRACE, & WIRE RW
designs
– Life testing for over 11 years
Triana RW electronics & lubricant are same
– Life testing for over 4 years
James Simpson - ACS
LRO has chosen to build RWs
in-house
RW is based on the on-going
GPM RW design
The GPM RW ETU is
currently undergoing followup vibe testing
– Results will be valid for
LRO
RW High Level Specs
– 0.16 Nm of torque
– 80 Nms of Momentum
Capacity
29
ACS Mass and Power Budget
SUBSYSTEM
ACS
COMPONENTS
Star Trackers
Inertial Measurement Unit
Reaction Wheels
Coarse Sun Sensors
ALLOCATION
(kg)
64.5
6.6
5.0
52.8
0.15
James Simpson - ACS
CURRENT
BEST ESTIMATE
(kg)
58.6
6.0
4.5
48.0
0.12
MARGIN
(%)
10.0%
10.0%
10.0%
10.0%
25.0%
30
PDE Power Budget
SUBSYSTEM
PDE
COMPONENTS
Control Modules
Inhibit control module
ALLOCATION
(kg)
15.4
12.3
3.1
James Simpson - ACS
CURRENT
BEST ESTIMATE
(kg)
14.0
11.2
2.8
MARGIN
(%)
10.0%
10.0%
10.0%
31
Propulsion Deployment Electronics (PDE)
Card
Aluminum construction
100 mil thickness
No backplane
Supports the addition of more cards
• Designed to fire thrusters in
support of any Delta-V or
Delta-H maneuver.
• Also contains inhibit relays
used to prevent unsafe
operations from occurring until
desired
– RF transmitters and
Propulsion system inhibited
until launch vehicle
separation
– Also notifies FSW via 1553
Inhibit
Unit
of launch vehicle separation
status
James Simpson - ACS
32
PDE Driving Requirements
PDE Req.†
Requirement
3.2.1.1
Independent redundant interfaces within the propulsion and
deployment system shall be continued to be independently
redundant within the PDE.
3.2.3.1
The PDE shall provide 3 independent electrical inhibits for
deployables.
3.2.3.2
The PDE shall provide 2 independent electrical inhibits for RF
transmitters.
3.2.5.3
The PDE shall be able to operate nominally following the
permanent shorting failure of a FET immediately following a
NSI firing.
3.2.6.1
Two 1553 commands shall be required to control any single
switch within the PDE.
†
PDE Requirements Document Number 431-RQMT-000421
James Simpson - ACS
33
Mission Critical Design Drivers
• Solar Array Deployment is mission critical
– Charging of the batteries is mission critical once on orbit
– Stowed solar array only has a third of its cells exposed
• Lunar Capture is Time Critical
– The start time for the Delta-V maneuver is critical
– Interruption of the Delta-V maneuver could be
catastrophic
• Minimize Power Consumption
– Goal is to turn off PDE between thruster burns
– Thruster burns occur roughly every two weeks
James Simpson - ACS
34
PDE Propulsion Interface
Reference Designators
Pressure
Tank
PVx-y = Pyro Valve #x, NSI #y
LVx = Latch Valve #x
T20N-x = 20N Thruster #x
T80N-x = 80N Thruster #x
PV1
PV2
Note: This simplified propulsion
schematic only shows the
elements relevant to PDE.
LV1
Propellant
Tank
LV6
Propellant
Tank
LV7
LV2
T20N-1
T20N-2
T20N-3
LV4
LV3
T20N-4
T20N-5
T20N-6 T20N-7 T20N-8
James Simpson - ACS
LV5
T80N-1
T80N-2
T80N-3
T80N-4
35
PDE Propulsion Requirements
• Actuators controlled by PDE
–
–
–
–
–
80 N thrusters (4)
20 N thrusters (8)
Dual coil, low pressure latch valves (6)
Dual coil, high pressure latch valves (1)
NASA Standard Initiators (4)
• 2 NSIs per Pyrotechnic Valve
• Telemetry read by PDE
– Latch valve status (7)
James Simpson - ACS
36
PDE Deployment Interfaces
• Actuators controlled
by PDE
SOLAR
ARRAY
– High Gain Antenna
RESTRAINT-1
• Separation Nuts (2)
(SAR1-1,2)
– NSIs (4)
RESTRAINT-2
(SAR2-1,2)
HIGH GAIN
ANTENNA
RESTRAINT-3
(SAR3-1,2)
RESTRAINT-1
– Solar Array
• Separation Nuts (4)
– NSIs (8)
(HGAR1-1,2)
RESTRAINT-4
RESTRAINT-2
(SAR4-1,2)
(HGAR2-1,2)
• Telemetry read thru
MAC in C&DH box
Note: NSIs are fired one at a time.
James Simpson - ACS
37
PSE
Output
Modules
Mod D Controller Electronics
SPARE
1A
15A
Mod D Load Bus
Mod C Controller Electronics
Mini-RF Transmitter
1A
PSE-OPM4
Mod C Load Bus
PSE-OPM3
5A
1A
15A
Mod B Load Bus
Mod B Controller Electronics
?A
1A
Ka-Band RF Transmitter
Separation Connector J1
Mod A Load Bus
SBW-1
SBW-2
SBW-3
SBW-4
Separation Connector J2
PSE-OPM2
Mod A Controller Electronics
2A
S-Band RF Transmitter
SBW-5
SBW-6
SBW-7
SBW-8
Separation
Break
Wires
15A
PSE-OPM1
PSE
15A
PDE Interface Block Diagram
Relay
Drivers
K4-3 K3-3
K6-3 K5-3
RELAY STATE MONITOR - TO GSE
PDE Inhibit Unit
SPARE
RF1
RF2
RF3
S-Band
Ka-Band
Mini-RF
PDE
Cards
1553 Bus
PDE
Modules
PDE
Mod-A
PDE
Mod-B
LVS-4
LVS-5
SBW-9
Actuator
Loads
(Thrusters)
Relay
Inhibit
Verification
Interface
K8-3 K7-3
K6-1
K5-1
K4-1
K7-1
K7-2
K8-2
K3-1
K2-1
K8-1
Inhibit
Relays
RELAY RESET - TO GSE
GSE I/F
To Reset
Relays
K2-3 K1-3
1Shot
K6-2
1Shot
K5-2
1Shot
K4-2
1Shot
K3-2
1Shot
K2-2
1Shot
K1-2
1Shot
K1-1
1Shot
PV1-1
SAR1-1
SAR3-1
HGAR1-1
SPARE
T20N-1
T20N-2
T20N-3
T20N-4
P LV4A C
PV1-2
SAR1-2
SAR3-2
HGAR1-2
SPARE
PDE
Mod-C
LVS-2
LVS-6
SBW-10
P LV2A C
P LV3A C
P LV6A C
PV2-1
SAR2-1
SAR4-1
HGAR2-1
SPARE
PDE
Mod-D
LVS-1
SPARE
SPARE
T20N-5
T20N-6
T20N-7
T20N-8
P LV4B C
PV2-2
SAR2-1
SAR4-2
HGAR2-2
SPARE
LVS-7
LVS-3
SPARE
P LV2B C
P LV3B C
P LV6B C
P LV5A C
P LV7A C
P LV5B C
P LV7B C
P LV1A C
T80N-1
T80N-2
P LV1B C
T80N-3
T80N-4
James Simpson - ACS
38
PDE Actuator Mapping
LVS-4
LVS-5
SBW-9
A
Signal
Input
FPGA
1553
PSE-OPM1
1A (2 pair)
DC/DC
PSE-OPM1
15A (4 pair)
Load Bus
LVS-1
SPARE
SPARE
Signal
Input
C
FPGA
1553
PSE-OPM3
1A (2 pair)
DC/DC
PSE-OPM3
15A (4 pair)
Load Bus
1.2A
1.2A
1.2A
1.2A
T20N-1
T20N-2
T20N-3
T20N-4
P LV2A C
1.2A
1.2A
1.2A
1.2A
12A
P LV4A C
P LV6A C
P LV5A C
P LV7A C
P LV1A C
T80N-1
T80N-2
PV1-2
SAR1-2
SAR3-2
HGAR1-2
P LV3A C
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
12A
12A
12A
12A
12A
12A
PV1-1
SAR1-1
SAR3-1
HGAR1-1
12A
SPARE
SPARE
12A
1.2A
T20N-5
P LV2B C
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
T20N-6
T20N-7
T20N-8
P LV4B C
1.2A
1.2A
12A
12A
12A
P LV5B C
P LV7B C
P LV1B C
PV2-1
T80N-3
T80N-4
PV2-2
12A
12A
12A
12A
SAR2-1
SAR4-1
HGAR2-1
SAR2-2
SAR4-2
HGAR2-2
SPARE
Reference Designators
HGARx-y - High Gain Antenna Release #x, NSI #y
LVx = Latch Valve #x (P = open coil, C = close coil)
LVSx = Latch Valve Status
PSE-OPMx - Power System Electronics Output Module #x
PVx-y = Pyro Valve #x, NSI #y
P LV3B C
P LV6B C
SPARE
12A
12A
12A
12A
12A
•
B
FPGA
–
1553
DC/DC
–
PSE-OPM2
1A (2 pair)
–
Load Bus
•
1.2A
1.2A
12A
12A
12A
Four identical and independent
cards
–
PSE-OPM2
15A (4 pair)
1.2A
1.2A
1.2A
1.2A
1.2A
12A
12A
12A
12A
Signal
Input
LVS-2
