Ukrainian Rocket Engine Technologies

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LIQUID-PROPELLANT ROCKET ENGINES​

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

 

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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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HISTORY OF SOLID ROCKET MOTORS​

The first solid rocket motor (SRM) designed by Yuzhnoye was a 15D15 main motor for the first stage of the 8K99 missile. The motor development started in 1963.

The first firing of the 15D15 motor was conducted in April 1965.

However, in October 1969, development of the 8K99 missile was stopped regardless of a series of completely successful launches. The experience in SRM development allowed working on new advanced approaches to the definition of the optimum configuration of future main solid rocket motors.

In 1969, Yuzhnoye started the development of the 15Zh43 ICBM and its first-stage motor, the 15D122.

A NUMBER OF BREAKTHROUGH TECHNOLOGIES WERE PROPOSED FOR THE MOTOR DESIGN:​

• Combined casing with a fiber-glass longitudinal-circumferential wound pipe and metal motor heads
• Case-bonded block of solid butyl-rubber propellant
• Central fixed nozzle partly submerged into the combustion chamber, with thrust vector control provided by injecting hot chamber gas into the submerged diverging section of the nozzle.

The 15D122 motor firing tests demonstrated the motor performance and characteristics of the thrust vector control system based on the hot-injection technique.

In this period, Yuzhnoye developed steering SRM for deployment of objects in space. The 15D161, 15D171, and 15D221 motors featured increased running time and control forces. The motors used end-burning charges made of metal-free propellant mixture with an original design of case bonding and a unique configuration of the seal in bearings of the rotating nozzles.

The next work Yuzhnoye performed in the SRM design was the development of the 3D65 main motor for the first stage of the 3M65 naval missile (designed by Makeyev design bureau). The motor design incorporated the most advanced engineering solutions:

• All-wound cocoon casing with a load-bearing shell made of high-strength organic fiber and embedded parts made of titanium alloy
• Case-bonded charge of high-energy butyl-rubber propellant
• Fixed nozzle with a three-axis thrust vector control system based on the hot-injection technique
• A number of engineering solutions determined by the specifics of using the motor in
• the naval missile (for both above-water and underwater launches).

In 1982, the 3D65 motor was accepted for series production.

In the mid-1970s, Yuzhnoye started developing main solid rocket motors for a silo-based 15Zh44 ICBM and a rail-mobile 15Zh52 ICBM: the 15D206 motor for the first stage and the 15D207 motor for the second stage.

To reduce the scope and period of development tests, the 15D206 motor was designed as a complete analog of the 3D65 motor. The modifications included increased thrust/consumption performance, increased throat diameter, and increased chamber pressure.

THE FOLLOWING NEW TECHNOLOGIES WERE USED IN THE 15D207 MOTOR:​

• Fixed nozzle with an extendable high-altitude exit cone
• Carbon-carbon composite materials in the throat liner
• Formulation of propellants with higher energy characteristics
• Block of solid propellant with high chamber volumetric loading efficiency.

Ground tests of the motors started in 1979.

However, in 1983, it was decided to stop the development of the 15Zh44 and 15Zh52 missiles and use them as the basis for the 15Zh60 and 15Zh61 missiles with higher performance characteristics and better nuclear hardness.

A new motor, 15D290, was designed for the second stage of rail-mobile 15Zh61 missiles. Better performance of the motor was attained by using a new high-energy propellant mixture and making a number of design decisions to increase the motor’s nuclear hardness.

In the 15Zh60 missile, the first- and second-stage main motor requirements necessitated the development of a fundamentally new motor, 15D305, for the first stage and the upgrade of the 15D339 motor for the second stage.

THE 15D305 MOTOR DESIGN WAS BASED ON THE FOLLOWING UNIQUE TECHNOLOGIES:​

• High-energy propellant based on cyclotetramethylene tetranitramine (octogene)
• Cocoon casing
• Central swiveling nozzle with a one-piece throat liner made of 3D reinforced carbon-carbon material.

For the 15D339 motor, a multifunctional coating to protect the casing against all nuclear effects was developed; mass efficiency of the motor and erosion resistance of the nozzle were improved.

In 1986-1988, development tests of the 15D290, 15D305, and 15D339 motors were completed, and series production started.

In 1988, Yuzhnoye design office was tasked with developing the 15Zh65 ICBM first-stage propulsion system, the 15D365.

