NPO Energomash Rocket Engines
Credits for providing information to Jim Glass
Boeing North American, and Jane's Space
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 |
| Fuel | UDMH |
| Feed Method | single turbopump feeding two gimballed
chambers |
| Thrust | 883 kN |
| Isp | 293 sec vacuum |
| Chamber Pressure | 75 atm |
| Burn Time | 125 sec |
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 |
| Fuel | UDMH |
| 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 |
| Fuel | UDMH |
| 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 |
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 |
| Fuel | UDMH |
| 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 |
ASI W9800156r1.3.
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