LVS-6
SBW-10
Signal
Input
LVS-3
–
SPARE
D
FPGA
Redundant PDE actuators spread
across the cards
–
LVS-7
–
1553
DC/DC
PSE-OPM4
1A (2 pair)
PSE-OPM4
15A (4 pair)
Load Bus
–
•
Each has its own 28V power
inputs
Separate PSE output modules
power each PDE card
Individual cards can be powered
down as required
Each has its own interface to the
1553 bus to the C&DH
Can lose any one card and meet
mission requirements
Deployment actuators all have
redundant NSIs
Latch valves have redundant
coils
Thruster are partially redundant
Can accommodate worst case
load of all thrusters on
simultaneously
SARx-y = Solar Array Release #x, NSI #y
SBWx = Separation Break Wire #x
T20N-x = 20N Thruster #x
T80N-x = 80N Thruster #x
James Simpson - ACS
39
Inhibit Unit
PDE Card (1 of 4)
PSE Output
Module (1 of 4)
Actuators
Inhibit Unit
(1 of 4 Pairs)
Control Electronics
C&DH
Breakwire A
Breakwire B
James Simpson - ACS
40
PDE Development Status
• Requirements review
– Completed 1/11/06
• Bread board schematic review
– Completed 1/30/06
• Bread board in layout
James Simpson - ACS
41
ACS Analysis
• Successfully completed PDR Level Peer Review in
September 2005
– Completed Analysis
•
•
•
•
•
•
•
•
Sensor and actuator placement
Sensor and actuator specification
Spacecraft Error Budget (Knowledge, Pointing, and Jitter Stability)
HGA Pointing Error Budget
Fuel Slosh Analysis
Momentum management analysis
Mass Properties Analysis and Verification
Rigid Body Stability Analysis
– Gain and Phase margin requirement met with 20% variance in mass
properties
• Roll and Yaw Maneuver Analysis
– Software Tool development completed and ready for new mass
properties
James Simpson - ACS
42
ACS Analysis
• Post Peer Review analysis
– ACS Flight Software begun and ongoing
– ACS High Fidelity Simulation begun and ongoing
– Monte Carlo analysis has recently started
• Analysis results show ACS design meets or exceeds all Level
2 and 3 ACS requirements
• Latest LRO configuration is not significantly different from
LRO configuration presented at GN&C Peer Review
– Despite new solar array design and increased mass properties, no
major obstacles expected
• Delta Peer Review will be held on April 26, 2006
James Simpson - ACS
43
ACS Pointing Budgets
Requirement ( arcsec, 3-sigma.)
C.B.E ( Bias + Random )
( arcsec )
X axis
Y axis
Z axis
60
60
60
31.0
34.9
31.4
29.0
25.1
28.6
Margin ( arcsec, 3-sigma)
Bias Errors ( Sum )
Random Errors (RSS)
6.1
6.1
6.1
Comments
Knowledge budget
2 ST
2. Instrument Disturbances
2.7
2.7
2.7
Analysis
3. Wheel Disturbances
10.7
16.7
11.5
Analysis
1. Knowledge Errors
12.0
12.0
12.0
4. HGA/SA Disturbances
SDO/Analysis
9.9
9.9
9.9
5. ACS Controller Errors
Sub-Totals (arcsec)
Analysis
12.0
12.0
12.0
James Simpson - ACS
10.2
19.0
10.2
22.9
10.2
19.4
44
Observing Mode Knowledge Error Budget
Requirement ( arcsec.)
X axis
Y axis
Z axis
30.0
30.0
30.0
23.4
25.7
25.7
6.6
4.3
4.3
C.B.E ( Bias + Random )
Margin
Bias Errors ( Sum )
Random Errors (RSS)
1 Kalman Filter error relative to
prime Star Tracker
5.4
6.0
6.0
2. Kalman Filter unmodeled
error sources
5.0
5.0
5.0
9.4
11.8
11.8
3. Mechanical Alignment Prime
ST relative to Optical Bench
7.0
7.0
7.0
3.5
3.5
3.5
12.0
12.0
12.0
11.4
13.7
13.7
Comments
Modified
Farrenkopf - 2 ST
Analysis
Mechanical
subsystem
Sub-totals (arcsec)
James Simpson - ACS
45
Observing Mode 4 sec Stability Error Budget
X axis
Y axis
Z axis
Requirement ( arcsec, 3-sigma )
20.0
20.0
20.0
C.B.E. ( Bias + Random )
(arcsec)
13.8
17.3
13.8
6.2
2.7
6.2
Margin ( arcsec )
Bias Errors (Sum)
Random Errors (RSS)
1. Instrument Disturbances
2.7
2.7
2.7
Mechanical
subsystem
2. Wheel Disturbances
2.2
10.7
2.2
Analysis
3. HGA Disturbances
1.0
1.0
1.0
SDO/Analysis
4. Solar Array Disturbances
9.7
9.7
9.7
SDO/Analysis
5. Solar Radiation Pressure
1.0
1.1
1.0
Analysis
6. ACS Controller
9.1
9.1
9.1
Analysis
Sub-Total (arcsec)
13.8
17.3
13.8
James Simpson - ACS
46
Observing Mode 0.1 sec Stability Error Budget
X axis
Y axis
Z axis
Requirement ( arcsec, 3-sigma )
10.0
10.0
10.0
C.B.E. ( Bias + Random )
( arcsec)
Margin ( arcsec )
7.2
9.0
7.2
2.8
1.0
2.8
Bias Errors (Sum)
Random Errors (RSS)
1. Instrument Disturbances
2.4
2.4
2.4
Mechanical
2. Wheel Disturbances
0.7
5.4
0.7
Analysis
3. HGA Disturbances
1.0
1.0
1.0
SDO/Analysis
4. Solar Array Disturbances
4.9
4.9
4.9
SDO/Analysis
5. Solar Radiation Pressure
1.0
1.1
1.0
Analysis
6. ACS Controller
4.5
4.5
4.5
Analysis
Sub-Total (arcsec)
7.2
9.0
7.2
James Simpson - ACS
47
Moments of Inertia Summary
Moments of Inertia
(Deployed, version J)
599.4
-1.90
21.75
Moments of Inertia
(Deployed, version β)
1216.13
-81.21
-66.07
-1.90
662.14
-14.44
-81.21
1143.35
-60.04
21.75
-14.44
685.02
21.75
-60.04
1456.59
ACS Requirement on Moments of Inertia
Ixz_mean = 5 kg-m2, Ixz_amp = 8 kg-m2
Iyz_mean = 30 kg-m2, Iyz_amp = 15 kg-m2
James Simpson - ACS
48
Sun Safe ACS Performance
Tip off case of 2 deg/sec per axis initial body rates using version J Moments of Inertia
angle between actual and desired sun vectors
momentum (Nms)
200
angle (deg)
150
eclipse
100
50
0
0
500
1000
1500
2000
2500
3000
time (sec)
Stowed array is pointed to the Sun within 6 minutes
James Simpson - ACS
49
Sun Safe ACS Performance
Tip off case of 1,2,2 deg/sec per axis initial body rates (YPR) using
Version β Moments of Inertia
a ng le b e twe e n a ctua l a nd d e sire d sun ve cto rs
momentum (Nms)
200
angle (deg)
150
e clip se
100
50
0
0
1000
2000
3000
tim e (se c)
array reaches
l bStowed
t
t l SundSafe
d attitude
i d within 16 minutes
t
James Simpson - ACS
50
Observing Mode Pointing Performance after
180 Deg Yaw Maneuver, Version J
Th Z err. (arc-sec)
Th Y err. (arc-sec)
Th X err. (arc-sec)
Pointing Control is kept within 60 arcsec/axis requirement
50
0
-50
1400
1500
1600
1700
1800
1900
2000
1400
1500
1600
1700
1800
1900
2000
1400
1500
1600
1700
Time (sec)
1800
1900
2000
50
0
-50
50
0
-50
James Simpson - ACS
51
Observing Mode Knowledge Error after
180 Deg Yaw Maneuver, Version J
Est. Z err. (arc-sec) Est. Y err. (arc-sec) Est. X err. (arc-sec)
Attitude Knowledge is kept within 30 arcsec/axis requirement
20
0
-20
1400
1500
1600
1700
1800
1900
2000
1400
1500
1600
1700
1800
1900
2000
1400
1500
1600
1700
Time (sec)
1800
1900
2000
20
0
-20
20
0
-20
James Simpson - ACS
52
Stability Performance after 180 Deg Yaw, Version J
•Attitude Errors (3σ) are within 20 arcsec requirement