THE 15D365 MOTOR DESIGN FEATURES INCLUDE:​

• Block of propellant with a swiveling control nozzle, following a circular motion pattern
• Organic plastic cocoon casing
• Case-bonded octogene-based propellant mixture.

Five firing tests were conducted, and the 15D365 motor was accepted for flight tests. However, due to collapse of the Soviet Union, all work on the 15D365 in Yuzhnoye was shut down.

Besides main and steering motors, Yuzhnoye developed 82 types of small motors, pressure accumulators, and gas generators, which were used for a broad range of applications:

• Missile ‘mortar’ launch
• Missile ‘mortar’ staging
• Missile ‘wag’ during a mortar launch
• Nose-cone inflatable tip geometry change
• Increase of a main motor nozzle altitude capability
• Separation and removal of various objects away from the missile trajectory
• Missile parts flight control
• Ejection of objects from the missile and support a specified velocity of their flight
• Spin stabilization.

Many years of successful operation have proven high reliability and high operating environment resistance of the developed SRM.

BASIC SPECIFICATIONS OF MAIN MOTORS​

Name	Vacuum thrust, kgf	Propellant	Vacuum specific impulse, kgf?s/kg	SRM loaded mass, kg	Missile/Launch Vehicle
15D15	70	PEKA-181D mixture	265,0	20,0	8K99 (SS-X-15) missile first-stage steering motor
3D65	205,8	Т-9BK-8E mixture	274	52,65	3М-65 naval missile second-stage steering motor
15D206	235,7	Т-9BK-8E mixture	271,2	52,45	15Zh43 and 15Zh52 (SS-24 Mod 1) missile first-stage steering motor
15D207	142,8	OPAL mixture	291,2	32,2	15Zh43 and 15Zh52 (SS-24 Mod 1) missile second-stage steering motor
15D290	145,7	START mixture	297,5	32	15Zh61 (SS-24 Mod 3) missile second-stage steering motor
15D339	147,3	START mixture	296,6	32,05	15Zh60 (SS-24 Mod 2) missile second-stage steering motor
15D305	310,8	OPAL mixture	280	51,57	15Zh60 (SS-24 Mod 2) missile first-stage steering motor
15D365	145	OPAL mixture	283	27,55	15Zh65 Universal (SS-X-27) missile first-stage steering motor

BASIC SPECIFICATIONS OF POST-BOOST MOTORS​

Name	Vacuum thrust, kgf	Propellant	Vacuum specific impulse, kgf?s/kg	SRM loaded mass, kg	Missile
15D171	260	T-9BKN-9K cryogenic mixture	232,0	250	15А15, 15А16 missiles
15D161	780	229,0	940	15А14 missile
15D221	910	234,5	1343

BASIC SPECIFICATIONS OF SMALL SRM​

Propellant	Composite propellant with a burning temperature of 1200 to 3300°K and aluminum content of 3 to 18% Ballistite (bibasic) propellant with a burning temperature of 1600 to 2800°K
Charge mass, kg	0.006 to 120
Bonding type	Mixture: inserted and case-bonded Ballistite propellant: inserted and glued into a heat-shielding barrel
Motor case material	High-strength steel, aluminum alloy, titanium alloy
Configuration	Classic design: straight or deflected nozzle Skewed nozzle Single-nozzle, two-nozzle, four-nozzle configuration with a plume spreader (for cartridge-pressure accumulator) Segner wheel shaped Ignition on nozzle side Ignition from electrical and mechanical pyrocartridges
Total thrust impulse, kgf?s	5 to 6 000
Thrust, kgf	7 to 16 680
Running time, s	0.1 to 60
Thrust (consumption) variation diagram shape	Rectangular Degressive, stepped Progressive ( 3-20 times) Thrust reversal with frequency of 15 Hz

 

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RD-843​

The RD843 engine generates thrust and provides two-axis flight control of launch vehicle upper stages and booster stages.

The RD843 is a single-chamber engine with a pressure-fed supply system of hypergolic propellants.

The engine gimbaling in two mutually perpendicular planes provides thrust vector control. The engine is capable of multiple (up to 5) burns in flight.

The engine starts and cuts off by the control system, commanding oxidizer and fuel electrohydraulic valves.

The RD843 ground test campaign included 74 tests, 140 ignitions, reaching a total of 8201 s, which is approximately 12 service lives on 4 engines. In 2012-2015, the engine successfully passed 6 flight tests onboard lightweight launch vehicles.

