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Aug.19, 1999
George T. RoachIntegration Mission Design CenterNASA- GSFCCode 543Greenbelt, MD. 20771301-286-2213FAX 301-286-0343E-Mail- [email protected]
Overview
• See if the dual spacecraft MAXIM configuration can be accommodated on a single U.S. launch vehicle.
• Present a trade study comparing launch vehicles
• Present a concept that can meet mission requirements:- Launch and orbit configuration
* Payload Instruments placements* Structural mass summary
Spacecraft Mechanical Overview
• The baseline will be 2 Mission Unique Spacecraft ( MUS): - Use of commercial parts and subsystems
- Accommodate the Detector Instruments and Optic Instruments - Provide accommodations for a dual wing fixed solar array panels (total area= 6 meters ^2),
for the detector spacecraft.
- Provide accommodations for a fixed semi ring solar array panels (total area= 7 meters ^2),for the optic spacecraft.
- The Optics spacecraft will interface to the Delta IV 5m dia. X 19.8 m long/ 17362 PAF
- The core structure will be made-up of aluminum honeycomb panels, aluminum truss, fittings, brackets, diagonals and stringers.
Note- Due to the time element in preparing this presentation, I baseline aluminum material, howeverdue to the thermal sensitivity the structure material most likely would be composite. This would increasecost and fabrication complexity, but would decrease mass. In either case, this is do-able in the design.
Launch Vehicle Trade StudyBased on Performance to C3 = 0; Mass = 4100kg;
Volume = 3.25 dia X 14.5 m
Access Mode Candidates
No Foreign Launch Vehicles Investigated (Space Policy-buy American)
Delta II and III do not meet C3 performance requirements
Delta IV:
S,M, M+(4,2 &5,2 config.) do not meet C3 performance requirements
M+ (5,4) config. meets C3 performance requirements but does not meet fairing volume envelope
Heavy meets C3 performance requirements and fairing volume envelope
Atlas V:
Core and core plus 2 strap-on do not meet C3 performance requirements
Core plus 4 or 5 strap-on meet C3 performance requirements but does not meet fairing volume envelop for the 5m short or medium PLF
Heavy meets C3 performance requirements and fairing volume envelope
Lockheed Martin studying (via a KSC task) whether the stretched version of the medium PLF (5.4 m dia.X 26.5 m) can be used on the medium EELV launch vehicles (Results expected in a couple of months)
Boeing studying (via a KSC task) whether the heavy fairing (5 m dia. X 19.8 m) can be used on the medium EELV launch vehicles (Results expected in a couple of months)
Vehicle Capacity to GTO (kg) 1 P/L Weight to C3=0 (kg) Comments
Delta II/7920 1300 910 Estimated (70% of GTO)Delta II/7925 1869 1308.3 Estimated (70% of GTO)
Delta III/6 GEMs 2690 1883 Estimated (70% of GTO)Delta III/9 GEMs 3810 2667 Estimated (70% of GTO)
Delta IV S 2180 1526 Estimated (70% of GTO)Delta IV M 3900 2730 Estimated (70% of GTO)Delta IV M+ (4,2) 5300 3849 (C3 = 0.3)Delta IV M+ (5,2) 4350 2786 (C3 = 0.3)Delta IV M+ (5,4) 6120 4159 (C3 = 0.3)Delta IV H 14200 9306 (C3 = 0.3)
Atlas V (MLV-A Government) 5380 3465
Atlas V Core (MLV-G Comercial) 4100 2698 P/L weight may be based on 5m short PLF2
Core + 2 Strap-ons 6000 4149 P/L weight may be based on 5m short PLF2
Core + 4 Strap-ons 7600 5278 P/L weight may be based on 5m short PLF2
Core + 5 Strap-ons 8200 5649 P/L weight may be based on 5m short PLF2