•Short Term Stability is under 10 arcsec requirement
•Long Term Stability under 20 arcsec requirement
Num Pnts
0
-20
0
20
X err. (asec)
Ang Err., 3 sig. =9.56
10000
5000
0
-10
5000
0
-10
0
Z err. (asec)
10
0
-20
0
20
X err. (asec)
4s Err., 3 sig. =3.18
10000
5000
0
-5
0
5
Y err. (asec)
4s Err., 3 sig. =3.28
10000
Num Pnts
Num Pnts
0
10
Y err. (asec)
Ang Err., 3 sig. =8.93
10000
5000
5000
0
-10
0
10
Z err. (asec)
James Simpson - ACS
Num Pnts
5000
Num Pnts
Num Pnts
Num Pnts
10000
Num Pnts
4s Err., 3 sig. =7.54
10000
Num Pnts
Ang Err., 3 sig. =20
0.1s Err, 3 sig. =0.195
10000
5000
0
-0.5
0
0.5
X err. (asec)
0.1s Err, 3 sig. =0.0859
10000
5000
0
-0.2
0
0.2
Y err. (asec)
0.1s Err, 3 sig. =0.0886
20000
10000
0
-0.2
0
0.2
Z err. (asec)
53
Observing Mode Pointing Error
Pointing Error Analysis Results well within 60 arcsec pointing requirement
Yaw Err. (arc-sec) Pitch Err. (arc-sec) Roll Err. (arc-sec)
Star Tracker models do not contain systematic errors
Version J
Version Beta
50
0
-50
500
1000
1500
2000
2500
3000
500
1000
1500
2000
2500
3000
500
1000
1500
2000
Time (sec )
2500
3000
50
0
-50
50
0
-50
James Simpson - ACS
54
Momentum Management during Nadir Pointing
(Observing Mode)
Inertia Tensor (Deployed at Beg of Life), Version J
⎡599.40 − 1.90 21.75 ⎤
⎢ − 1.90 662.14 − 14.44⎥
⎥
⎢
⎢⎣ 21.75 − 14.44 685.02 ⎥⎦
James Simpson - ACS
Kg-m^2
55
Momentum Management during Nadir Pointing
(Observing Mode)
HRW2 (N-m-s)
HRW1 (N-m-s)
Angular Momentum RWA, Ixz = 5, Ixz amp = 8, Iyz = 30, Iyz amp = 15 kg-m2
HRW3 (N-m-s)
•
Version β Inertia Tensor has a requirement on the applicable cross product of inertia terms:
Ixz_mean = 5 kg-m2, Ixz_amp = 8 kg-m2,
Iyz_mean = 30 kg-m2, Iyz_amp = 15 kg-m2
Results show the requirement for 14 days between unloading is met
given inertia tensor requirements are met
HRW4 (N-m-s)
•
100
0
-100
0
2
4
6
8
10
12
14
0
2
4
6
8
10
12
14
0
2
4
6
8
10
12
14
0
2
4
6
8
(days) 80 N-m-s
HwTime
heel
10
12
14
100
0
-100
100
0
-100
100
0
-100
Inertia Tensor (Deployed at Beg of Life), Version β
James Simpson - ACS
56
Maneuver Performance 20 Deg Roll, Pointing
Error
Roll maneuver analysis using Version J Moments of Inertia
Roll Maneuver completed in less than 20 minutes
Roll Error
(deg)
10
0
-10
-20
0
100
200
300
400
500
600
700
800
0
100
200
300
400
500
600
700
800
0
100
200
300
400
Time (sec)
500
600
700
800
Pitch Error
(deg)
0.04
0.02
0
-0.02
Yaw Error
(deg)
0.2
0
-0.2
James Simpson - ACS
57
Maneuver Performance 180 Deg Yaw, Pointing Error,
Version J
Z pnt. err. (deg)
Y pnt. err. (deg)
X pnt. err. (deg)
Yaw maneuver complete in 20 minutes
50
0
-50
0
200
400
600
800
1000
1200
1400
1600
1800
2000
0
200
400
600
800
1000
1200
1400
1600
1800
2000
0
200
400
600
800
1000 1200
Time (sec)
1400
1600
1800
2000
20
0
-20
-40
100
0
-100
-200
180 Deg Yaw maneuver analysis using Version J Moments of Inertia
James Simpson - ACS
58
Delta-V/Delta-H Control Performance
•
Phase Plane Plot
0.5
X-axis
Y-axis
Z-axis
Requirement
0.4
0.3
Rate (deg/sec)
0.2
0.1
0
-0.1
-0.2
-0.3
-0.4
-0.5
-6
-4
-2
0
2
Attitude Error (deg)
4
James Simpson - ACS
6
Initial Conditions and
Assumptions
– Deployed configuration with
Full fuel
– Initial rates: (0,0,0) deg/sec
– Initial System Momentum:
(31, -132, 21) N-m-s
– Commanded wheel torque:
0.15 N-m/wheel
– Simulation Time of 20 min
– IRU Model
• Read-out Noise = 1 (1σ)
arcsec
• Angle Random Walk =
1.8 arcsec/(sec)0.5
• Rate Random Walk =
0.001 arcsec/(sec)1.5
59
Rigid Body Margins
Mode
Bandwidth (Hz)
Gain Margin (db)
Phase Margin (deg.)
X
Y
Z
X
Y
Z
0.02
0.019
0.021
16.5
16.1
16.8
14.5
12.9
16.4
16.5
16.1
16.8
68.7
72.0
64.2
71.9
68.3
75.6
68.7
72.0
64.2
Safe-Sun
(w/o gyros)
- Nominal
- +20%
- -20%
0.013
0.013
0.013
12.4
12.6
12.4
15.4
14.5
15.9
14.5
15.1
14.1
54.2
53.9
53.6
78.0
76.6
79.1
55.2
54.9
54.5
Delta-V/Delta-H
- Nominal
- +20%
- -20%
0.116
0.137
0.101
21.7
19.7
23.2
21.7
19.7
23.2
21.7
19.7
23.2
34.8
37.4
32.5
34.8
37.4
32.5
34.8
37.4
32.5
Observing
- Nominal
- +20%
- -20%
0.04
0.033
0.049
14.6
13.8
17.4
14.6
13.8
17.4
14.6
13.8
17.4
45.3
42.2
49.8
45.3
42.2
49.8
45.3
42.2
49.8
Safe-Sun
( with gyros)
- Nominal
- +20%
- -20%
First structural mode required to be no less than 1 Hz
Controller bandwidth requirement to be at least 1 decade below first structural mode
Rigid body stability analyses are single axis and insensitive to inertia changes
Rigid body stability margins are 6 db ( goal of 12 db) in gain and 30 degrees in phase margins
Variation of +/- 20% is study of mismatch between ACS estimate and true inertias
James Simpson - ACS
60
Onboard Ephemerides
•
•
•
•
•
Within GNC FSW, onboard Ephemerides models are in separate
process from and supplied to ACS every cycle
Spacecraft Ephemeris
– Interpolated
– Output: Time, ECI, J2000, position and velocity vectors
Solar Ephemeris
– Analytical
– Output: Time, ECI, J2000 position and velocity vectors
( velocity used for ST aberration correction )
Lunar Ephemeris
– Interpolated
– Output: Time, ECI, J2000 position and velocity vector
HGA Ka Band Ground Station Ephemeris (White Sands)
– Analytical
– Output: Time, ECI, J2000 position vectors
James Simpson - ACS
61
HGA Pointing Error Budget
Requirement (3σ)
0.130 °
C.B.E. (3σ)
0.111°
Margin (3σ)
0.019°
Bias Errors (sum)
Random Errors (RSS)
ACS Subtotal
0.000
0.040
Alignment Subtotal
0.039
0.050
Control Subtotal
0.008
0.008
Totals
0.047
0.065
James Simpson - ACS
62
High Gain Antenna (HGA)
•
Body frame and Gimbal Frame are co-aligned
when gimbals are at zero degrees
±90 Degrees
of Rotation
(Pitch)
[A]GF
BCS = [I ]3 x3
•
Gimbal Frame to Antenna frame is given by
Ant
[A(α , β )]GF
= [A(β )]Roll [A(α )]Pitch
•
Calibration of HGA identical to SDO scenario
– Pointing accuracy is 0.3 deg of ground station
– Scan pattern around ground station location
– Ground station measure signal strength
– Compute pointing biases from signal strength
– Multiple orbits identifies scale factor error
James Simpson - ACS
±90 Degrees
of Rotation
(Roll)
+Z S/C Frame
63
Solar Array (SA)
•
Body frame and Gimbal Frame are co-aligned
when gimbals at zero degrees
[A]GF
BCS = [I ]3 x3
•
Gimbal to Solar Array frame
[A(el, Az )]SA
GF = [A( Az ) ]Roll [A(el ) ]Pitch
+Z
+X +Y
S/C Body Frame
Solar Array Roll Gimbal
Solar Array Pitch Gimbal
James Simpson - ACS
64
Testing
• All procured items (STs, MIMU, CSSs) will be acceptance
tested at the vendor facilities.