BASIC SPECIFICATIONS​

Engine mass, kg	16,5
Propellants: Oxidizer Fuel	Nitrogen tetroxide Unsymmetrical dimethylhydrazine
Vacuum thrust, kgf	250
Vacuum specific impulse, s	315,5
Propellant mixture ratio	2,0
Gimbal angle, ang. deg	±10
Running time in flight, s	Up to 700

 

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RD-861K​

The RD861K engine generates thrust and provides pitch and yaw control of the Cyclone-4 third stage powered flight. The engine is based on a highly reliable RD861 serial engine used in the Cyclone-3's third stage.

The open-cycle RD861K is a single-chamber engine with a turbo-pump-feed system using hypergolic propellants. The engine gimbaling in two mutually perpendicular planes provides thrust vector control.

The RD861K is capable of multiple (up to 5) burns in flight. The turbopump assembly turbine uses oxygen-rich gas as a working fluid.

Ground development tests of the engine are nearing completion.

BASIC SPECIFICATIONS​

Engine mass, kg	207
Propellants: Oxidizer Fuel	Nitrogen Tetroxide Unsymmetrical Dimethylhydrazine
Vacuum thrust, kgf	7916
Vacuum specific impulse, kgf·s/kg	330
Propellant mixture ratio	2,41
Gimbal angle, ang. deg	±5
Running time in flight, s	Up to 480

 

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RD-840​

Apogee main engine is single-chamber, single-mode, multiple firing, with propellant-pressurized feed system.

It is designed for SC add-on injection into orbits, including geostationary orbit.

​

BASIC SPECIFICATIONS​

Engine mass, kg	4.3
Propellant components:	АТ+НДМГ
Vacuum thrust, kgf	40,7
Vacuum specific impulse, s	315
Propellant mixture ratio	2,05
Number of firings	100
TRL	6

 

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RD-860, RD-860L, DU-802​

RD860 and RD860L engines with pneumatic pump propellant feed system are designed for LV upper stages, space tugs and unmanned takeoff and landing modules.

RD860L engine consists of two blocks (EB-1 – dual-chamber, stationary, EB-2 – RD860 engine with gimballing).

The engines are developed on the basis of PS802 proven technologies, and provide for multiple firing.

BASIC SPECIFICATIONS​

Engine model	РД860	РД860L
Mass, kg	4.3	106
Propellant components:	АТ+НДМГ
Vacuum thrust, tf (main/throttle)	0,5/0,25	1/0,25
Vacuum specific impulse, s	322,5	316
Propellant mixture ratio	2,32
Number of firings	2
TRL	6

PS802 propulsion system with the use of pneumatic pump propellant feed system. The propulsion system is developed for Krechet AST, provides additional injection of different payloads including into lunar trajectory.

It has high mass energy characteristics, has a capability of multiple firing during flight and duration of in-flight functioning for 10 days and more.

BASIC SPECIFICATIONS​

Mass, kg	165,4
Propellant components:	АТ+НДМГ
Vacuum thrust, tf	0,45
Vacuum specific impulse, s	322,5
Propellant mixture ratio	2,25
Number of firings	10
TRL	7

 

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RD-805​

The main engine is single-chamber, single-mode, single firing, with turbopump propellant feed system, made according to oxidizing generator gas afterburning scheme.

The engine is developed on the basis of proven units of Zenit LV RD-8 high-reliable series-production steering engine.

ОСНОВНЫЕ ТЕХНИЧЕСКИЕ ХАРАКТЕРИСТИКИ​

Engine mass, kg	80
Propellant components:	О2+ kerosene
Vacuum thrust, tf	2,15
Vacuum specific impulse, s	345
Propellant mixture ratio	2,5
Number of firings	1
TRL	4

 

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RD-809K​

The RD809K engine generates thrust and controls flight of the upper stages of the Mayak family rockets.

The engine is based on the Cyclone-4 rocket third-stage RD861K sustainer chamber and the Zenit rocket second-stage RD-8 steering engine assemblies. The staged-combustion RD809K is a single-chamber, single-mode engine with a turbo-pump-feed system

The engine gimbaling in two planes provides stage flight control. The engine is capable of multiple (up to 4) burns in flight.

