Propulsion System for Manned Lunar Vehicle

Propulsion engineers of Yuzhnoye design office carried out an important and complex task: they developed an 11D40 propulsion system for the lunar landing vehicle.

The 11D410 propulsion system consisted of an RD-858 main engine and an RD-859 backup engine. The propulsion system would provide a soft landing on the surface of the Moon, liftoff from the Moon, and injection of the lunar vehicle into an elliptical orbit of the Moon’s artificial satellite.

The lunar vehicle would fly with a crew onboard; therefore, the most stringent requirements were placed upon the engine reliability. The reliability had to be proven by a great number of tests that simulated full-scale conditions. To provide a soft landing on the Moon and liftoff, the RD-858 engine features a dual-burn capability and two thrust levels: a main mode and a deep-throttling mode. A throttling range is ±9.8% in the main mode and ±35% in the deep-throttling mode. Such deep throttling required specific structural modifications to provide the engine chamber stability with reliable cooling.

The two-chamber RD-859 backup engine features one thrust level with the throttling range of ±9.8%.

The most stringent requirements were applied to the engine turbopump assemblies, and specifically to the face seals separating the oxidizer pump from the turbine. A significant number of tests were required to select the most reliable and efficient friction pair. The structure proved to be robust: the turbopump assemblies had a life estimated at thousands of seconds.

To provide reliable cooling, the chamber’s high-temperature flux area features machined helical flutes with optimal variable cross-section on complex-geometry parts.

The number of ignitions per engine reached twelve instead of two required in flight. The backup engine features a unique capability of ignition after a three-second period between cutoff and reignition. Processes of the engine cutoff, chamber pipeline emptying, and reignition after the three-second pause were thoroughly studied to prove the behavior convergence. The reignition parameters during tests were the same as those of the first ignition. None of the existing engines with a turbopump feed system was able to provide such performance. For liquid-propellant engines with turbopump feed systems providing a wide range of throttling, these engines featured quite a high specific impulse for such thrust level. The propulsion system mass and dimensions go to prove high design efficiency, even taking into account the integrated engine performance control and throttling systems. A total mass of the engines is 110 kg for a total thrust of 4100 kgf. For comparison, the mass of the Ariane-5 upper-stage engine exceeds 100 kg at 2700 kgf.

The test campaign was extensive: 181 RD-858 engines with a total running time of 253281 seconds and 181 RD-859 engines with a total running time of 209463 seconds. Eleven 11D410 propulsion systems were tested, with emergency simulated.

On the whole, the liquid-propellant propulsion system of the lunar landing vehicle is among the most reliable in its class. Three propulsion systems were successfully tested in Earth orbit onboard the T-2K spacecraft launched by the R-7 launch vehicle. 

Main engines

Name

Vacuum thrust, kgf

Propellants

Vacuum specific impulse, kgf?s/kg

Mass, kg

Missile/Launch Vehicle

RD853

47680

Oxidizer: nitric acid + 27% N2O4

Fuel: UDMH

300,7

485

8K66 (SS-7) missile second stage

RD854

7700

Oxidizer: NTO

Fuel: UDMH

312,2

100

8K69 (SS-9-2) boost stage: deceleration and control of the orbital spacecraft in all stabilization axes

RD857

14000

Oxidizer: NTO

Fuel: UDMH

329,5

190

8K99 (SS-15) missile second stage

RD861

8026

Oxidizer: NTO

Fuel: UDMH

317

123

11K68 (Cyclone-3) launch vehicle third stage: thrusting and control in powered flight in all stabilization axes

RD862

14544

Oxidizer: NTO

Fuel: UDMH

331

192

15А15 and 15А16 (SS-17-1 and SS-17-2) missile second stages

RD864

2060

Oxidizer: NTO

Fuel: UDMH

309

199

15А18 (SS-18-2) missile: two thrust modes and control in all stabilization axes during the post-boost vehicle flight

RD866

513,5

Oxidizer: NTO

Fuel: UDMH

323,1

125,4

Space tug engine; installed in 15Zh44, 15Zh52, 15Zh61, 15Zh60  missile nose cones

RD868

2371

Oxidizer: NTO

Fuel: UDMH

325

125

Zenit and Cyclone-4 apogee stages

RD869

2087

Oxidizer: nitric acid +

Fuel: UDMH

313

196

Space tug engine; 15А18М (SS-18-3) missile third stage flight control in all stabilization axes

 

History of Liquid-Propellant Rocket Engines

In 1958, development of steering engines for the first and second stages of the 8K64 ICBM became Yuzhnoye’s first experience of independent development of liquid-propellant rocket engines (LPRE). The main feature of this missile was a new fuel, unsymmetrical dimethylhydrazine (UDMH), used for the first time in combination with the AK-27 oxidizer. UDMH became the main fuel for several LPRE generations.

