Engine Model: J-2S. Manufacturer Name: J-2S. Designer: Rocketdyne. Developed in: 1967. Propellants: Lox/LH2. Thrust(vac): 1,138.500 kN (255,945 lbf). Isp: 436 sec. Burn time: 475 sec. Mass Engine: 1,400 kg (3,000 lb). Diameter: 2.01 m (6.60 ft). Length: 3.38 m (11.08 ft). Chambers: 1. Chamber Pressure: 30.00 bar. Area Ratio: 40.00. Oxidizer to Fuel Ratio: 5.50. Country: USA. Status: Developed 1965-1969. J-2 version proposed for Saturn follow-on vehicles, using results of the J-2X technology program. The engine was simplified while offering improved performance. The J-2S also provided the basis for X-33 linear aerospike engine thirty years later. It was resurrected yet again in 2005 when NASA proposed it to power the trans-lunar injection stage of its Cargo Launch Vehicle. Finally, a further development of the engine was selected to power the second stage of the Ares I Crew Launch Vehicle for the Orion space capsule.
The original J-2X program, conducted between 1964 and 1967, developed most of the simplification features of the J-2S, including the tap-off, and pump-fed engine cycle.
The J-2S (J-2 Simplified) engine was originally developed as a replacement for the J-2 Saturn vehicle upper stages, stages 2 and 3 on the Saturn V, and Stage 2 on the Saturn IB. The intent of the design changes was not only to provide performance upgrades to the engine but to greatly simplify the production and operation of the engine. The J-2S engine and components were developed between 1965 and 1972 and the effort was based on experimental engines tested between 1964 and 1968 (the J-2X engine series). The J-2S program consisted of six flight configuration engines tested at both sea level and vacuum conditions in 273 tests for a total operational experience of 30,858 seconds. At the completion of the program the engine was fully developed and ready to go into certification for flight operations.
The nominal thrust, in vacuum, of the engine is 116,000 kgf with a specific impulse of 436 seconds, with a 40:1 nozzle expansion ratio. Baseline operation was at a mixture ratio of 5.5, oxidizer to fuel, with the capability to operate at mixture ratios of 5.0 and 4.5 on command for optimized propellant utilization during the mission. All engine interfaces were located such that the engine could be used as a direct substitute, in form, fit and function, for the J- 2. The version proposed by NASA in 2005 would have a specific impulse of 451.5 seconds and a thrust of 124,511 kgf (evidently with a greater expansion ratio nozzle).
Throttling capability was added as an option for applications other than the Saturn program. The engine also included a low thrust operational feature known as Idle Mode. This was to be used for propellant tank settling, on-orbit maneuvering, and rapid engine chill down prior to firing.
The J-2S simplification was primarily in the change from a gas generator engine cycle to simpler tap-off engine cycle. This cycle allowed for the elimination of the gas generator by supplying hot gas for the turbines from the combustion chamber, diluted considerably with liquid hydrogen. This simplification eliminated some of the timing difficulties associated with the start-up of multiple combustion devices. The oxidizer in this cycle was pumped to pressure and ducted into the injector. The fuel was pumped to pressure and used to cool the combustion chamber and immediately diluted with some of the cold hydrogen from the coolant circuit. This warm hydrogen and steam were used to power the high-pressure ratio turbines and were then ducted back to the nozzle at a suitable location to be expanded with the engine's exhaust plume.
The engine start-up cycle was driven by a solid propellant gas generator, which was electronically initiated at the proper moment in the start sequence. These solid propellant turbine starters were arranged in a manifold of three units for those configurations requiring multiple starts. Each cartridge was fired in sequence and did not require protection from the ignition of the adjoining cartridges.
The proper propellant utilization mixture ratio was controlled by a valve, which varied oxygen recirculation flow around the oxidizer pump. For lower mixture ratio, more oxidizer was recirculated. Idle mode was made possible by adding an injector, thus maintaining a stable pressure drop in each of these elements even during the low propellant flows seen during operation in this low thrust mode. Throttling was accomplished by closing down the hot gas valve, thus reducing the available power of the turbine. The effectiveness of this technique was verified by a hot-fire test.
The J-2S, while not a qualified in-production engine, was not a paper engine. It was an engine with significant flight heritage in the J-2 Saturn program engine, and it had significant ground test experience. It was not put into production only because a follow-on order for Saturn launch vehicles never materialized. Because of the excellent flight history of its heritage system, and because of its "almost ready for flight" status, the J-2S was considered by NASA for a number of applications after its development. Detailed consideration was given during the Space Shuttle development for its use, as well as in the early stages of the NASA Advanced Launch System Program (ALS).
It was estimated by ATK Thiokol in 2005 that restarting the J-2S program, including engine fabrication, design and reliability verification, certification, and production, would require four years. Although no J-2S tooling was known to exist, modern soft tooling could be developed quickly and less expensively than the original hard tooling. There was an existing manufacturing and supplier network in place to support a J-2S restart.