The turbopump assembly turbine uses oxygen-rich gas as a working fluid. A starting fuel is used to ignite propellants in the chamber and the gas generator. The engine is currently in development.

BASIC SPECIFICATIONS​

Engine mass, kg	330*
Propellants: Oxidizer Fuel	Liquid oxygen Kerosene
Vacuum thrust, kgf	10000
Vacuum specific impulse, kgf·s/kg	352
Propellant mixture ratio	2,62
Engine gimbal angle, ang.deg	±5
Total running time in flight, s	Up to 600

 

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RD-835​

RD835 main engine with turbopump propellant feed system, made according to generator gas afterburning scheme, and in-flight two-times firing. The engine has high mass-energy characteristics. This engine is designed for the use in most advanced LV of Mayak type, developed by Yuzhnoye SDO. The engine is developed on the basis of Zenit LV proven technologies. RD836 engine can be developed on its basis for LV first stage.

BASIC SPECIFICATIONS​

Engine mass, kg	830
Propellant components	О2+kerosene
Vacuum thrust, tf	50
Vacuum specific impulse, s	355
Propellant mixture ratio	2,6
Number of firings	2
TRL	3

 

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RD-870​

RD870 engine block (EB) with turbopump propellant feed system is made according to generator gas afterburning scheme. The EB is developed on the basis of Zenit LV proven technologies. RD872 engine can be developed on its basis for LV second stage with thrust ~90 tf.

BASIC SPECIFICATIONS​

Engine mass, kg	1353
Propellant components:	О2+ kerosene
Ground thrust, tf	79,3
Vacuum thrust, tf	89,4
Ground specific impulse, s	301,5
Vacuum specific impulse, s	340,0
Propellant mixture ratio	2,684
Number of firings	1
TRL	6

 

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RD-801​

The RD801 engine generates thrust and controls flight of the Mayak launch vehicle first stage. The engine is based on technologies tested on the Zenit launch vehicle.

The RD801 is a single-chamber, dual-mode, single-burn engine with a turbo-pump-feed system. The engine uses oxygen-rich staged combustion.

The engine gimbaling in two mutually perpendicular planes provides thrust vector control.

A starting fuel is used to ignite propellants in the chamber and the gas generator.

The turbopump assembly turbine uses producer gas as a working fluid.

A 492-ton-thrust engine cluster, comprised of four RD801 engines can be developed based on the RD801.

The engine is currently in development.

BASIC SPECIFICATIONS​

Engine mass, kg	1630
Propellants: Oxidizer Fuel	Liquid oxygen Kerosene
Thrust, kgf: Sea-level Vacuum	122224 136566
Vacuum specific impulse, kgf·s/kg: Sea-level Vacuum	300,7 336
Propellant mixture ratio	2,65
Engine gimbal angle, ang.deg	±6
Running time in flight, s	Up to 200

 

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RD-810​

The RD810 generates thrust and controls flight of launch vehicle first stages. The staged-combustion engine is based on proven Zenit launch vehicle technologies. The RD810 is a single-chamber, dual-mode single-burn engine with a turbo-pump-feed system.

The engine gimbaling in two mutually perpendicular planes provides thrust vector control. A starting fuel is used to ignite propellants in the chamber and the gas generator.

The turbopump assembly turbine uses oxygen-rich gas as a working fluid. The RD810 engine can be used for development of an 800-ton-thrust engine cluster incorporating four RD810 engines. The engine is currently in development.

BASIC SPECIFICATIONS​

Engine mass, kg, not more than	2800
Propellants: Oxidizer Fuel	Liquid oxygen Kerosene
Thrust, kgf: Sea-level Vacuum	191314 214639
Vacuum specific impulse, kgf·s/kg: Sea-level Vacuum	299 335,5
Propellant mixture ratio	2,65
Gimbal angle, ang.deg	±8
Running time in flight, s	Up to 300

 

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RD-815​

Main engine with turbopump propellant feed system, made according to generator gas afterburning scheme. The engine is developed on the basis of Zenit LV proven technologies. The engine has high mass-energy characteristics.

This engine is designed for the use in most advanced LV of Mayak type, developed by Yuzhnoye SDO. Cluster engine line with thrust 500 to 1000 tf can be developed on the basis of this engine.

