Final Version Gabe Karpati May 17, 2002 Micro-Arcsecond X-ray Imaging Mission, Pathfinder (MAXIM-PF) System Overview

Final Version

Gabe KarpatiMay 17, 2002

Micro-Arcsecond X-ray Imaging Mission, Pathfinder (MAXIM-PF)

System Overview

MAXIM-PF, May 13-17, 2002Goddard Space Flight Center

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Final Version

Requirements & AssumptionsBaseline ConfigurationOptions ConsideredComments, Issues, Concerns

Outline


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Final Version

Requirements & Assumptions Study Overview

Mission objective X-ray interferometry mission, a pathfinder to full MAXIM

Original requirements As formulated in the Prework and in K. Gendreau’s “going-in-13may02.ppt”

Original requirements modified during the study Lifetime for Phase 1: 1 yr required / 50 targets (1wk/target);

Lifetime for Phase 2: 3 yrs required / 4 yrs goal (3 wks/target)

Additional constraints, challenges 2015 launch

Primary purpose of this study Identify mission drivers and breakpoints Identify technologies required Subsystem configuration, mass and cost estimates

Length of study 5 days


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Requirements & Assumptions Major Driving Requirement Areas

High precision pointing Centroid image of a laser beacon for microarcsec LOS alignment Point by referencing microarcsec image of stars or use GPB-like

microarcsec grade Super-Gyro Multi s/c formation flying

Orbital dynamics: Formation acquisition and control; Orbits; Transfer to L2

Propulsion: Thrust needs to vary by several orders of magnitude ACS: Position control to microns over 100’s of m, and to cm’s over

20000 km, knowledge to microns; Retargeting issues Software

To accommodate all functions Verification

Functional and performance verification 1 g environment Thermal control

Handle two thermally very dissimilar mission Phases with one h/w Control to .1 degree to maintain optical figure “STOP” CTE effects

Communication Complex communications web: Detector to Ground; Hub to Detector;

Hub to FFs; FF to FF; Rough ranging using RF


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Baseline Configuration Experiment Overview

Observatory configuration One Hub spacecraft, one Detector spacecraft, six Free Flyer spacecraft Hub communicates with Detector and the Free Flyers Detector communicates with ground

Phase 1: 100 microarcsec Science 2 formation flying objects at 200 km

Phase 2: 1 microarcsec Science Hub surrounded by 6 identical Free Flyers in a circle of 200-500 m,

Detector at 20,000 km Distance from Hub to Detector: RF ranging course & time of flight for

fine ranging and control (~5m) Align Hub and Detector using Superstartracker that centroids the image

at the Detector of a LISA - like laser beacon mounted on Hub (microarcsec)

LOS pointing: reference beacon image to image of stars in background w/ Superstartracker or use GPB - like Super-Gyro (microarcsec)

HUB to FF’s distance: w/ RF ranging course; Laser interferometer fine w/ corner cubes on Hub (~10 um);

FF position: use FF startrackers (~arcsecs)looking at LED on Hub


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•LOS to target knowledge to ~0.1 milliarcsec (~15 microns @ 20,000 km)

FreeFlyer S/C•Pitch, Yaw control to ~1 arcsec•Pitch, Yaw Knowledge to arcsecs •Roll Control to 30 milliarcsecs

Optics Hub S/C• Pitch, Yaw, control

to ~ 1 arcsec, roll control to arcmins

• Pitch, Yaw, Roll Knowledge to +/- 1 arcsecond

Diagram courtesy of K. Gendreau


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Continuous full sun Battery required for safe Phase only

Transfer to L2 Takes up to 6 months All S/C are attached together High thrust chemical propulsion Transfer stage is jettisoned at L2

Communication web HUB to Free Flyers HUB to Detector All Space-Ground communications performed by Detector spacecraft IP, 50 Kbps; One contact day @ DSN 5 Mbps Ranging for collision avoidance


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Baseline Configuration Overview

LV Delta IV 4040 (4M fairing) Delta IV 4240 Delta IV 4450

C3 (km2/s2):- .7

2780kg

C3 (km2/s2):- .7 4,135kg

$97M

C3 (km2/s2):- .7 4,650kg

Orbit Orbit Type Eclipse Boresight

L2, Lissajous, 800,000 km

, ~6 months period

No eclipse. Full Sun, Avoid Moon

shadow

Anywhere within a 30 deg halfangle

cone, perpendicular to sunline

Mission Liftoff Mass Life Size Consumables

~3000 kg Ph1: 1 yr req / 50 targets

(1wk/ target);

