Unlike most liquid-fueled rocket engines in service at the time, the J-2 was designed to be restarted once after shutdown when flown on the Saturn V S-IVB third stage.

After the crew verified that the spacecraft was operating nominally, the J-2 was re-ignited for translunar injection, a 6.5 minute burn which accelerated the vehicle to a course for the Moon.

[2] The thrust chamber was constructed of 0.30 millimetres (0.012 in) thick stainless steel tubes, stacked longitudinally and furnace-brazed to form a single unit.

The chamber was bell-shaped with a 27.5:1 expansion area ratio for efficient operation at altitude, and was regeneratively cooled by the fuel.

Fuel entered from a manifold, located midway between the thrust chamber throat and the exit, at a pressure of more than 6,900 kPa (1,000 psi).

[2] The thrust chamber injector received the propellants under pressure from the turbopumps, then mixed them in a manner that produced the most efficient combustion.

The dome provided a manifold for the distribution of the LOX to the injector and served as a mount for the gimbal bearing and the augmented spark igniter.

The ASI operated continuously during entire engine firing, was uncooled, and was capable of multiple reignitions under all environmental conditions.

[2] The fuel turbopump, mounted on the thrust chamber, was a turbine-driven, axial flow pumping unit consisting of an inducer, a seven-stage rotor, and a stator assembly.

It was a high-speed pump operating at 27,000 rpm, and was designed to increase hydrogen pressure from 210 to 8,450 kPa (30 to 1,225 psi) (absolute) through high-pressure ducting at a flowrate which develops 5,800 kW (7,800 bhp).

Power for operating the oxidizer turbopump was provided by a high-speed, two-stage turbine which was driven by the exhaust gases from the gas generator.

The four-vane rotor in the hydrogen system produced four electrical impulses per revolution and turned approximately 3,700 rpm at nominal flow.

The six-vane rotor in the LOX system produced six electrical impulses per revolution and turned at approximately 2,600 rpm at nominal flow.

During engine operation, either LOX was tapped off the oxidizer high-pressure duct or helium was provided from the vehicle stage and routed to the heat exchanger coils.

The gaseous hydrogen imparted initial spin to the turbines and pumps prior to gas generator combustion, and the helium was used in the control system to sequence the engine valves.

The electrical sequence controller was a completely self-contained, solid-state system, requiring only DC power and start and stop command signals.

The primary package instrumentation measures those parameters critical to all engine static firings and subsequent vehicle launches.

The instrumentation system was designed for use throughout the life of the engine, from the first static acceptance firing to its ultimate vehicle flight.

[2] After an interval of 0.450 seconds, the start tank discharge valve was closed and a mainstage control solenoid was actuated to:[2] Energy in the spark plugs was cut off and the engine was operating at rated thrust.

[2] During mainstage operation, engine thrust could be varied between 780 and 1,000 kilonewtons (175,000 and 225,000 lbf) by actuating the propellant utilization valve to increase or decrease oxidizer flow.

[2] To provide third stage restart capability for the Saturn V, the J-2 gaseous hydrogen start tank was refilled in 60 seconds during the previous firing after the engine had reached steady-state operation (refill of the gaseous helium tank was not required because the original ground-fill supply was sufficient for three starts).

A source evaluation board was formed to nominate a contractor from five bidding companies, and approval was given on 1 June 1960 for Rocketdyne to begin development of a "high-energy rocket engine, fuelled by LOX and hydrogen, to be known as the J-2".

The final contract, awarded in September 1960, was the first to explicitly require the design "insure maximum safety for crewed flight.

"[4] Rocketdyne launched the development of the J-2 with an analytical computer model that simulated engine operations and aided in establishing design configurations.

Confidence in the design was regained in August, however, when the same stage, S-IVB-201, performed flawlessly on a full-duration firing of 452 seconds, which was the first engine test sequence to be controlled entirely by computers.

Reliability and development testing continued on the engine, with two uprated versions being used by NASA in the later flights of the Apollo program.

[6][7] While work on the J-2S continued, NASA also funded a design effort to use the J-2S turbomachinery and plumbing to a toroidal combustion chamber with a new aerospike nozzle.

What became a different engine with a similar name, called the J-2X,[10][11] was chosen in 2007 for the Project Constellation crewed lunar landing program.

[14] The new J-2X is designed to be more efficient and simpler to build than its Apollo J-2 predecessor, and cost less than the Space Shuttle Main Engine (SSME).

[18] Project Constellation was cancelled by President Barack Obama on October 11, 2010,[19] but development of the J-2X has continued for its potential as the second stage engine for the new, heavy-lift Space Launch System.