• In-House components (RWs, PDE) will be tested at every
level
– Board level
– ETU box level
– Flight box level
FlatSat testing
I&T facility
• FlatSat testing will include ETU testing with flight software
for verification credit.
• I&T testing will include copper path testing & polarity testing
– Additional goal of “flying LRO for a week” with real-time, closed
loop testing
– Star Tracker and MIMU stimulator modifications are being
researched in support of this real-time, closed loop testing
James Simpson - ACS
65
Goddard Dynamic Simulator
• 6 DOF simulator for
– Environment dynamics
– Math models of instruments,
ACS HW, C&DH, HGA,
Solar Array
– Propulsion Loads
• Uses ground system to
communicate with the
spacecraft
James Simpson - ACS
66
ACS Risks
Rank &
Trend
5
L
I 4
K
E
L 3
I
H
O 2
O
D
Approach
Risk Title
Risk ID
1
↓
M
Parts Procurement
Manpower Support
ACS-HW33
2
→
R
CSS Cold Survival
Temperature Test.
ACS-HW109
1
2
1
2
1
3
4
5
CONSEQUENCES
Criticality
L&C Trend
Approach
High
↓
Decreasing (Improving)
M – Mitigate
Med
↑
Increasing (Worsening)
W – Watch
Low
→
Unchanged
A – Accept
New since last month
R – Research
*
1. Parts Procurement engineers are
overwhelmed, at times, with the demands
from multiple projects and multiple
subsystems. More personnel are
recommended to avoid delays in schedule.
2. Performance testing of the CSS over the full
LRO operating temperature range is being
researched. One option could be to add
heaters on array mounted CSSs, design test
facility to support full operating
temperature testing range or demonstrate
survival of CSS in cold case.
James Simpson - ACS
67
Preliminary ACS Reliability Study
Single String (Mission Life = 14 months = 10220 hrs)
Failure Rates and Percentages
Components
Coarse Sun Sensors (6 out of 10)
Reaction Wheels (3 out of 4)
Star Trackers (1 out of 2)
Gyro
Total Failure Rate =
ACS Subsystem Reliability
Component
Block
Reliability Redundancy
0.999975
0.976912
0.975581
0.977767
1.000000
0.996899
0.999404
Component
Failure Rates
Percent of Total
2.4100E-09
2.2856E-06
2.4190E-06
2.2000E-06
0.000%
5.598%
1.075%
93.33%
2.3573E-06
100.00%
0.97415
James Simpson - ACS
68
Preliminary PDE Reliability Study
Failure Rates and Percentages
Assembly/Board
Connectors (38)
Relay (26)
Sep - Switch (8)
FBWs (4)
PDE Modules (3-out-of-4)
Total Failure Rate =
PDE Subsystem Reliability
Card
Block
Reliability Redundancy
0.995737
0.993379
0.995920
0.997958
0.984152
0.998524331
Failure Rates
Percent of Total
1.1000E-08
2.5000E-08
5.0000E-08
5.0000E-08
1.5631E-06
17.55%
13.16%
24.02%
4.09%
0.35%
4.9825E-06
100.00%
0.98165
James Simpson - ACS
69
Documents
•
Level 3 Requirements
–
•
Procurement Documents
–
–
–
•
Star Tracker (All documents under final CM review)
• 431-SPEC-000217
• 431-LIST-000428
• 431-SOW-000216
IRU (All documents in Procurement Office)
• 431-SPEC-000219
• 431-LIST-000419
• 431-SOW-000218
CSS(All documents in Procurement Office)
• 431-SPEC-000214
• 431-LIST-000279
• 431-SOW-000215
PDE
–
•
431-SPEC-000162 LRO ACS Specification
• Awaiting closing of actions
431-RQMT-000421
• Under CM review
RW
–
431-ICD-000148
• Under CM Review
James Simpson - ACS
70
ACS Schedule
Item
UID Responsible
Date
Status
ACS Analysis
1029
GSFC/591
8/30/2006
Ongoing
CSS
1964
Unknown
1/24/2007
RFP Released
Star Tracker
1968
Unknown
9/7/2007
RFI
Complete/RFP
Release pending
IRU
1972
Honeywell
6/13/2007
Sole Source
pending/ RFP
pending
RW
1984
GSFC/596
8/8/2007
ETU Testing
PDE
1980
GSFC/596
8/17/2007
Breadboard
Development
James Simpson - ACS
71
GN&C Peer Review
Request for Actions (RFA) Status
• LRO GN&C Peer Review
produced 29 RFAs
– 3 closed
– 26 open
James Simpson - ACS
72
Conclusion
• ACS is very well defined and meets requirements
– ACS has completed a PDR Level Peer Review
– Level 3 Requirements are baselined
• Level 4 Requirements are under review
– ACS Performance Analysis illustrates a robust design
• Latest mass properties do not pose any considerable threat to
design
– Hardware choices are very mature
• Procurements are well under way
• Any new designs (PDE) are very mature
– Schedule is defined and achievable
– Few risks exist and are being worked
James Simpson - ACS
73
Backup Slides
PDE Actuator Mapping
LVS-4
LVS-5
SBW-9
A
Signal
Input
FPGA
1553
PSE-OPM1
1A (2 pair)
DC/DC
PSE-OPM1
15A (4 pair)
Load Bus
LVS-1
SPARE
SPARE
Signal
Input
C
FPGA
1553
PSE-OPM3
1A (2 pair)
DC/DC
PSE-OPM3
15A (4 pair)
Load Bus
1.2A
1.2A
1.2A
1.2A
T20N-1
T20N-2
T20N-3
T20N-4
P LV2A C
1.2A
1.2A
1.2A
1.2A
12A
P LV4A C
P LV6A C
P LV5A C
P LV7A C
P LV1A C
T80N-1
T80N-2
PV1-2
SAR1-2
SAR3-2
HGAR1-2
P LV3A C
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
12A
12A
12A
12A
12A
12A
PV1-1
SAR1-1
SAR3-1
HGAR1-1
12A
SPARE
SPARE
12A
1.2A
T20N-5
P LV2B C
1.2A
1.2A
1.2A
1.2A
1.2A
1.2A
T20N-6
T20N-7
T20N-8
P LV4B C
1.2A
1.2A
12A
12A
12A
P LV5B C
P LV7B C
P LV1B C
PV2-1
T80N-3
T80N-4
PV2-2
12A
12A
12A
12A
SAR2-1
SAR4-1
HGAR2-1
SAR2-2
SAR4-2
HGAR2-2
SPARE
Reference Designators
HGARx-y - High Gain Antenna Release #x, NSI #y
LVx = Latch Valve #x (P = open coil, C = close coil)
LVSx = Latch Valve Status
PSE-OPMx - Power System Electronics Output Module #x
PVx-y = Pyro Valve #x, NSI #y
P LV3B C
P LV6B C
SPARE
12A
12A
12A
12A
12A
LVS-2
LVS-6
SBW-10
•
B
FPGA
1553
DC/DC
PSE-OPM2
1A (2 pair)
• 2 NSI each
– Seven dual coil latch
valves with status
indicator
Load Bus
1.2A
1.2A
12A
12A
12A
LVS-7
Signal
Input
D
PSE-OPM4
1A (2 pair)
•
PSE-OPM4
15A (4 pair)
James Simpson - ACS
High gain antenna
deployment
– Two release points
– 2 NSI per release point
Load Bus
SARx-y = Solar Array Release #x, NSI #y
SBWx = Separation Break Wire #x
T20N-x = 20N Thruster #x
T80N-x = 80N Thruster #x
Solar array deployment
– Four release points
