The original RD-0120 engines, Russian equivalent to the US Space Shuttle Main Engine, are mothballed at Baikonur. Despite various proposals for their use, future production is unlikely.
Credit: © Mark Wade
Kosberg lox/lh2 rocket engine. 1961 kN. Energia core stage. Design 1987. Isp=455s. First operational Russian cryogenic engine system, built to the same overall performance specifications as America's SSME, but using superior Russian technology.
The result was an engine of similar size, thrust, and specific impulse but lower cost. Feed Method: 35,000 rpm dual-stage turbopump. A single pre-burner burns fuel-rich at 527 Celsius to drive the single-shaft high pressure turbopump. Some of the pre-burner gas drives the oxygen low pressure pump. The fuel low pressure pump is driven by GH2 from the main chamber cooling loop.. Throttle range is 45%. Throat diameter 261 mm, exit diameter 2420 mm. The original RD-0120 engines are mothballed at Baikonur.
Although the SSME may have been the starting point, Soviet engine technology led that of the United States in many other detailed points of liquid rocket design. By the mid-1960's the USA had practically abandoned development of liquid fuel engines, with the sole exception of the SSME. The US military preferred to use solid rocket motors for missile and booster stage applications. Russian rocket engineers had spent their entire lives perfecting military liquid fuel rockets and had never favoured solid fuel. Therefore Russian Liquid Oxygen/Kerosene and N2O4/UDMH engines were of much higher performance than those in the US. On the other hand the Soviet Union had not developed any Lox/LH2 engines over 40 tonnes thrust and none actually been flown in space. The contribution of unique Soviet technology and the inevitable changes that occurred during development resulted in the MKS RD-0120 main engine being different in detail from the SSME while retaining the same performance.
In the first stages of the development of the RD-0120, different basic engine schemes were evaluated before a single-shaft turbo-pump for both liquid hydrogen and liquid oxygen was selected (the SSME had separate turbo-pumps for each fuel component). Use of a single pump simplified the engine control system and manufacturing, but also required more detailed and sophisticated methods of design and optimization then were available to the Americans. Another principal difference was the absence of resonance chambers, which were used on the SSME for suppression of high frequency vibrations in combustion chamber. The start sequence developed for the RD-0120 was and remains completely unique.
On the other hand Russian engineers observe that the SSME designers used some technologies that were not used previously in the USA but were common in Russia. The best example was the milled combustion chamber, widely used on Russian engines, but never before on American engines.
By 1999 the US was studying incorporation of more RD-0120 technology into the SSME:
The channel-wall nozzle is a proposed replacement for the current SSME nozzle. Employing a process developed in Russia and used for the Russian RD-0120 rocket engine, flat stock is roll formed into a conical shape, which serves as the nozzle liner. The liner is slotted to form channels for the nozzle's liquid hydrogen coolant to flow through. A jacket is then installed over the liner and welded at the ends. The entire assembly is then furnace brazed. The channels in the liner take the place of the 1,080 tubes that regeneratively cool the current SSME nozzle. The channel-wall nozzle is a relatively simple design that has fewer parts and welds than the current complex SSME nozzle. (The current SSME nozzle takes two-and-one-half years to build, costs $7 million, and is currently flown no more than 12 to 15 times because of safety concerns related to hydrogen leaks.) NASA expects the channel-wall nozzle to be more reusable than the current nozzle and to have less risk of critical failure. The new nozzle is also expected to improve engine performance slightly (although any gain in payload capacity may be canceled by the increased nozzle weight), to cost less and take less time to produce, and to cost less to operate. NASA and Rocketdyne (through Aerojet) have spent $0.8 and $1.2 million respectively to study this upgrade, and development could start at the beginning of 1999. The proposed upgrade would cost an estimated $63 million over four years for development and testing, plus an additional $71 million to build 18 certification and production nozzles.
Drawing on this blend of mature American technology and Soviet innovation, the RD-0120 had a relatively trouble-free development program. The final engine represented for the Soviet Union new technical solutions in engine reliability, control, throttleability, and performance. These were the first fully throttleable Soviet engines, and their first production Lox/LH2 engines.
Application: Energia core stage.
Thrust (sl): 1,517.100 kN (341,058 lbf). Thrust (sl): 154,702 kgf. Engine: 3,450 kg (7,600 lb). Chamber Pressure: 218.00 bar. Area Ratio: 85.7. Thrust to Weight Ratio: 57.97. Oxidizer to Fuel Ratio: 6.
AKA: RO-200; RD-0120; 11D122.
More... - Chronology...
Status: Design 1987.
Unfuelled mass: 3,450 kg (7,600 lb).
Height: 4.55 m (14.92 ft).
Diameter: 2.42 m (7.93 ft).
Thrust: 1,961.00 kN (440,850 lbf).