BASIC SPECIFICATIONS​

Engine mass, kg	3200
Propellant components:	О2+ kerosene
Ground thrust, tf	251,3
Vacuum thrust, tf	274,5
Ground specific impulse, s	306,7
Vacuum specific impulse, s	335
Propellant mixture ratio	2,7
Number of firings	1
TRL	4

 

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RD-880​

The engine is dual-chamber, two-mode, single firing, with turbopump propellant feed system, made according to oxidizing generator gas afterburning scheme.

The engine is developed on the basis of Zenit LV proven technologies.

RD880.1 single-chamber engine with thrust 500 tf can be developed on the basis of РД880.2 engine.

BASIC SPECIFICATIONS​

Engine mass, kg	6179
Propellant components:	О2+kerosene
Ground thrust, tf	502,6
Vacuum thrust, tf	549
Ground specific impulse, s	306,7
Vacuum specific impulse, s	335
Propellant mixture ratio	2,7
Number of firings	1
TRL	3

 

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LIQUID-PROPELLANT THRUSTERS LOW-THRUST​

Liquid-propellant thrusters are incorporated in liquid-propellant jet propulsion systems and generate thrust by propellant exhaust through a supersonic nozzle.



Cyclone-3 Rocket 100 N Thruster Cyclone-3 Rocket 30 N Thruster



Okean-O Spacecraft 30 N Thruster Cyclone-4 Rocket 30 N Thruster

BASIC SPECIFICATIONS​

Parameter	Cyclone-3 thruster	Cyclone-3 thruster	Okean-O thruster	Cyclone-4 thruster
Propellants: Oxidizer fuel	Nitrogen Tetroxide Unsymmetrical Dimethylhydrazine
Nominal thrust, N	100	30	30	30
Specific impulse in continuous running at nominal propellant inlet pressure, m/s	2600	2000
Chamber pressure, MPa	0,7	0,1
Nominal propellant inlet pressure, MPa	1,35	Oxidizer - 0,53
Fuel - 0,25
Nominal mixture ratio	1,6	1,3	1,6	1,85
Guaranteed number of ignitions	1300	50000	30000
Total running time, s	200	3600	5600
Duration of a single ignition, s	0,21 / 165	0,21 / 600	0,21 / 2000
Supply voltage, V	28
Mass, kg	1,45	1,3	1,16

 

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SOLID-PROPELLANT MAIN ENGINES​

MAIN SOLID ROCKET MOTORS

The main functions are:

• thrust application;
• producing control force along the yaw and pitch channels with rotating control nozzle (RCN);
• performance of all SRM characteristics under the specified external influences in every phase of operation.

These SRM are used as the first and second stages of the space launch vehicle (SLV).

Basic characteristics:

Propellant	composite, high-energy
External diameter, mm	up to 2000
Thrust, ton-force	up to 150
Casing	carbon fiber-reinforced plastic, cocoon type
Nozzle	rotating control

MAIN SRM FOR SHORT-RANGE MISSILE

The main functions are:

• thrust application;
• performance of all main motor characteristics under the specified external influences in every phase of operation.

The short-range missiles are intended for killing enemy targets within the operational penetration limit relative to the front line.

Basic characteristics:

Propellant	composite, high-energy
External diameter, mm	up to 1000
Thrust, ton-force	up to 20
Casing	carbon fiber-reinforced plastic, cocoon type
Nozzle	fixed

MAIN SRM FOR MEDIUM-RANGE AND SHORT-RANGE SURFACE-TO-AIR MISSILES

The main functions are:

• thrust application;
• performance of all main motor characteristics under the specified external influences in every phase of operation.

The surface-to-air missiles are parts of the air defense missile complex and intended for killing various air targets.

Basic characteristics:

Propellant	composite, high-energy
External diameter, mm	up to 450 up to 300
Casing	high-strength steel
Nozzle	fixed

SRM FOR MULTIPLE LAUNCH ROCKET SYSTEM

Multiple launch rocket systems are one of the widespread and promising types of the ground tactical missile-artillery armament; they are capable of engaging various targets and executing various tasks depending on the head module.

Basic characteristics:

Propellant	composite, high-energy
External diameter, mm	from 122 to 400
Casing	high-strength steel
Nozzle	fixed

MAIN SRM FOR REBOOST STAGE

It is intended for increasing payload mass injected into orbit by SLV.