– 2 NSI per release point
SPARE
1553
DC/DC
•
LVS-3
FPGA
Propulsion
– Eight 20N thrusters
– Four 80N thrusters
– Two pyrotechnic valves
PSE-OPM2
15A (4 pair)
1.2A
1.2A
1.2A
1.2A
1.2A
12A
12A
12A
12A
Signal
Input
•
Two separation break wire
monitors
75
PDE Module Card Layout
• Module 6U plus card component placement
James Simpson - ACS
76
Acronyms
AIDR
Acceleration Insensitive Drift Rate
DILS
Deliverables Item List and Schedule
FDC
Fault Detection and Correction
IRU
Inertial Reference Unit
MIMU
Miniature Inertial Measuring Unit
NEA
Noise Equivalent Angle
PDE
Propulsion Deployment Electronics
RW
Reaction Wheel
SOW
Statement of Work
James Simpson - ACS
77
LRO PDR
Electrical Systems
Philip Luers
February 8, 2006
Electrical Systems Agenda
•
•
•
•
•
•
•
•
Introduction
Document Flow
Interfaces
Timekeeping
Power
Grounding
EMI
Radiation
Philip Luers - Electrical Systems
2
Introduction
• Electrical Systems
– Includes electronics components, interconnect harnessing,
structural chassis grounding system, grounding of external
surfaces, and elements that provide shielding
• Drivers
–
–
–
–
–
Heritage subsystems
Thermal control system
High Data Rates
Common mode noise control
Charging environment
Philip Luers - Electrical Systems
3
Electrical Document Tree
Mission Requirements Document
(431-RQMT-000004)
General Thermal
Subsystem
Specification
(431-SPEC0000091)
Heaters,
Thermistors
EELV PPGs
LV Requirements
Payload
Electrical Systems
Specification
(431-SPEC-000008): Power,
1553, SpaceWire, 1 pps, RS422, Grounding, Isolation,
Shielding, Charging
C&DH
Electrical Systems
Requirements
Document
(431-RQMT-000140)
Power
Comm
Specific Interfaces
Connectors, pinouts
GN&C
Released
Draft
Planned
Philip Luers - Electrical Systems
4
Payload Electrical Doc. Flow
CRaTER EICD (431-ICD-000094)
DLRE EICD (431-ICD-000095)
Payload
LAMP EICD (431-ICD-000096)
LEND EICD (431-ICD-000097)
LOLA EICD (431-ICD-000098)
LROC EICD (431-ICD-000099)
Mini-RF EICD (431-ICD-000152)
Released
Draft
Electrical Systems
Requirements
Document
(431-RQMT-000140)
Planned
Philip Luers - Electrical Systems
5
C&DH, Power,Comm Elec. Doc. Flow
C&DH EICD (431-ICD-000141)
LRO SpaceWire Spec. (431-SPEC-000103)
C&DH
C&DH Req (431-RQMT-000168)
MSS Spec (431-SPEC-000381)
Timing Spec (431-SPEC-000212)
Timing Validation Plan (431-PLAN-000TBD)
EPS Spec (431-SPEC-000013)
PSE Req (431-RQMT-000016)
PSE EICD (431-ICD-000142)
Solar Array Spec (431-SPEC-000037)
Solar Array EICD (431-ICD-000150)
Battery Spec (431-SPEC-000032)
Battery EICD (431-ICD-000151)
Power
Specific Interfaces
Connectors, pinouts
S-band Spec (431-SPEC-000121)
Comm
Released
Ka-band Mod Spec (431-SPEC-000138)
Communications EICD (431-ICD-000146)
Ka-band TWTA Spec (431-SPEC-000178)
Draft
Electrical Systems
Requirements
Document
(431-RQMT-000140)
Planned
Philip Luers - Electrical Systems
6
GN&C Elec. Document Flow
PDE Rqmt (431-RQMT-000421)
PDE EICD (431-ICD-000143)
Star Tracker Spec (431-SPEC-000217)
Star Tracker EICD (431-ICD-000144)
IMU Spec (431-SPEC-000219)
IMU EICD (431-ICD-000145)
Propulsion Spec (431-SOW-000017)
Propulsion EICD (431-ICD-000147)
Reaction Wheels EICD (431-ICD-000148)
GN&C
Gimbal Controller EICD (431-ICD-000149)
CSS Spec. (431-SPEC-000214)
Specific Interfaces
Connectors, pinouts
Released
Draft
Electrical Systems
Requirements
Document
(431-RQMT-000140)
Planned
Philip Luers - Electrical Systems
7
LRO Simplified System Block Dia.
Mini-RF
Backup
20MHz Clock
H/W Decoded
Command
Discretes
LROC
SpaceWire
Network
LAMP
USO 9600
S-Xpndr
Comm
Hi-Rate
Tlm
DIB
LAMP Sci. & HK
Omnis
Low-Rate
Cmds & Tlm
HGA
Ka-Xmtr
ATA
SBC
20MHz Clock
Discretes
Thermistors
Closed Loop Htrs
LEND
HK / IO
DDA
Power Bus
LOLA
ST(2)
HGA
Gimbals
MAC
Unsw. + 28V
Diviner
+ 28V
USO 9500
GIMBAL
CONROL
IRW(4)
CSS(10)
LVPC
C&DH
IMU
LRO Simplified
Systems Block Dia.
1-24-06
MIL-STD-1553 Network
CRaTER
PDE
PMC
Battery
Prop/Dep-A
SAM
P
Solar
Array
Sw. and
Unsw.
+28V Pwr
Services
Prop/Dep-C
OM-B
Prop/Dep-D
OM-C
SA
Gimbals
OM-D
GIMBAL
CONTROL
Propulsion
Prop/Dep-B
OM-A
Inhibit Unit
PSE
P
R
R
P
SA & HG
Deploy
Actuation
Vehicle
Separation
Break Wires
Philip Luers - Electrical Systems
8
Payload Interfaces
CRaTER EICD (431-ICD-000094)
LEND EICD (431-ICD-000097)
DLRE EICD (431-ICD-000095)
LOLA EICD (431-ICD-000098)
LROC NAC-L (Sw +28V)
LROC NAC-R (Sw +28V)
LROC WAC (Sw +28V)
LROC Decon (Sw +28V)
LROC Decon Heater (Sw +28V)
Instrument Survival Heater Bus
(Unsw +28V)
NAC-L
NAC-R
WAC
LROC
40Mbps Tx/4Mbps Rx
SpaceWire to/from C&DH SBC
SCS
Thermistors (8)
LAMP EICD (431-ICD-000096)
LROC EICD (431-ICD-000099)
Mini-RF EICD (431-ICD-000152)
Philip Luers - Electrical Systems
9
C&DH Interfaces
SpaceWire Port0
SBC
NC (SpaceWire Port1)
NC (SpaceWire Port2)
NC (SpaceWire Port3)
SBC, PSE,
C&DH LVPC
SBC
Discrete Commands
Mini-RF
J5
J10
S-Comm J2
J3
J1
S-band Tlm
Umbilical
S-band Tlm
EELV (Umbilical)
J6
NC (JTAG Test)
J7
NC
J11
J17
Ka-band Tlm
J14
J18
Ka-band Ctl/Status
NC (SpaceWire Port2)
J12
SpaceWire Port0
NC (SpaceWire Port2)
NC (SpaceWire Port3)
Ka-Comm
J16
Ka-band Modulator
Ka-band Tlm
J13
J26
J28
LAMP Hi-speed & Low-speed
J31
J29
J27
LAMP
CRaTER, LEND, DLRE,
LOLA, PSE, S/C Test Panel
1 pps
20 MHz
HK/IO
US0
9500
J34
J32
J33
NC (JTAG Test)
NC
20 MHz
Analog-2
Analog-3
Analog-4
SpaceWire Port0
S-Comm
Ka-Comm
LROC
HK/IO
SpaceWire Port1
SpaceWire Port2
SpaceWire Port3
J38
J40
J39
J41
MAC
ATA Ultra-5
ATA Ultra-5
C&DH Power (Unsw +28V)
Comm. Power (Unsw +28V)
S-Comm
Discrete Commands
1553 Bus B
J37
J42
J43
J48
NC (JTAG Test)
J44
J49
1553 Bus A
SBC
J50
US0
9600
NC (JTAG Test)
J36
C&DH Ctl Htr (Sw. +28V)
C&DH Ctl Htr
1553 Bus B
J46
J47
DDA
1553 Bus A
J35
J45
J53
DIB
J51
Ka-band TWTA
S/C Test Panel
J15
NC (JTAG Test)
J30
Analog-1
H/W Decoded Commands from S-comm to:
0.