Specific impulse: 455 s.
Specific impulse sea level: 359 s.
Burn time: 600 s.
First Launch: 1976-90.
Number: 10 .
Associated Launch Vehicles
Albatros Unique Russian space shuttle design of 1974. Hydrofoil-launched, winged recoverable first and second stages. Hydrofoil would have been propelled to launch speed by the launch vehicles rocket engines, using a 200 tonne fuel store in the hydrofoil. Advantages: launch from the Caspian Sea into a variety of orbital inclinations, variations in launch track possible to meet range safety requirements. Proposal of Alexeyev/Sukhoi OKBs. More...
Vulkan Super heavy-lift version of Energia with six strap-on boosters, and in-line upper stages and payloads. The concept was put on the back burner when Energia / Buran development begun. More...
Energia The Energia-Buran Reusable Space System (MKS) began development in 1976 as a Soviet booster that would exceed the capabilities of the US shuttle system. Following extended development, Energia made two successful flights in 1987-1988. But the Soviet Union was crumbling, and the ambitious plans to build an orbiting defense shield, to renew the ozone layer, dispose of nuclear waste, illuminate polar cities, colonize the moon and Mars, were not to be. Funding dried up and the Energia-Buran program completely disappeared from the government's budget after 1993. More...
Interim HOTOL Initiated by a British Aerospace team led by Dr Bob Parkinson in 1991, this was a less ambitious, scaled-back version of the original HOTOL. The single-stage to orbit winged launch vehicle using four Russian rocket engines. It was to have been air-launched from a Ukrainian An-225 Mriya (Dream) aircraft. Interim HOTOL would separate from the carrier aircraft at subsonic speeds, and would then pull up for the ascent to orbit. It would return via a gliding re-entry and landing on gear on a conventional runway. Interim HOTOL suffered from the same aerodynamic design challenges as HOTOL and went through many, many design iterations in the quest for a practical design. More...
Associated Manufacturers and Agencies
Kosberg Russian manufacturer of rocket engines. Kosberg Design Bureau, Russia. More...
Lox/LH2 Liquid oxygen was the earliest, cheapest, safest, and eventually the preferred oxidiser for large space launchers. Its main drawback is that it is moderately cryogenic, and therefore not suitable for military uses where storage of the fuelled missile and quick launch are required. Liquid hydrogen was identified by all the leading rocket visionaries as the theoretically ideal rocket fuel. It had big drawbacks, however - it was highly cryogenic, and it had a very low density, making for large tanks. The United States mastered hydrogen technology for the highly classified Lockheed CL-400 Suntan reconnaissance aircraft in the mid-1950's. The technology was transferred to the Centaur rocket stage program, and by the mid-1960's the United States was flying the Centaur and Saturn upper stages using the fuel. It was adopted for the core of the space shuttle, and Centaur stages still fly today. More...
Haeseler, Dietrich, Information material from Chemical Automatics Design Bureau, Voronezh 1993 via Dietrich Haeseler.
Golubev, A A, KB KhimAvtomatiki - Straniy Istorii, Vol. 1, Voronezh 1995 via Dietrich Haeseler.
Gontcharov, Orlov, Rachuk, Rudis, Shostak, McIllwain, Starke, Hulka, "Tripropellant Liquid Rocket Engine Technology Using a Fuel-Rich Closed Power Cycle", ONERA, June 1995 via Dietrich Haeseler.
Russian Arms Catalogue, Vol 5 and 6, Military Parade, Moscow via Dietrich Haeseler.
Albatros Carrier Aircraft Lox/LH2 propellant rocket stage. Loaded/empty mass 1,250,000/210,000 kg. Thrust 7,840.00 kN. Vacuum specific impulse 455 seconds. Configuration: delta wing with wingtip vertical stabilizers and canards. Engine type and performance, empty weight estimated. More...
Energia Core Lox/LH2 propellant rocket stage. Loaded/empty mass 905,000/85,000 kg. Thrust 7,848.12 kN. Vacuum specific impulse 453 seconds. More...
Energia EUS Lox/LH2 propellant rocket stage. Loaded/empty mass 77,000/7,000 kg. Thrust 1,962.03 kN. Vacuum specific impulse 455 seconds. More...
Interim HOTOL Lox/LH2 propellant rocket stage. Loaded/empty mass 250,000/33,100 kg. Thrust 7,840.00 kN. Vacuum specific impulse 455 seconds. More...
Vulkan 1 Lox/LH2 propellant rocket stage. Loaded/empty mass 800,000/80,000 kg. Thrust 7,450.00 kN. Vacuum specific impulse 452 seconds. Original version of Energia core as used on Vulkan booster, with in-line upper stages and payloads. Developed 1974-1976; cancelled when Energia / Buran development begun. More...
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