In the event, NASA was unable to resist 'improving' the J-2S, and by early 2007 the engine for the second stage of the Ares 1 Crew Launch Vehicle was the redesignated and substantially different J-2X.
Engine Model: J-2X. Designer: Rocketdyne. Developed in: 2012. Application: Ares I launch vehicle second stage. Propellants: Lox/LH2. Thrust(vac): 1,310.000 kN (294,490 lbf). Isp: 448 sec. Burn time: 431 sec. Mass Engine: 2,430 kg (5,350 lb). Diameter: 3.05 m (10.00 ft). Length: 4.70 m (15.40 ft). Chambers: 1. Area Ratio: 80.00. Oxidizer to Fuel Ratio: 5.50. Country: USA. Status: In development 2006-2012. The J-2X concept can be traced back to 2005. It began as an update to the J-2S, a simplified version of the J-2 used on the Saturn launch vehicles. But by the time NASA and its subcontractor had studied the design, applied "essential" current technical specifications, materials, and standards, they wound up with a design with 20% more thrust, 3% greater specific impulse, but nearly double the weight of the original.
The J-2S had been developed and tested in the early 1970s for follow-on production runs of the Saturn IB and Saturn V launch vehicles. It was shelved after further production was cancelled. Some J-2S hardware was revived for use on the J-2+ linear aerospike engine for the X-33 in the 1990's, but this was also cancelled. It should be noted that the J-2X designation was also used for the J-2 technology program conducted in 1964-1967 that led to the J-2S.
As a "risk reduction" measure, NASA originally decided to develop two variants of the engine, designated J-2X and J-2XD. The J-2X would be developed first, for the Ares I upper stage and low-earth orbit operations. It had a lower combustion chamber pressure, using gas generators for the turbopumps derived from the original J-2 engine, and had a nominal thrust of 1220 kN. The J-2XD would be developed only after the J-2X development was completed (perhaps never if basic J-2X performance was acceptable). It would have a nominal thrust of 1310 kN, achieved by using a higher chamber pressure and a modification of the gas generator designed for the RS-68 engine that powered the Delta IV. Both engine versions used of the Mark 29 gas generator cycle series turbopump design developed for the J-2S and further developed in the 1990's for the J-2+. The turbine exhaust gas would go into the nozzle, augmenting thrust and also providing film cooling to the nozzle extension.
The main injector was of a new design, but contained coaxial elements similar to the J-2 engine. The main combustion chamber had a copper liner with milled channels and a HIP-bonded jacket using methods developed for the RS-68. The two-part exhaust nozzle used Volvo Aero technology developed for the Vulcain engine that powered the Ariane 5. The regeneratively-cooled upper nozzle used a patented Volvo Aero sandwich design. The lower nozzle extension was cooled by supersonic film injection of turbine exhaust gases, developed for the Vulcain 2 engine. The regeneratively cooled nozzle had the same 40-to-1 expansion ratio as the J-2S. The nozzle extension increased the expansion ratio to 80:1 and this larger expansion ratio was the main contributor to the engine's high specific impulse.
The ignition system consisted of an augmented spark igniter similar to the J-2 design, but incorporating improvements derived from the Space Shuttle Main Engine. The J-2X used J-2 valve designs updated for NASA's 2006 seal materials and structural requirements. A few heritage actuators were used where feasible, but most were new or modified to incorporate NASA's 2006 fault tolerance requirements. The J-2X incorporated a new digital engine controller designed to meet fault tolerance and failure detection, isolation, and recovery requirements. The system used open-loop controls to ensure that the design was as simple as possible.
Even with these substantial "modifications" to the original J-2 design, the J-2X was still found to vastly cheaper and less risky than the original concept of modifying the Space Shuttle Main Engine for in-flight restart.
The J-2X for the initial Ares I application could be relatively simple - it might not even have to be restartable. However for the Ares V lunar mission Earth Departure Stage the engine would have to be restarted after loitering in Earth orbit for up to 95 days. This would require development of new-technology methods of reliquefying boil-off of the stage's cryogenic liquids, and elaborate thermal protection systems to handle the more extreme conditions.
Development began in October 2005 with the assumption that a derivative of the Shuttle SSME would be used for the Ares I upper stage and a J-2 derivative for the Ares V Earth Departure Stage. By January 2006 the launch vehicle configurations were refined to drop the SSME on cost grounds, and use the J-2X in both stages. A J-2X Upper Stage Engine Element Office was created within NASA. They convened a 'grey beard' team of surviving J-2 engineers from the Apollo era to discuss historical problems with the engine and suggest possible design approaches for improvement. The team also provided a valuable list of "lessons lived' during the Apollo era in testing, hardware evaluatios, engineering fixes and problem-solving approaches.
The Preliminary Requirements Review of the conceptual engine design and development planning was completed in June 2006. Testing of subsystem components using modifications of existing hardware began as early as April 2006, and a full-scale version of the new injector was to be completed and begin testing in 2007. The original program plan called for the Preliminary Design Review in May 2007, followed by the Critical Design Review in May 2008. The first production-type article would begin testing in 2010, leading to Design Certification for flight in 2012.