S-comm Self Reset
•
1C&DH LVPC POR (Reset)
•
C&DH LVPC Power Cycle
•
PSE “All OFF”
•
PSE “SAM Normal Mode”
S-band Xpndr Ctl/Status
NC (SpaceWire Port1)
SpaceWire Port0
S-band Transponder
J8
J9
SpaceWire Port0
NC (SpaceWire Port3)
SBC
S-band Tlm/Commands
J4
Discrete Commands
S-Comm,
PSE
H/W Decoded Commands from SBC to:
0.
S-comm Reset
•
PSE PMC Reset
•
PSE “Jumper ON”
•
PSE “SAM All ON”
NC (Relay Comm)
J52
J25
J21
J24
USO 9600 (Sw +15V)
LVPC
J20
J23
J22
C&DH
J19
RF Switch Commands
NC (C&DH LVPC Test)
RF Transfer Switch
C&DH EICD (431-ICD-000141)
NC (Comm LVPC Test)
MSS Spec (431-SPEC-000381)
USO 9500 (Sw +28V)
Philip Luers - Electrical Systems
10
PSE Interfaces
PSE Discrete Commands (from):
1.
PSE PMC Reset (SBC)
2.
PSE “Jumper ON” (SBC)
3.
PSE “All OFF” (S-comm)
4.
PSE “SAM All ON” (SBC)
5.
PSE “SAM Normal Mode”
(S-Comm)
PSE EICD (431-ICD-000142)
Philip Luers - Electrical Systems
11
GN&C Interfaces
Ka TX Sw +28V
S-band TX (Sw + 28V)
Mini-RF PA (Sw +28V)
4 SAR – red (8)
CSS Spec. (431-SPEC-000214)
C&DH EICD (431-ICD-000141)
Inhibit
Unit
2 HGAR – red (4)
Ka TX Sw +28V
ELV Sep
Relay
S-band TX Sw + 28V
IRW1-4 (Sw +28V)
5 lb thruster control (8)
ELV Sep
Relay
IRW (1-4)
Mini-RF PA Sw +28V
MIL-STD-1553 Network
PDE-A Sw +28V
Reaction Wheels EICD (431-ICD-000148)
20 lb thruster control (4)
7 LV (open/close) – red (14)
ELV Sep
Relay
2 PV – red (4)
Prop/Dep-A
Prop/ Deploy Sw +28V
MIL-STD-1553 Network
PDE-B Sw +28V
Prop/Dep-B
Prop/ Deploy Sw +28V
MIL-STD-1553 Network
ELV Sep
Relay
PDE-C Sw +28V
IMU Spec (431-SPEC-000219)
Prop/Dep-C
Prop/ Deploy Sw +28V
IMU EICD (431-ICD-000145)
PDE-D Sw +28V
MIL-STD-1553 Network
ELV Sep
Relay
Prop/Dep-D
Prop/ Deploy Sw +28V
MIL-STD-1553 Network
ELV Sep
Relay
Star Tracker Spec (431-SPEC-000217)
PDE
PDE Rqmt (431-RQMT-000421)
Propulsion Spec (431-SOW-000017)
PDE EICD (431-ICD-000143)
Propulsion EICD (431-ICD-000147)
Star Tracker EICD (431-ICD-000144)
Philip Luers - Electrical Systems
12
Comm. and Deployables Interfaces
14 Segments
MIL-STD-1553 Network
SA-GC
(Sw +28V)
25 Modules per Panel
2 Strings per Module
24 Cells per String
Solar Array
Gimbal
Controller
S/A
Deploy
Az SA
Gimbal
El SA
Gimbal
S/A
Deploy
27 Modules per Panel
2 Strings per Module
24 Cells per String
S/A
Deploy
S/A
Deploy
25 Modules per Panel
2 Strings per Module
24 Cells per String
S-band Spec (431-SPEC-000121)
Ka-band Mod Spec (431-SPEC-000138)
Ka-band TWTA Spec (431-SPEC-000178)
Solar Array Spec (431-SPEC-000037)
Communications EICD (431-ICD-000146)
Solar Array EICD (431-ICD-000150)
Gimbal Controller EICD (431-ICD-000149)
Philip Luers - Electrical Systems
13
Data and Signal Interfaces
•
MIL-STD-1553 bus – 19 RTs
– Dual redundant, transformer coupled, with 1:1.41 (stub to bus) ratio,
Trompeter PL3155AC on SC harness
•
SpaceWire
– ESA ECSS-E-50-12 standard, with 26 AWG harness, 9-pin connectors
•
LVDS
– Ka-Comm to Ka-band modulator
– LAMP High-speed telemetry
– 20 MHz USO signals
•
RS-422 UART
– Command and telemetry to/from S-band XPNDR and Ka-band Modulator
– LAMP Low Speed Control/Housekeeping
•
RS-422 Synchronous Serial (clock and data)
– LAMP High Speed telemetry (1 MHz)
– S-band telemetry to S-band XPNDR
– S-band telemetry to EELV
Philip Luers - Electrical Systems
14
Data and Signal Interfaces
•
RS-422 Discretes
–
–
–
•
•
+28V Discretes
– Deployment NSIs, Prop Pyro-valve NSIs, Prop latch valves, Thruster valves, Battery
Relay
ATA Ultra-DMA 5
–
•
To/from Disk Drive Assembly
Discrete telemetry
–
–
•
1 Pulse-per-second (pps) to 5 instruments and PSE
HW Decoded pulse commands to C&DH LVPC and PSE
SBC Discrete commands to S-comm and PSE
Latch valve status, separation break wires (gathered by PDE)
Battery relay status, bypass status (gathered by PSE)
Analog telemetry
–
–
–
–
–
S-311-P18-05A Thermistor, high-strength 26 AWG (gathered by MAC)
118MF PRT, high-strength 26 AWG (gathered by MAC)
Coarse Sun Sensors (gathered by MAC)
Pressure transducers (gathered by MAC)
Battery cell voltages (gathered by PSE)
Philip Luers - Electrical Systems
15
Data and Signal Interfaces
Philip Luers - Electrical Systems
16
Data Flow Block Diagram
Philip Luers - Electrical Systems
17
Time Maintenance
•
Requirements
– The Orbiter and ground system shall provide knowledge of the Orbiter time with
respect to UTC to an accuracy of 3 ms.
– Provide a mechanism for maintaining Orbiter time within 100 ms of UTC at all times
•
Mission Elapsed Time (MET)
– MET hardware counter on C&DH HK/IO card, continuously counting, LSB = 15
microseconds
• Resettable only via C&DH power cycle hardware decoded command
•
S/C Time Correlation
– S-comm card detects receipt of command Barker Code preamble, generates TickIn1 to
S-comm SpaceWire core
• Generates corresponding TickOut1 on HK/IO card
– HK/IO card latches MET of every spacecraft command received
– FSW periodically gathers MET of last ground command from HK/IO card
– FSW periodically telemeters “S/C time of last command” to ground
• S/C Time = MET + Spacecraft Time Correction Factor (STCF)
•
Operationally
– Ground station records UTC of command transmission for every command
– MOC must wait for S/C time of last command to be telemetered before sending
another ground command
– Compensate for known delays and compare S/C time to UTC to determine error
Philip Luers - Electrical Systems
18
Time Distribution
•
1 Pulse-per-second
–
1 Pulse-per-second (1 pps) distributed by C&DH HK/IO via RS-422 to 5 instruments, PSE and
Spacecraft Test Panel.
•
–
1 pps drives SpaceWire TickIn0 on HK/IO card
•
•
Rising edge on MET 1 second rollover, subseconds field = 0.
