Design Considerations for Space Transfer Vehicles Using Solar Thermal Propulsion
Abstract
The economical deployment of satellites to high energy earth orbits is crucial to the ultimate success of this nation's commercial space ventures and is highly desirable for deep space planetary missions requiring earth escape trajectories. Upper stage space transfer vehicles needed to accomplish this task should ideally be simple, robust, and highly efficient. In this regard, solar thermal propulsion is particularly well suited to those missions where high thrust is not a requirement. The Marshall Space Flight Center is, therefore, currently engaged in defining a transfer vehicle employing solar thermal propulsion capable of transferring a 450 kg payload from Low Earth Orbit (LEO) to Geostationary Earth Orbit (GEO) using a Lockheed Launch Vehicle (LLV3). The current design uses liquid hydrogen as the propellant and employs two inflatable elliptical off-axis parabolic solar collectors to focus sunlight onto a tungsten/rhenium black body type absorber. The concentration factor on this design is projected to be approximately 1,800:1 for the primary collector and 3:1 for the secondary collector for an overall concentration factor of nearly 5400:1. The engine, which is about twice as efficient as the best currently available chemical engines, produces 8.9 N of thrust with a specific impulse (Isp) of 860 sec. Transfer times to GEO are projected to be on the order of one month.
Document Details
- Document Type
- Technical Report
- Publication Date
- Jul 01, 1995
- Accession Number
- ADA409502
Entities
People
- William J. Emrich Jr.
Organizations
- National Aeronautics and Space Administration