Ph2: 3 yrs req (3 wks/ target) /

4 yrs goal

5 yrs; Phases 1 & 2 combined


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Baseline Configuration Overview


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Baseline Configuration Instrument Resources Summary

X- ray Detector System $k kg

Detector Unit 15000 50Electronics Unit 5000 50

X- ray Detector System Total 20000.0 100.0

Mirror Module (one unit) $k kg

Bench / Houing 5

2 Mirrors ttl 100

6 Actuators ttl 30

6 Mechanisms ttl 60

1 Shutter 1

Electr/ Harness 10

Thermal 1

I nstr Eng 40

S/ C level I &T .1 FTE 15

Assy Tech .2 FTE 30

One Mirror Module Total 292.0 5.0


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Baseline Configuration Metrology System Resources

Summary

Hub Metrology System $k kg

LI SA Laser Beacon 5000 20

Hub's LED's (6) and cornercubes (6) Total 400 5

Det Metrology System $k kg

LOS Superstartracker / Gyro Package 70000 120

FF Metrology System $k kg

FF's Ranging I nterferometers, one unit 2000 10


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Baseline Configuration S/c Mass Summaries

HUB TOTAL MASS [kg]

Actual Contingent

PAYLOAD TOTAL 70.0 85.3

BUS SUBSYSTEMS DRY TOTAL 198.6 238.4

Observatory Dry Mass 268.6 323.6

PROPELLANT 62.0 74.4

WET MASS (BUS + PAYLOAD) 330.64 398.0

DET TOTAL MASS [kg]

Actual Contingent






FF ea. TOTAL MASS [kg]

Actual Contingent







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Baseline Configuration Mission Mass Summary

MISSION TOTAL MASS [kg]Actual Contingent

Hub Wet 330.6 398.0

Detector Wet 619.1 753.9

Total 6 Freeflyers Wet 1824.2 2192.1

Transfer Module Wet 500.0 600.0

LIFTOFF MASS 3274.0 3944.0

Delta 4040 capability to L2 2780.0 2780.0

Margin [kg] -494.0 -1164.0

Margin [%] -17.8 -41.9

Delta 4240 capability to L2 4135.0 4135.0

Margin [kg] 861.0 191.0

Margin [%] 20.8 4.6


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Baseline ConfigurationPayload Cost [$M]

Actual w/ Cont. Cont %

9 Mirror Modules 2.6 3.4 30%Metrology System: LI SA Laser, 6 cornercubes, 6 LEDs 5.4 7.0 30%Payload GSE incl inclPayload I &T 18.3 22.8 25%Payload Support incl incl

Payload Total 26.3 33.2

Hub s/c

Hub s/c


X-ray Detector System Total 20.0 26.0 30%LOS Superstartracker / Gyro Package 70.0 100.0

Payload GSE incl incl

Payload I &T incl incl

Payload Support incl incl


Detector s/ c

Detector s/ c


11 Mirror Modules 3.2 4.2 30%FF's Ranging I nterferometers 2.0 2.6 30%Payload GSE incl inclPayload I &T incl inclPayload Support incl incl


Freeflyer s/ c

Freeflyer s/ c


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Baseline ConfigurationHub S/c Subsystems Cost [$M]

HUB SPACECRAFT BUS SUBSYSTEMS* [$M]


Mechanical Cost 0.6 0.7 20%Mechanical Labor 1.1 1.3 20%ACS Hardware 2.7 3.3 20%ACS Design, Analysis, SS I &T 5.2 6.2 20%Thermal Hardware 0.7 0.8 20%Thermal Labor 1.0 1.2 20%Propulsion Hardware 1.0 1.2 20%Propulsion Labor 2.0 2.4 20%

Power Hardware 1.5 1.8 20%

Power Labor 2.8 3.4 20%

C&DH Hardware 5.0 6.0 20%

C&DH Labor 2.0 2.4 20%RF Communication Hardware 4.5 5.4 20%RF Communication Labor (I &T, CTV test) 5.0 6.0 20%Flight Sof tware Hardware 0.2 0.2 20%Flight Sof tware Labor 10.8 12.9 20%

Spacecraft Bus Total 46.0 55.2

Hub s/c


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Baseline ConfigurationDetector S/c Subsystems Cost [$M]

DETECTOR SPACECRAFT BUS SUBSYSTEMS* [$M]








Detector s/ c


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Baseline ConfigurationOne FF S/c Subsystems Cost [$M]

FREEFLYER SPACECRAFT BUS SUBSYSTEMS* [$M]








Freeflyer s/ c


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Baseline Configuration Overall Cost Summary [$M]

TOTAL MI SSI ON COST [$M]Actual w/ Cont.