Resultant TickOut0 received by C&DH SBC, LROC and Mini-RF
“Time at tone will be” message
–
–
FSW gathers current MET from HK/IO, either at boot-up or every second
FSW adds 1, sends a “time at tone will be” message to users
•
•
•
•
CRaTER, DLRE, LEND (S/C Time) – via 1553,
LOLA (MET) – via 1553
LAMP (S/C Time) – via low-speed command/housekeeping interface
LROC & Mini-RF (S/C time) – via SpaceWire
1pps
(internal
to C&DH)
tPLH
tPLH
Parameter
Value
Meaning
tPLH, tPHL
2-30 nanoseconds
Delay through driver
tTLH, tTHL
1-13 nanoseconds
Rise time, Fall time
tH
35 +/1 microseconds
Pulse width
t1s
1 s +/- 10 nanoseconds
Pulse to pulse timing
t1s
tH
1pps90%
10%
1pps+
tLH
tHL
Philip Luers - Electrical Systems
19
Power Requirements
•
All non-critical unswitched loads shall be fused
– All critical unswitched loads shall be unfused
•
•
•
•
•
•
All switched loads protected by SSPC, derated to 80% of maximum
sustainable current
No switched returns
Different PSE OM provide redundant power where applicable
Ground test will verify all operational configurations and all copper paths
in distribution system
No damages to Orbiter due to misconfiguring power system
Voltage drop will be verified
– Anticipate < 1 volt drop between PSE (22-35 V) and load (21-35 V)
– < 5 watt power dissipation estimated in S/C harness
• not including SA and battery harness, which are accounted for in energy balance
Philip Luers - Electrical Systems
20
Power Distribution Capability
Subsystem
Switched
CRaTER
1 (1A)
DLRE
2 (2 x 2A)
LAMP
2 (2 x 2A)
LEND
1 (1A)
LOLA
1 (2A)
LROC
5 (5A, 4 x 1A)
Mini-RF
2 (5A, 2A)
S-XPNDR
1 (2A)
Ka-band TX
1 (10A)
C&DH
1 (2A)
HGA-GC
1 (2A)
SA-GC
1 (2A)
PDE
4 (4 x 2A)
Unswitched
Payload
Deployables
Communications
GN&C
C&DH
Thermal
1 (2A)
2 (5A, 2A)
Service Type
Per Output
Module (OM)
Total PSE
4 (4 x 15A)
Unswitched 2A
1
4
Prop Catbed
2 (2 x 2A)
Unswitched 5A
2
8
IMU
1 (2A)
Switched 1A
2
8
ST(2)
2 (2 x 1A)
Switched 2A
6
24
IRW(4)
4 (4 x 5A)
Switched 5A
2
8
Thermal
10 (3x10A, 2x5A, 5x2A)
Switched 10A
1
4
Spare
2 (2 x 2A)
Switched 15A
1
4
TOTAL
48
15 (12/3)
60 (48/12)
Prop
Deploy
9 (7x5A, 2x2A)
12
Total
Philip Luers - Electrical Systems
21
Power Distribution Assignments
#
Rating
OM-1
OM-2
OM-3
OM-4
1
UNSW 2A
S-band RX
C&DH S-Comm
Thermal:Inst Surv Htrs-2
Thermal:SC Ess Htrs-7
2
UNSW 5A
C&DH
Thermal:Inst Surv Htrs-1
Thermal:SC Ess Htrs-1
Thermal:SC Ess Htrs-2
3
UNSW 5A
Thermal:SC Ess Htrs-3
Thermal:SC Ess Htrs-4
Thermal:SC Ess Htrs-5
Thermal:SC Ess Htrs-6
4
SW 5A
IRW-1
IRW-2
IRW-3
IRW-4
5
SW 10A
Thermal:Prop Htrs-RED
Thermal:Prop Htrs-PRI
Thermal:Inst Op Htrs-1
Ka-band TX
6
SW 15A
Prop
Prop
Prop
Prop
7
SW 2A
Spare
HGA-GC
SA-GC
Thermal: Inst Surv-3
8
SW 1A
LROC-NAC-L
CRaTER
LEND
LROC-WAC
9
Removed
10
SW 5A
LROC-DECON
MiniRF-PA
Thermal: C&DH Ctl
Thermal: C&DH Ctl
11
SW 1A
ST-2
ST-1
LROC-NAC-R
LROC-SCS
12
SW 2A
S-band TX
PDE-B
DLRE Inst
Thermal:SC Op Htrs-2
13
SW 2A
IMU
LAMP-RED
DLRE DREB
Spare
14
SW 2A
MiniRF-Main
LOLA
LAMP-PRI
Prop Catbed-PRI
15
SW 2A
Thermal:SC Op Htrs-1
Prop Catbed-RED
PDE-C
PDE-D
16
SW 2A
PDE-A
Thermal: Deploy Htrs
Thermal:Gimbal Htrs PRI
USO (9500)
Deploy
Deploy
Payload
Deployables
Communications
GN&C
C&DH
Thermal
Philip Luers - Electrical Systems
Deploy
Deploy
22
+28V Backplane
Battery ON/OFF Line
SA-GC Sw +28V
HGA-GC Sw +28V
Power Bus
Power Distribution Diagram
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23
Power Distribution
•
•
Simple, centralized power switching for
most subsystems
CD&H Secondary Switching:
S-Band Comm
Ka Comm
100GB
DIB
DDA
100GB
SBC
– USO 9600 (+15V), backup oscillator
HK/IO
• Switched in C&DH LVPC, controlled by FSW
MAC
C&DH Controlled
Heaters
– Disk Drive Assembly
– Thermal: C&DH Controlled Heaters
S-comm
LVPC
• Switched on MAC, controlled by FSW
C&DH
– C&DH: SBC, HK/IO, MAC, DIB, DDA
ELV
Separation
Break Wires
• Can be power-cycled by momentary switch
actuated by hardware decoded command sent
from ground
•
USO
9600
20 MHz
Clock
C&DH
• Switched on DIB, controlled by FSW
S-band TX Pwr
S-band TX Pwr
Ka-band TX Pwr
Ka-band TX Pwr
Mini-RF PA Pwr
Mini-RF PA Pwr
PDE Mod-A Loads
PDE Secondary Switching
PDE Mod-B Loads
PDE Mod-C Loads
– S-band TX, Ka-band TX, Mini-RF PA
PDE Mod-D Loads
Inhibit Unit
• RF Hazards through PDE Inhibit Unit (IU):
PDE Mod-D
– Deployment and Propulsion Actuators
PDE Mod-D Loads
PDE Mod-D
• Deploy and Prop hazards through PDE IU and
PDE Modules
PDE Mod-C
PDE Mod-C Loads
PDE Mod-C
PDE Mod-B
PDE Mod-B Loads
PDE Mod-B
PDE Mod-A
OMs
PSE
Philip Luers - Electrical Systems
PDE Mod-A Loads
PDE Mod-A
PDE
24
Electrical Controls of Inhibit Devices
Control for inhibits of RF hazards (S-band TX, Ka-band TX, and Mini-RF PA),
propulsion hazards (valves), and deployment hazards (NSIs)
PDE inhibit unit, through separation break wires provides 2 independent controls of power to control
inhibits of hazards, C&DH provides third level of independent control.
ELV
Separation
Break Wires
S-band TX Pwr
S-band TX Pwr
Ka-band TX Pwr
Ka-band TX Pwr
Mini-RF PA Pwr
For any one hazard, two relays
must close to energize – each
relay controlled by different
break wire
Mini-RF PA Pwr
PDE Mod-A Loads
PDE Mod-B Loads
PDE Mod-C Loads
PDE Mod-D Loads
Inhibit Unit
In this example, two break wires
must open, indicating separation,
two close two latching relays to
energize any load controlled by
PDE module D
1
PDE Mod-D
2
PDE Mod-D Loads
PDE Mod-D
PDE Mod-C
PDE Mod-C Loads
PDE Mod-C
PDE Mod-B
PDE Mod-B Loads
PDE Mod-B
PDE Mod-A
OMs
PSE
PDE Mod-A Loads
PDE Mod-A
C&DH via bus commands to the PSE
and the PDE provides third level of
independent control
3
PDE
Philip Luers - Electrical Systems
25
System Grounding
• Component Grounding
– Primary return isolated from secondary returns (> 1 Mohm). Only
high frequency (i.e. capacitor) connection is allowed
– Secondary grounds tied to chassis ground with low impedence paths
at multiple points on PWB. Analog grounds tied at one point if
possible.