Basic characteristics:

Propellant	composite, high-energy
External diameter, mm	up to 1250
Thrust, ton-force	up to 10
Casing	carbon fiber-reinforced plastic, cocoon type
Nozzle	rotating control

 

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SOLID-PROPELENT SPECIAL ENGINES​

PULSED CONTROL SRM
Usage of micro SRM provides short-time control pulses along the yaw and pitch channels

Basic characteristics:

Propellant	composite, high-energy
Operating time, s	0.045 … 0.05
Thrust, kgf	up to 500
Case	high-strength steel

LAUNCH-BOOST SRM

It provides high launch thrust due to usage of propellant with high specific momentum and increased combustion rate.

Basic characteristics:

Propellant	composite, high-energy
Operating time, s	5
Thrust, kgf	up to 8000

ROLL CONTROL SRM

It provides control force along the roll channel due to distribution valve and intrachamber pressure regulator.

Basic characteristics:

Propellant	composite, low-temperature
Operating time, s	75...90
Thrust, kgf	up to 30

SRM FOR METEOROLOGICAL ROCKET SOFT LANDING

Usage of sloping nozzles and multigrain charge with inhibitor eliminates the thermal effect on the recoverable payload.

Basic characteristics:

Propellant	composite
Operating time, s	0,27
Thrust, kgf	up to 3500

STAGE SLOWDOWN SRM

Usage of sloping nozzles eliminates the influence of combustion products blast on the rocket structure.

Basic characteristics:

Propellant	composite
Operating time, s	0,3...04
Thrust, kgf	up to 500
Nozzle	fixed, skewed

SRM FOR SPACECRAFT DEPLOYMENT

It is used to perform advanced spatial maneuvers during spacecraft deployment.

Basic characteristics:

Propellant	composite, metal-free
Operating time, s	120...270
Thrust, kgf	up to 300-1200
Угол вращения сопел	±60 deg

LAUNCH SOLID-FUEL GAS GENERATOR

It provides the expulsion of the missile from the transportation-and-launch container during the mortar launch.

Basic characteristics:

Propellant	composite
Operating time, s	2,6...3,5
Thrust, kgf	up to 300

 

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HALL-ENGINE-BASED ELECTRIC JET PROPULSION SYSTEM​

The Hall-engine-based Electric Jet Propulsion System (EJPS) is designed for spacecraft orbit correction, stabilization, and interplanetary flights.

BASIC SPECIFICATIONS​

Working medium - xenon
Maximum bottle pressure at 50°С, MPa	13
Total thrust impulse, kN·s	50 - 1000
Nominal Hall engine thrust, mN	1 - 66
Thrust excursion in all operating environments, %	15
Specific impulse of Hall engines in continuous running mode in nominal environment, m/s	2200 -22000
Engine thrust-to-power ratio, W/mN	16-20

 

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AMMONIA JET PROPULSION SYSTEMS​

The ammonia propulsion system (APS) is designed for correction of SC injection error in orbital period and orbital period keeping during active lifetime.

AMMONIA PROPULSION SYSTEM FOR SICH-2M SC​

BASIC SPECIFICATIONS​

Total thrust pulse, N·s (kgf·s):
in hot mode– in cold mode –	39240 (4000) 13780 (1405)
One engine thrust in hot and cold modes, N (gf)	0,049 (5)
Thrust deviation from nominal value, %:
in hot mode– in cold mode –	± 15 ± 22
Number of APS firings during ALT	up to 1000
Number of engines, pcs.	2
Duration of one firing, min	20, no more
Specific impulse, m/s (s):
in hot mode– in cold mode –	2500 (255) 882 (90)
Direct-current power voltage, V	от 24 до 34
Maximum power consumption of APS at nominal power voltage of 28 V, W	472
APS mass (dry), kg	40

AMMONIA PROPULSION SYSTEM FOR MICROSAT SC​

BASIC SPECIFICATIONS​

Total thrust pulse, N·s (kgf·s):
in hot mode– in cold mode –	7000 (714) 2500(255)
One engine thrust in hot and cold modes, N (gf)	0,049(5)
Thrust deviation from nominal value, %:
in hot mode– in cold mode –	±15 ±22
Number of APS firings during ALT	up to 500
Number of engines, pcs	2
Длительность одного включения двигателя, мин	15, no more
Duration of one firing, min
in hot mode– in cold mode –	2500 (255) 882 (90)
Direct-current power voltage, V	от 24 до 34
Maximum power consumption of APS at nominal power voltage of 28 V, W	375
APS mass (dry), kg	24