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Able-Star
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Gross Mass: 4,470 kg (9,850 lb). Empty Mass: 590 kg (1,300 lb). Thrust (vac): 35.088 kN (7,888 lbf). Isp: 278 sec. Burn time: 296 sec. Propellants: Nitric acid/UDMH. Isp(sl): 210 sec. Diameter: 1.40 m (4.50 ft). Span: 1.40 m (4.50 ft). Length: 5.88 m (19.29 ft). Country: USA. No Engines: 1. Engine: AJ10-104. Cost $ : 5.000 million. First Flight: 1960. Last Flight: 1965. No Launched: 19. No Failed: 3. Used as: Thor Able-Star. As is almost always the case in such programs, the Air Force requested increases in the propulsion system capabilities in an effort to meet their ever-expanding mission requirements. As a result, the stainless steel version of the basic Able engine was selected, and it was uprated to increase thrust 34.7 kN to 37.0 kN and to increase the duration 2-1/2 times (easily done with the stainless steel thrust chamber) - and this configuration was called Ablestar. The Ablestar also included modifications that allowed in-space restarting - a first in the industry. The time required for developing and qualifying the Ablestar propulsion system was eight months, most of which was needed for the design, development and qualification of the much larger propellant tanks and titanium helium spheres. These remarkably short development times was a result of the basic simplicity of the Able design - mainly the low chamber pressure, hypergolic propellants, and gas pressurized propellant tanks. This simplicity also resulted in a number of additional very desirable features:
- The ability to achieve rapid, relatively low cost modifications, and high reliability
for a variety of missions
- The ability to shift back to the aluminum thrust chamber and injector which
provided an extremely good thrust to weight ratio (180, based on 3500 kgf thrust and a weight of 19.5 kg)
- Very light weight tankage based on heat treated 410 stainless steel
- Easily adjustable run time (in the stainless steel version) based on simply varying
the length of the cylindrical section of the tanks.
In addition, the basic philosophy of pressure fed, low chamber pressure and ablative (rather than regeneratively cooled) thrust chambers for upper stage engines produced outstanding reliability and scalability. In a vacuum engine, a low chamber pressure still provides a reasonable expansion ratio, and thus reasonable performance. Secondly, low chamber pressure allows use of a very simple, pressure fed propellant system with relatively light and inexpensive tanks. Thirdly, the low chamber pressure results in lower heat transfer rates, thus making ablative chambers more practical - and they are inherently less expensive, and much mere reliable. And finally ablative chambers greatly simplify restarts in a vacuum environment because there are essentially no problems with cooling jacket and manifold fill times or coking in the coolant system.
Used on Launch Vehicles:
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© Mark Wade, 1997 - 2007 except where otherwise noted.
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