PROJ ECT MANAGEMENT 17.6 22.0

PRE- LAUNCH SCIENCE TEAM SUPPORT 4.2 5.3

PAYLOAD TOTAL COST 147.6 166.0

SPACECRAFT BUS SUBSYSTEMS 231.2 277.5

MISSION SYSTEMS ENGINEERING 10.8 13.0

ATLO & MISSION READINESS 18.1 21.7

LAUNCH VEHICLE 97.0 97.0

GROUND SYSTEM DEVELOPMENT 5.6 6.7

MISSION READINESS VERIFICATION 1.0 1.3

OPERATIONS FOR 3 YEARS 14.4 18.0

MISSION TOTAL, Full Acct 547.5 628.4


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Additional Issues To ConsiderSmaller RSDO Busses

RSDO On-Ramp II in force RSDO On-Ramp IV selection in process

Several new buses added, to increase choice

Spectrum Astro SA 200B, Bus dry mass = 90 kg Payload Power (OAV) (EOL) / Mass Limit: 86 W / 100 kg

Orbital - Microstar, Bus dry mass = 59 kg Payload Power (OAV) (EOL) / Mass Limit: 50 W / 68 kg

Ball BCP 600, Bus dry mass = 203 kg Payload Power (OAV) (EOL) / Mass Limit: 125 W / 90 kg

Orbital - Leostar, Bus dry mass = 263 kg Payload Power (OAV) (EOL) / Mass Limit: 110 W / 101 kg

Surrey - Minisat 400, Bus dry mass = 207 kg Payload Power (OAV) (EOL) / Mass Limit: 100 W / 200 kg

TRW - T200A, Bus dry mass = 242 kg Payload Power (OAV) (EOL) / Mass Limit: 94 W / 75 kg

SA 200B

BCP 600


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Additional Issues To Consider Bigger RSDO Busses

Swales EO-SP (new in RSDO II catalog) Bus dry mass = 370 kg Payload Power (OAV) (EOL) / Mass : 80 W / 110kg

Spectrum Astro SA 200HP Bus dry mass = 354 kg Payload Power (OAV) (EOL) / Mass Limit: 650 W / 666 kg

Lockheed Martin - LM 900 Bus dry mass = 492 kg Payload Power (OAV) (EOL) / Mass Limit: 344 W / 470 kg

Orbital StarBus Bus dry mass = 566 kg Payload Power (OAV) (EOL) / Mass Limit: 550 W / 200 kg

Orbital – Midstar Bus dry mass = 580 kg Payload Power (OAV) (EOL) / Mass Limit: 327 W / 780 kg

Ball BCP 2000 Bus dry mass = 608 kg Payload Power (OAV) (EOL) / Mass Limit: 730 W / 380 kg

EO-1

Midstar

SA200HP -DS1


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Comments, Issues and Concerns I&T, Requirements Verification

Environmental verification Standard, per GEVS

Any end-to-end testing / verification of the critical subsystems is very difficult or near-impossible in a 1 g environment E-E verification of orbit maintenance and formation flying capabilities

near-impossible E-E verification of metrology system near-impossible E-E verification of X-ray beam focus and alignment is difficult

Reasonable trades must be made on verification approaches, goals, and requirements That alone is a very significant body of work


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Maturity, Technologies, TRL

MAXIM is feasible ! MAXIM does not factor in any unrealistic technology expectations or

technologies un-envisionable today Fairly mature and serious plans, even for the metrology

Still, a staggering amount of technology development is required: Metrology system: H/w and s/w elements

Superstartracker GPB - like Super-Gyro for pointing

Software Formation flying and “virtual-one-body” telescope control software Analysis and simulation techniques

Propulsion system Very low thrust technologies, extremely variable force thrusters

