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AJ10-118

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AKA: LR52-AJ-1. Status: Out of Production. Number: 14 . Thrust: 33.80 kN (7,599 lbf). Unfuelled mass: 90 kg (198 lb). Specific impulse: 271 s. Burn time: 115 s. Diameter: 0.84 m (2.75 ft).

Engine originally developed for the Vanguard launch vehicle, and subsequently developed for use on the Able and Delta upper stages and as the Apollo Service module engine.

The initial Vanguard engine was not a derivative of the Aerobee thrust chamber. This had been Aerojet's original proposal, but had not been accepted by the Navy. The tubular (spaghetti)-walled chamber configuration was mostly influenced by the design experience from the Titan thrust chamber. The Vanguard thrust chamber had both an aluminum and backup stainless steel version. The latter was essentially mandated by Ed Elko (then head of the Thrust Chamber Section) and other thrust chamber elder statesmen within the company, because Aerojet had no experience with aluminum tubular thrust chambers. However, the weight advantage of the aluminum chamber was so great that the first fabrication order included 15 of the aluminum version and only 3 of the stainless steel. The two versions were virtually identical in external configuration, but there were minor differences relating to physical properties of the materials, and welding details.

The main proponent and original designer of the aluminum thrust chamber and injector was Harry Meyers, who had been hired from Bell Aircraft and had been a major contributor to that company's Rascal and Agena engines. A key element of the Vanguard engine was the injector. With a chamber internal diameter of about 8 inches, it consisted of a single 5.5 in. diameter circle of 1-on-l orifices (fuel on oxidizer, with the oxidizer aimed inward). Numerous iterations were required to obtain the desired balance of atomization and mixing necessary to achieve the target performance and stability. What finally did the trick was the addition of a single ring of fuel orifices near the center of the injector, to modify the gas circulation pattern. This was suggested by Sid Rumbold, based on his earlier work at M W Kellogg Company. Considerable analytical and tutorial efforts in the fields of heat transfer and fluid flow were conducted by Ed Elko and John Beerboom in trying to solve the problems of the early chamber wall burnouts and hot spots. This greatly improved the understanding of the situation, and the solution eventually came about by a combination of the injector spray pattern correction with the improved Titanium Carbide (TiC) wall coatings.

By the Fall of 1957, with the concurrence of the customer, the development work on the aluminum thrust chamber was suspended. The last block of second stage propulsion systems had to be released, and the continuing burnout problems with the tungsten carbide coated aluminum tubular chamber made it unacceptable to replace the stainless steel version. However the Air Force selected the aluminum chamber for a subsequent version of the engine used on their Thor-Able program.

The contractually required performance was met with the steel chamber and a respectable profit of more than 5% was realized, despite the early program problems.

Able was the first of many engine and application programs that flowed from the Vanguard experience base. These included Able, Ablestar, Delta, Fat Delta, the Japanese N II, and applications or offshoots such as Hydra, Saint (Satellite Intercept), and other classified programs. Included in all this were numerous upratings and incremental changes in the thrust chambers, tanks, and complete systems. Derivative programs included Transtage and Apollo SPS, and ultimately, the Shuttle OME. Delta thrust chamber assemblies of a considerably advanced configuration were still being produced by Aerojet well into the 21st Century - a total of over 50 years of continuous activity in this family.

The associated large number of different missions, vehicles, stages, and thrust chamber assemblies, and modifications thereof, has led to a nomenclature problem, and considerable confusion as to program details, relationships, and relative timing. A major example of this is that in the early years the Air Force called the vehicles that they procured "Thor-Able" or "Thor-Ablestar," but, NASA called all their Thor-based vehicles "Delta." No matter what they were called, they were all really Vanguard second stages, either with the original or larger diameter tanks. In those days Able or Ablestar meant Air Force, and Delta meant NASA. However, several years later, the name Delta was also applied to Aerojet's ablative thrust chambers and stages, even though some were procured by the Air Force.

Continuing development of the Vanguard aluminum thrust chamber assembly resulted in selection of this system by the Air Force for use with a Thor booster that was to be used to demonstrate the Atlas guidance system, and to explore nose cone reentry problems. This was called the Able program, and it began in November 1957. Thor was basically a single stage IRBM built by Douglas Aircraft that used essentially the same thrust chamber assembly as Atlas, and reached flight status before Atlas. Space Technology Laboratories (STL), and later the Aerospace Corporation (which was formed from part of STL in 1960), acted as system manager for the Thor-Able program and its Air Force successors. The Able system included the thrust chamber assembly, valves, tanks, pressurizing system, and any additional components to make up a complete second stage. The oxidizer was changed from the WFNA used in Vanguard, to RFNA. The first few Thor-Ables were delivered before the formation of NASA.

The minor modification of the Vanguard aluminum tube thrust chamber to meet the Able requirements was accomplished in the record time of only three months. The major effort during this time was the testing of six aluminum tube thrust chambers for durations longer than the full burn time. This was done to develop confidence that the expected burn-through failure in the throat would occur at least 30% beyond the nominal duration, that it would be repeatable, and that the total impulse would be within specification limits. This was accomplished, and it provided the first opportunity for Aerojet's aluminum tube bundle engine to perform successfully in space.

Engine: 90 kg (198 lb). Chamber Pressure: 7.00 bar. Area Ratio: 40. Propellant Formulation: WFNA/UDMH. Thrust to Weight Ratio: 38.3. Coefficient of Thrust vacuum: 13.7657076924562. Coefficient of Thrust sea level: 8.05142197817045.