• Single Point Ground (SPG)
– Primary DC current returns to SPG in the PSE, and the PSE will be
tied to spacecraft structure via current shunt
– Solar Array, Battery, SAS, DPC returns are tied to the PSE SPG via
current shunts (PSE backplane)
– Primary returns from downstream switches are back to point of origin
• Subsystem Bonding
– Metal-to-metal contact between component baseplate/mounting feet
to Al face sheet
– Grounding straps will be used where necessary
Philip Luers - Electrical Systems
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System Grounding
•
Structure Grounding
– Bus structure panels, propulsion module, and other primary structure
members will be electrically bonded (< 2.5 milliohm) to form spacecraft
common ground
– Solar array panels, gimbals, booms will be connected to spacecraft structure
with ground straps
– Components on the Instrument Module (composite) will be connected to
spacecraft module with ground straps
Philip Luers - Electrical Systems
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System Grounding
•
Surface Charge Grounding
– Conductive areas to bleed bulk charges, grounded by various methods to spacecraft
structure
– External surfaces surface conductivity < 109 ohms/square
– No more than 10 surfaces >6cm2 in any 1 m2 area
– Insulating films such as Kapton shall be less than 5 mil thick and assembled to bleed
off surface charge, multiple ground straps, 2 ground tabs for every m2, 1 addl ground
tab for each additional m2.
– Conductive paint/finish on antenna surface and solar array panel
– Conductive surface coating on solar cell cover glass frame, with grounding
– Unavoidable dielectrics accepted by waiver only, considering impacts to system.
– Comprehensive list of external surface properties will be assembled and maintained
Philip Luers - Electrical Systems
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System Grounding
•
Ground Straps
– Fabricated of single layer 5 mil copper with 5:1 (length:width) where possible
• Sufficiently low impedence from 0 to 40 MHz
• Multi-layered copper strap as alternative
– Rectangular strip contact area, leveraging SDO design where possible
• Ground strap and ground lug with 80 mm2 contact area as alternative
•
Protection from ESD-Induced Noise
– 40 dB overall shielding from external environment
– Joints, seams and seals between panels shall be sealed with copper tape, conductive
adhesive or EMI gaskets
– No gaps or holes larger than 2.5 cm
– All internal harnesses routed away from uncovered openings
– External circuits filtered at entrance to spacecraft structure wherever possible
• Filtered at electronics box as alternative, with harness separately shielded and routed
– Aluminum tape (LG-1055 or equivalent) at least 1 mil thick, wrapped with 50%
overlap on each 360-degree wrap over the previous wrap shall be used as harness
bundle shield.
Philip Luers - Electrical Systems
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Grounding Diagram
Philip Luers - Electrical Systems
30
EMI Specifications
CE01/CE03 Conducted Emissions Limits
Comp
RF
CE01
X
CE03
X
Orbiter
ES Spec.
STD-7000
461C
X
3.3.1.1
2.5.2.1
2
X
3.3.1.1
2.5.2.1
3
X
3.3.1.2
2.5.2.1
4
CE01 (20 Hz - 14 KHz)
CE06
CE03 (14 KHz-50 MHz)
140
120
Instrument or Component Level
Test Limits (Differential Mode)
100
X
X
3.3.2
2.5.2.2
17
CS01
X
X
X
3.3.3.1
2.5.3.1a
6
CS02
X
X
3.3.3.2
2.5.3.1a
7
CS03
X
3.3.3.3
2.5.3.1b
8
CS04
X
3.3.3.4
2.5.3.1c
9
CS05
X
3.3.3.5
2.5.3.1d
10
X
3.3.3.6
2.5.3.1e
11
80
dBuA
RE02
Instrument or Component Level Test
Limits (Common Mode)
60
CS06
X
RS03
X
X
X
3.3.4
2.5.3.2
40
20
Bandwidth (Hz)
0
1.00E+01
21
1.00E+02
5 Hz
1.00E+03
500 Hz
1.00E+04
X
3.3.5
CS01/CS02 Limits
CS01
•
•
RE02 Radiated Emissions Limits
CS02
30 GHz
60
dB uV/m
5
V rms
1.00E+08
80
6
1.5 kHz
40 Hz 2.8 Vrms
3
2
1
50
40
30
50 kHz
1.0 Vrms
400 MHz
20
14 kHz
Orbiter Levels
408–430
2091.3967 +/- 6
MHz
MHz
Delta IV and Atlas(LRO
V S-band RX)
Components and Instruments Levels
10
0.1
0.001 0.01
1
10
1,000
100
(10 kHz)(100 kHz)
(1 kHz)
10
0.01
0.1
1
100
1,000
10,000
Frequency (MHz)
Frequency (MHz)
CS01/CS02 (injection of energy into power lines) shall be performed on all subsystems
•
RE02 shall be performed on all subsystems
CS03 (two signal modulation), CS04 (rejection of undesired signals), and CS05 (cross
•
RE from components ON at launch shall not exceed lower line
modulation) shall be performed on all RF receiving components
•
RE from components OFF at launch shall not exceed upper line
CS06 (powerline transient) shall be performed on all subsystems
Philip Luers - Electrical Systems
0.00001 0.0001
(10 Hz) (100 Hz)
•
1.00E+07
CE01/CE03 on all +28V pwr & return lines, in diff. and common mode
CE06 will be performed on all RF receivers and transmitters
70
•
•
1.00E+06
50 KHz
FREQUENCY (Hz)
Self
Comp
4
1.00E+05
5 KHz
100,000
31
EMI/EMC RS
LRO Operational Test Limits
Frequency Range
Test Level
Requirement Source
14 KHz – 2 GHz
2 V/m
GSFC-STD-7000
2 GHz – 12 GHz
5 V/m
GSFC-STD-7000
12 GHz – 28 GHz
10 V/m
GSFC-STD-7000
2.271.2 GHz +/- 5 MHz
7 V/m
LRO S-Band Transmitter
25.5 GHz – 28.0 GHz
10 V/m
LRO Ka-Band Indirect Radiation
Delta IV Operational Test Limits
Frequency Range
Test Level
Requirement Source
14 kHz – 40 GHz
20 V/m
Delta IV Launch Pad Environment
2241.5 MHz +/- 650 kHz
40 V/m
Delta IV Second Stage S-band T/M
5765 +/- 6 MHz
40 V/m
Delta IV Second Stage C-band beacon (transmit)
Atlas V Operational Test Limits
•
•
•
•
Frequency Range
Test Level
Requirement Source
14 kHz – 40 GHz
20 V/m (0.5 V/m TBD)
Launch Pad Environment
2211 MHz +/- 4 MHz
77 V/m
Atlas V Second Stage S-band T/M
5765 +/- 6 MHz
117 V/m (*TBD)
Atlas V Second Stage C-band (PEAK transmit)
5765 +/- 6 MHz
4.2 V/m
Atlas V Second Stage C-band (AVG transmit)
RS03 shall be performed on all subsystems
Subsystems that are OFF at launch may be powered OFF during Launch Site / Vehicle RS Tests
Delta IV or Atlas V levels will be used, not both
Atlas Second Stage C-band PEAK transmit represents worst-worst case, contractor will perform analysis of payload on vehicle and provide
realistic test limits when required
Philip Luers - Electrical Systems
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Radiation
•
•
•
Radiation Environment for LRO (431-SPEC-000020) and Radiation Requirements for the
LRO (431-RQMT-00045) have been released
TID requirement 10.8 krad-si (14 month mission, including 2x safety factor)
Single Event Effects (SEE) include Single Event Upsets (SEUs), Single Event Transients
(SETs), Single Hard Events (SHEs), Single Event Latchups (SELs), Single Event Burnouts
(SEBs), Single Event Gate Ruptures (SEGRs), and Multiple Bit Upsets (MBUs)
– SEE Immunity (for destructive effects) is defined as a LETth > 75 MeVcm2/mg, which
must be demonstrated by heavy ion testing or analysis
• No SEE shall cause permanent damage to a system or subsystem.
• For any component that is not immune to SEL or other destructive condition an analysis shall
demonstrate that the SEL probability of occurrence is negligible in the LRO mission
environment
– SEE Immunity (for non-destructive effects) is defined as LETth > 37 MeVcm2/mg,
which must be demonstrated by heavy ion testing or analysis. For components with
LETth < 37 MeVcm2/mg:
• the criticality of a component in its specific application must be defined. Please refer to the
Single Event Effect Criticality Analysis (SEECA) document(431-REF-000273) for details. A
SEECA analysis or a FMEA should be performed at the system level.
• The improper operation caused by single particle events like SEU, SET and MBU shall be
reduced to acceptable levels.
• Systems engineering analysis of circuit design, operating modes, duty cycle, device criticality
etc. shall be used to determine acceptable levels for that device. Means of gaining acceptable
levels include part selection, error detection and correction schemes, redundancy and voting
methods, error tolerant coding, or acceptance of errors in non-critical areas.
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