Verification approaches and technologies for FF LAI missions Simulators

Low CTE optical/structural materials General TRL Level of MAXIM key technologies today is 2-3


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Tall Poles

Tall Pole 1: Multi s/c formation flying ACS: Position control to microns over 100’s of m, and to cm’s over

20000 km, knowledge to microns; Retargeting issues Orbital dynamics: Formation acquisition and control; Orbits; Transfer to

L2 Metrology System: swarm sensors, interferometric range sensors,

beacon detecting attitude sensors Tall Pole 2: High precision pointing

Centroid image of a laser beacon for microarcsec LOS alignment Point by referencing microarcsec image of stars or use GPB-like

microarcsec grade Super-Gyro Tall Pole 3: Software

To accommodate all required functions Tall Pole 4: Propulsion

Continuous smooth micro-thrusters Thrusters force variable by orders of magnitude


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Tall Poles

Tall Pole 5: Verification science Theoretical “risk-science” assessment on feasible verification vs.

available resources Functional and performance verification in 1 g environment “STOP” CTE effects

Tall Pole 6: Thermal control Control to .1 degree to maintain optical figure Handle two thermally very dissimilar mission phases with one h/w

Tall Pole 7: Communication Complex communications web: Detector to Ground; Hub to Detector;

Hub to FFs; FF to FF; Rough ranging using RF Tall Pole 8: Mirror element actuators & software

General TRL Level of key technologies today is 2-3


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Additional Issues To Consider

Startracker on FF opposite the Hub – Sun line would stare at Sun Since 6 FF’s are 60 degrees apart, roll entire formation, to have two FFs

closest to Hub – Sun line at equal 30 degrees This concept doesn’t work for a higher number of FF’s, unless FF startracker

FOV is sufficiently narrowed (complicates access to star-field) Structural-Optical-Thermal effects

Not fully addressed yet Thermal control to 1.5 mK required – not trivial ! Lower CTE optical/structural materials?

Structural stability between the attitude sensor and the instrument It is good practice to mount the attitude sensors and the instrument on a

common temperature controlled optical table Free Flyers station fixed

Free Flyer station clocking position in circle around Hub is constrained To change position, while keeping mirrors in alignment requires rolling the FF s/c Rolling of FF s/c is disallowed for sun / anti-sun sides must be pointed right

Mounting FF Mirror Assemblies on turntable would allow repositioning of any FF s/c to any station


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Additional Issues To Consider

Other mission orbits should be fully explored Earth leading/trailing drift away orbit at .1 AU/year Distant retrograde orbits Solar-libration: “kite-like” solar sail “floating” on a toroid-like pseudo-

libration surface which envelops L1 between Sun-Earth Calibration Plan

Calibration may be a major requirements driver, must be factored in early on

Communications network architecture Communications between constellation elements: much refinement is

required TDRSS at L2? Servicing at L2?

Explore synergies and joint funding possibilities w/ other LAI missions at L2

Servicability at L2 Design shouldn’t of the bat preclude future serviceability Coordinate w/ servicing planners


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Supporting Data

Systems spreadsheet tool: “LAI-MAXIM-PF_System_Sheets.xls” System configuration summaries Mass and cost rollups and detailed ISIS subsystem data Quick propulsion calculator Prework information

WBS template: “Generic_WBS_Template_by_GSFC_NOO.doc” Full NASA mission’s complete Work Breakdown Structure Compiled by GSFC New Opportunities Office

Useful web sites Access to Space at http://accesstospace.gsfc.nasa.gov/ provides launch

vehicle performance information and other useful design data. Rapid Spacecraft Development Office at http://rsdo.gsfc.nasa.gov/

provides spacecraft bus studies and procurement services.


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System Summary

GSFC Contact: Keith Gendreau Phone Number: 301/286-6188 Mission name and Acronym: MAXIM-Pathfinder Authority to Proceed (ATP) Date: Dec 2007 Mission Launch Date: 2015 Transit Cruise Time (months): n/a Mission Design Life (months): 48 Length of Spacecraft Phase C/D (months): 72 Bus Technology Readiness Level (overall): 3 S/C Bus management build: TBD Experiment Mass: 3000 kg

Documents

Final Version Gabe Karpati May 17, 2002 Micro-Arcsecond X-ray Imaging Mission, Pathfinder (MAXIM-PF) System Overview