Rocket Engine Specifications
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NPO Energomash Rocket Engines

Credits for providing information to Jim Glass
Boeing North American, and Jane's Space

NPO Energomash

The Gas Dynamics Lab. OKB GDL was created in 1928 in Leningrad (now St. Petersburg) for R&D and production of liquid propellant and electric rocket engines.

Note RD-100 through RD-120 will not be listed.


Application to be identified
Dry Mass 141/119 kg
Length 2205/1700 mm
Maximum Diameter 1020/780 mm
Engine Cycle closed gas generator
Oxidizer LOX
Fuel Kerosene
Mixture Ratio (O/F) 2:6
Feed Method turbopump
Thrust 19.9/19.6 kN vacuum
Isp 365/360 sec vacuum
Expansion Ratio 19.25/18.75
Chamber Pressure 116.1 atm
Burn Time 900 sec maximum


Application SL-16 stage, SL-17 Energia strap-ons
First Flown 13 April 1985 (SL-16 sub-orbital test)
Dry Mass 8755 kg
Length 3.56 m
Maximum Diameter 399 cm
Engine Cycle staged combustion
Oxidizer LOX at 432 kg/s
Fuel Kerosene at 166.2 kg/s
Thrust 7903.77 kN vacuum and 7256.56 kN sea level (but can be throttled back to 4413 kN vac)
Isp 308 sec sea level and 336 sec vacuum
Expansion Ratio 36.8
Chamber Pressure 242 atm; 250 atm maximum
Chamber Mass 480 kg
Burn Time 140-150 sec


NPO Energomash proposes development of an RD-170 2-chamber version for the U.S. market.

RD-180 is proposed for use on the Atlas around the year 2000.


Application SL-7 Cosmos satellite launcher stage 1, SS-4 Sandal IRBM stage 1
First Orbital Flight March 1962
Number Flown 144 on orbital missions
Dry Mass 645 kg
Engine Cycle Open
Oxidizer Nitric Acid
Fuel Kerosene
Feed Method single turbopump driven by H2O4 gas generator feeding four fixed chambers
Thrust 635 kN sea level, 730 kN vacuum
Isp 230 sec sea level, 264 sec vacuum
Chamber Pressure 45 atm
Chamber Internal Diameter 480 mm, throat diameter 176 mm
Burn Time 140 sec


Two-chamber high altitude engine developed during 1958-61 for the SS-9 Scarp ICBM second stage.

Application SS-9 Scarp stage 2; Tsyklon stage 2?
First Orbital Flight June 1977?
Dry Mass 665 kg
Engine Cycle Open
Oxidizer Nitric Acid
Feed Method single turbopump feeding two gimballed chambers
Thrust 883 kN
Isp 293 sec vacuum
Chamber Pressure 75 atm
Burn Time 125 sec

RD-251 & RD-252

NPO Yuzhnoye's SS-18 ICBM is powered by four 1128 kN chambers (RD-251) on stage 1, and the 755 kN RD-252 (no indication of number of chambers) on stage 2. Both stages employ NTO/UDMH.


The first known Soviet closed cycle and N2O4 and UDMH engine, the gimballed single chamber RD-253 provides the first stage power for Proton's six clustered external tanks.

Appliation Proton stage 1
First Flown 16 July, 1965
Dry Mass 1280 kg
Length 2.72 m
Maximum Diameter 1.50 m
Engine Cycle closed
Oxidizer N2O4
Propellant Flow Rate about 528 kg/s; mixture ratio 2.67
Feed Method 18.7 MW turbopump driven by pre-burner gas
Thrust 1745 kN vacuum (1474 kN sea level), uprated from 1635 kN (1460 kN sea level).
Profile 40% full thrust in 0.1 sec, held at 40% for 2 sec, then to 100% in 0.1 sec
Isp 317 sec vacuum, 285 sec sea level
Chamber Pressure 150 atm
Burn Time 130 sec


Although apparently not tested beyond component level, the RD-270 was designed Chelomei's UR-700 lunar vehicle.

Appliation launcher stage 1
Dry Mass 4470 kg (5603 kg wet)
Length 4.85 m
Maximum Diameter 3.3 m
Engine Cycle closed staged
Oxidizer N2O4
Propellant Flow Rate unknown
Feed Method turbopump
Thrust 6713 kN vacuum, 6272 kN sea level
Isp 322 sec vacuum, 301 sec sea level
Chamber Pressure 266 atm
Burn Time unknown


NPO Energomash has been working on this tri-propellant engine since 1989 to propel a small 22t shuttle-type vehicle designed by NPO Molniya for release from an An-225 at altitude.

Mass 3990 kg assembly, 1840 kg each dry
Length 5.4 m extension down, 3.8 m extension up
Diameter 2.4 m
Engine Cycle closed staged combustion
Oxidizer LOX
Fuel mode 1: hydrogen and kerosene; mode 2: hydrogen
Mixture Ratio (O/F; mass) mode 1: 5.27:1 kerosene and 13.2:1 hydrogen; mode 2: 6.1:1
Flow Rate (each chamber, kg/s) mode 1: 388.4 LOX, 29.5 LH2, and 73.7 kerosene; mode 2: 148.5 LOX and 24.7 LH2
Thrust (each chamber, vac) 1960 kN mode 1 and 785 kN mode 2 (throttle 40-100%)
Isp mode 1: 415 sec and 330 sea level; mode 2: 460 sec
Chamber Pressure mode 1: 290 atm; mode 2: 122 atm
Expansion Ratio 170 extension down and 70 extension up

KB Khimautomatiki

Originally the bureau of Semyin A Kosberg (1903-65), formed in 1941, then in 1974, it was re-organized under the above name.


Energia's four single chamber core engines are the first operational Soviet (now Russian) cryogenic systems.

Application SL-17 Energia core, 4 engines
First Flown 15 May, 1987
Number Flown 8 to end of 1993
Dry Mass ~ 3.5 t
Length 4550 mm
Maximum Diameter 2420 mm
Engine Cycle closed
Oxidizer LOX
Fuel LH2
Mixture Ratio (O/F) 6.0:1
Feed Method 35,000 rpm dual-stage turbopumps on each chamber. 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.
Thrust 1961 kN vacuum, 1451 kN seas level; throttle range is 45%
Expansion Ratio 85.7:1 (throat dia 261 mm, exit dia 2420 mm)
Isp 455 sec vacuum
Chamber Pressure 220 atm
Burn Time about 480 sec (600 sec max operational)


Three versions of this hypergolic engine are used on Proton. Stage 2 (RD-0208) clusters three RD-465 and one RD-468 (the latter differing only by carrying the gas generator). Stage 3 (RD-0212) employs the RD-473 single engine version with four verniers.

Application SL-12/13 Proton stages 2 & 3
First Flown July 1965 (Stage 2), 1967 (Stage 3)
Engine Cycle closed (oxidiser pre-burner gas routed to main chamber after driving turbine)
Oxidizer Nitrogen Tetroxide
Thrust 594 kN vacuum (Stage 2 version), 593.6 kN vacuum (Stage 3, plus 31.5 kN added by four verniers)
Isp 327.4 (Stage 2) & 325.3 (Stage 3)sec vacuum
Chamber Pressure 148 atm
Burn Time about 210 sec

Rocket Engine Specifications

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