RESEARCH ON SPACECRAFT AND POWERPLANT INTEGRATION PROBLEMS. Second Quarterly Report, July 26 to October 26, 1963

Technical Report · Sat Oct 31 00:00:00 EST 1964

DOI:https://doi.org/10.2172/4121413· OSTI ID:4121413

Larson, J W

Major emphasis is on the design of the 1-Mw(e) Rankine cycle powered spacecraft. The heliocentric phases of planetary missions are analyzed and results obtained for Jupiter, Saturn, Uranus, Neptune, and Pluto. Effort is continuing on Mercury heliocentric, out-of-ecliptic, and solar probe trajectories. An interesting result of the heliocentric studies is the reduction of data to a characteristic distance relating mission time, specific impulse, and spacecraft mass ratio. A LEADER program is prepared for use in optimizing spacecraft parameters for selected NAVIGATOR missions. The design study of spacecraft configuration centered about the choice between a deployable and a non-deployable radiator for the 1-Mw(e) turboelectric system. The launch requirements are satisfied by the conical radiator configuration. The radiator weight, primary plus auxiliary cooling, is estimated at 9,800 pounds, which is 9.8 lb/kw(e) and 3.4 lb/ft/sup 2/ of radiating surface, including headers, feeders, structure, and other support items. Preliminary shielding studies indicate that the shield is not a major spacecraft item. Scattering from the radiator to the payload is significant and scatter shielding must be provided. The shield can probably be passively cooled. The electrical system is quite complex because of the many services performed. A first pass circuit diagram is studied. The guidance and control portion of the electrical system consists of using some of the electric thrusters for attitude control. The power generation system is estimated at about 8,000 pounds for a 1-Mw(e) system, exclusive of reactor, shield, and radiators. Thus, the complete powerplant specific weight is around 28 lb/kw input to the electric thrusters or other payload items. The study is centered on the 2600-lb turbogenerator assembly. Electric power and dimensional requirements of various electric thrusters are estimated. The principal state-ofthe-art limitation is electrode erosion, which limits efficiency, size, and weight. Payload weight data are amplified. (auth)

• Research Organization: General Electric Co. Missile and Space Div., Valley Forge Space Technology Center, King of Prussia, Penna.
• DOE Contract Number: NAS 3-2533
• NSA Number: NSA-18-006510
• OSTI ID: 4121413
• Report Number(s): 63SD886
• Resource Relation: Other Information: Orig. Receipt Date: 31-DEC-64
• Country of Publication: United States
• Language: English

Related Subjects

REACTOR TECHNOLOGY
CIRCUITS
CONFIGURATION
CONTROL
COOLANT LOOPS
COOLING
CURRENTS
DIAGRAMS
EFFICIENCY
ELECTRICITY
ELECTRODES
ELECTRONIC EQUIPMENT
EROSION
FAILURES
GAS COOLANT
GENERATORS
HEAT EXCHANGERS
JUPITER
MASS
MATERIALS TESTING
MECHANICAL STRUCTURES
MERCURY
NEPTUNE
NEUTRONS
ORBITS
PLANETS
PLANNING
PLUTO
POWER PLANTS
PROPULSION
RADIATION PROTECTION
RANKINE CYCLE
REACTOR CORE
REACTOR SAFETY
RESEARCH REACTORS
ROCKETS
SCATTERING
SHIELDING
SHIELDING MATERIALS
SPACE FLIGHT
SPACE VEHICLES
STANDARDS
SUN
SURFACES
THERMODYNAMICS
TRANSPORT
TURBINES
VELOCITY
VOLUME
WEIGHT