Atlas V Heavy 12485 8040
1 = Value from ISIR Report or Boeing/LMA generated information2 = Going to 5 m medium PLF, expected performance degradation between 59.3 to 66.6 kg
Launch Vehicle Performance
Baselined 8/18/99
For this study the Delta IV/5m/19.8m was baseline
Launch vehicle will:
• Meet lift-off (mass) requirements +20%(actually approx.. 50%, as of 8/19 AM)
• Meets volume requirements
LAUNCH VEHICLE
Delta IV (H) 5m diameter x 19.8m long
Launch Fairing Removed
Detector Spacecraft(2.2m)
Optics Instruments(10m)
Optic Spacecraft Systems(2.2m)
Baffling
LAUNCHCONFIGURATION
15.5 m
16.4 m
Payload
Spacecraft
ORBIT CONFIGURATION
Optic Spacecraft
Detector Spacecraft
Solar Array(7 m^2, projected area)
Different viewsof the Optic space-craft
Spacecraft sub-System This view the spacecraft subsystems removed
OPTIC SPACECRAFT (DIFFERENT VIEWS)
This is the far end away fromthe detector spacecraft
This is the viewing end towardthe detector spacecraft
Solar Array
Propulsion Tanks(4)
OPTIC SPACECRAFT (ON ORBIT)
Range lasers (4)
Aspect Interferometer
Walter Telescope
Aspect Interferometer
Walter Telescope Electronics
X-Ray Interferometer
OPTIC Instrument(rotated)
OPTIC Instrumentwith support structure
Note- The X-Ray Interferometerhas been removed form clarity
OPTIC INSTRUMENT
Stowed
Orbit
Payload
Spacecraft
Spacecraft Subsystems are mounted inthis volume
FixedSolar Array
(6m^2 shown)
DETECTOR SPACECRAFT
Detector / CCD/ QC Cryogenics and Baffle
Range Sensors and Baffle
DETECTOR PAYLOADINSTRUMENTS
Detector Baffle
Range Sensors
Range Sensors Baffle
Detector / CCD/ QC Cryogenics
Enlarged View of Baffle
DETECTOR SPACECRAFT
Payload Volume
TOP LEVEL SPACECRAFT BUS MASS SUMMARY
Structure 1051.7
Structure (Optics SC) : 906.6Cylinder (3m dia, x 10m long) 554.0Misc Attachment, Cylinder 110.8Stiffening Structure (Cylinder / Detector S/C) 143.0Thrust Tube 58.0Gusset (8) 34.0Misc Attachment, Thrust Tube/Gussets 6.8
Structure (Detector SC) : 145.2Core Structure - Detector S/C 40.0Misc Attachment 10.0Interface Truss Structure 95.2
Mass(kg)/perunit
Total Mass(kg)
TOP LEVEL SPACECRAFT BUS MASS SUMMARY
Mass(kg)/perunit
Total Mass(kg)
Structure (composite) 737.5
Structure (Optics SC) : 638.3Cylinder (3m dia, x 10m long) 365.6Misc Attachment, Cylinder 110.8Stiffening Structure (Cylinder / Detector S/C) 94.4Thrust Tube 38.3Gusset (8) 22.4Misc Attachment, Thrust Tube/Gussets 6.8
Structure (Detector SC) : 99.2Core Structure - Detector S/C 26.4Misc Attachment 10Interface Truss Structure 62.8
Amendment 8/20/99
Amendment 8/20/99
SYSTEM LEVEL MASSSUMMARY(preliminary)
Conclusions:
• Mechanically this mission is feasible:
- As of this phase of the study the dual spacecraft configuration can be accommodated on a single launch vehicle.
* Using the Delta IV, DIV-H 5 m diameter x 19.8 m long fairing* Meets the minimum 20% mass margin* Meets the volume margin
• Launching two spacecraft always lower the frequencies due to the two separations, however with the extra mass margins you can increase the structural integrity of the structure byadding structural members and increasing cross sectional areas.
• The separation system between the Optic spacecraft and Detector spacecraft has yet to be worked out.This is not a trivial matter, I’m not aware of a 3 m + diameter marman ring at this time.
Note- According with the Titan IV folks, anything that fit's inside this fairing the PAF can handlewith reference to C.G. locations.