In 1960, successful development of the first steering LPRE allowed starting the development of a new, more complex and multifunctional RD-853 engine for the 8K66 missile second stage.

In 1961, Yuzhnoye started the development of steering engines for the 8K67 missile first and second stages. The engines used a new propellant combination: nitrogen tetroxide (NTO) and UDMH.

In 1962, design and tests started on an open-cycle RD-854 engine burning NTO and UDMH for the 8K69 ICBM orbital weapon unit deorbit propulsion system. For the first time in Soviet propulsion engineering, a pipe nozzle for the engine chamber was developed and introduced into production.

In 1964, Yuzhnoye started developing an RD-857 main engine for the 8K99 missile second stage: for the first time ever, the engine featured afterburning of a fuel-rich generator gas in the combustion chamber. The RD-857 was also the first engine with thrust vector control by generator-gas injection into the supersonic section of the nozzle.

Yuzhnoye also took part in the Soviet lunar program. In 1965, they started developing a propulsion system (Block E) for the lunar landing vehicle of the 11А52 lunar booster. Developed in Yuzhnoye, the lunar vehicle propulsion system consisted of an RD-858 main engine and an RD-859 backup engine. The propulsion system would provide a soft landing on the surface of the Moon, liftoff from the Moon, and injection of the lunar vehicle into the elliptical orbit of the Moon’s artificial satellite. On the whole, the lunar landing vehicle LPRE was one of the most reliable in its class. Three propulsion systems were successfully tested in Earth orbit onboard a special-purpose T-2K spacecraft launched by the Soyuz launch vehicle.

Development of the RD-861 engine for the Cyclone-3 launch vehicle third stage started in 1966. This engine features quite high mass and energy characteristics.

In 1976, during development of the 15А18 ICBM, Yuzhnoye started developing the RD-864, four-chamber open-cycle engine burning NTO and UDMH. The engine featured two thrust modes, main and throttling, with a multiple mode selection capability (up to 25 times). This engine was the first to use high-precise and high-speed regulator assemblies using high-pressure counter-jets.

The RD-869 engine of the 15А18М ICBM was derived from the RD-864 engine. The RD-869 had even better performance characteristics.

Development of the Zenit-2 launch vehicle in 1977 was the new milestone for Yuzhnoye. This launch vehicle burns cryogenic propellants: kerosene and liquid oxygen. For the first time in propulsion engineering, a steering engine using these propellants was designed as a staged combustion engine. The experience gained in the LPRE design and the introduction of advanced engineering solutions helped designing the RD-8 steering engine with good mass and energy characteristics, high reliability, and a long service life.

Steering engines

Name

Thrust, kgf

Propellants

Vacuum specific impulse, kgf?s/kg

Mass, kg

Missile/Launch Vehicle

RD851

28 850 (sea-level)

Oxidizer: Nitric acid + 27% N2O4

Fuel: UDMH

279

403

8K64 (SS-7) missile first stage control in all stabilization axes 

RD852

4 920 (vacuum)

Oxidizer: nitric acid + 27% N2O4

Fuel: UDMH

255

133

8K64 (SS-7) missile second stage control in all stabilization axes

RD855

29 100 (sea-level)

Oxidizer: NTO

Fuel: UDMH

292

320

8K67 (SS-9-1; SS-9-2) missile and Cyclone launch vehicle first stage control in all stabilization axes

RD856

5 530 (vacuum)

Oxidizer: NTO

Fuel: UDMH

280,5

112,5

8K67 (SS-9-1; SS-9-2) missile and Cyclone launch vehicle second stage control in all stabilization axes

RD863

28 230 (sea-level)

Oxidizer: NTO

Fuel: UDMH

301

310

15А15 and 15А16 (SS-17-1 and SS-17-2) missile first stage flight control

RD8

8 000 (vacuum)

Oxidizer: Liquid oxygen

Fuel: Kerosene

342

380

Zenit launch vehicle second stage flight control in all stabilization axes

 
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