Rocket engine rd. Rocket engines

Assembly shop of rocket engines of JSC Kuznetsov / Photo: TASS, Sergey Fadeichev

The first two of 60 RD-181 rocket engines under a contract worth $1 billion were supplied by the Russian side to the American company Orbital Sciences Corporation, RSC Energia told RIA Novosti.

“On July 16, the first two engines were delivered to the United States; the delivery date for the next batch has not yet been announced,” the press service said.

Earlier, TASS reported that the United States would need up to ten years and about $3 billion to create its own rocket engine, similar to the Russian RD-181. This was stated by the President of RSC Energia Vladimir Solntsev.

“If we talk about creating an engine in America, then any power is capable of creating it, the only question is: why? According to my estimates, this will require about $3 billion, and maybe more, and seven to ten years,” said V. Solntsev. He emphasized that the outcome may be ambiguous, and it is not a fact that everything will work out. At the same time, the head of the corporation noted, no one “canceled or killed international cooperation.”

Vladimir Solntsev / Photo: www.rg.ru

Earlier, V. Solntsev reported that NPO Energomash plans to supply the American company Orbital Sciences with 60 new RD-181s, a contract has already been signed for 20 of them. The contract amount is about $1 billion, and this amount includes a whole range of services. V. Solntsev participated in the transaction on behalf of NPO Energomash, since he remains the executive director of the association.

The head of Energia also explained that Energomash and the United Rocket and Space Corporation received all the necessary permits for the transaction.

The RD-181 will be used on the first stage of the Antares rocket, manufactured by Orbital Sciences. Previously, these missiles were equipped with AJ-26 engines, based on the Soviet NK-33.

Russia adapted RD-180 engines to American needs

Solntsev noted that Russian specialists adapted the RD-180 engines supplied to the United States for Atlas rockets for manned launches. "Over the last two years, we have completed a contract with the United Launch Alliance to adapt the RD-180 engine for the United States manned program. Both (launcher) Atlas and Space X won the tender, and we have fulfilled this contract and adapted this engine for manned programs. launches," he said.

“Atlas-5 can now acquire a new history, a new face,” noted V. Solntsev.

The US military appropriations bill for fiscal year 2015 includes an amendment from Republican Senator John McCain that would ban further purchases of the RD-180 for missiles. Atlas puts into orbit not only civilian but also military satellites, and many lawmakers argue that US national security interests have become dependent on Russian technology.

At the same time, the law provides for the allocation of $220 million for the development of an American engine to replace the RD-180. The US Air Force, which is responsible for space launches, has been asked to solve the problem within five years. The Pentagon, however, openly admits that they do not know how to use the allocated funds - it is advisable to create a new engine only for a specific launch vehicle.

This is currently being done by the United Launch Alliance corporation, which launches Atlas. The US military believes that it should continue this work.

Technical information

RD-180- closed-cycle liquid rocket engine with afterburning of oxidizing generator gas after the turbine, equipped with two combustion chambers and two nozzles. Developed and produced by OJSC NPO Energomash named after Academician V.P. Glushko. Fuel - kerosene, oxidizer - liquid oxygen. The cost of one engine as of 2010 is $9 million.

RD-180 was created in the mid-90s based onengine RD-170. In 1996, General Dynamics acquired the rights to use the engine. It was first used on May 24, 2000, as the first stage of the Atlas IIA-R launch vehicle, a modification of the Atlas IIA rocket. The letter "R" in the rocket's name indicates the use of a Russian-made engine. The rocket was later renamed Atlas III.

After the first launch, additional work was carried out to certify the engine for use on the universal rocket module of the core stage of the Atlas-5 rocket. The RD-180, which was used on the test bed, was shown at the 23rd G8 meeting (June 1997, Denver, USA). The engine consists of two chambers, a turbopump unit, a fuel booster pump unit, an oxidizer booster pump unit, a gas generator, an automation control unit, a cylinder unit, an automation drive system, a steering drive system, a fuel flow regulator in the gas generator, an oxidizer throttle, a fuel throttle, a starting oxidizer and fuel shut-off valves, two ampoules with starting fuel, a starting tank, an engine frame, a bottom screen, emergency protection system sensors, a heat exchanger for heating helium to pressurize the oxidizer tank.

Today, all engine production is concentrated in Russia. The sale was carried out by a joint venture between Pratt & Whitney and NPO Energomash, called RD AMROSS JV. The acquisition and installation were carried out by United Launch Alliance. Since May 2014, the conclusion of new contracts has been temporarily stopped by a court order due to a claim by a competitor, SpaceX; deliveries of engines under old contracts continue.

RD-180 engine on a test bench at the Marshall Space Center (USA) / Photo: / ru.wikipedia.org

Tactical and technical indicators

Type	liquid rocket engine
Fuel	kerosene
Oxidizer	liquid oxygen
Combustion chambers	2
Application	"Atlas III and Atlas V" (first stage)
Development	RD-181
Weight and dimensions characteristics
Total weight, kg	5 950
Dry weight, kg	5 480
Height, mm	3 600
Diameter, mm	3 200
Performance characteristics
Traction, tf:	vacuum: 423.4

By the beginning of work on the 11D520 and 11D521 engines, NPO Energomash (former names OKB-456 and KB EM) had experience in creating engines with high pressure in the CS, built in a closed circuit and operating on high-power components (AT and UDMG).

In particular, for ballistic missiles, engines 15D119 (RD-263/264) with thrust P z = 1040 kN (106 t) and pressure in the compressor compartment of 20.6 MPa, and 15D168 (RD-268) with thrust P z = 1147 kN (117 t) were created ) and with a pressure in the compressor chamber of 22.6 MPa. In the process of working on these engines, the plant at the design bureau improved the technology of steel casting of complex power parts (for example, pump housings and automation units, which were previously made of non-ferrous metals). To eliminate the occurrence of combustion instability in the liquid-propellant rocket engine chamber, plastic anti-pulsation partitions were installed, installed on the mixing head and contributing to the attenuation of pressure pulsations.

A certain groundwork was also provided by the development of the 8D420 (RD-270) engine with a thrust of 640 tons and a pressure in the compressor compartment of 26.1 MPa, operating according to the gas-to-gas circuit. Among other things, special TPU parking seals were developed for this engine to ensure multiple starts, and to reduce the weight and dimensions of the TPU, a design of booster pumps was developed with turbine blades located directly on the pump impeller-screw.

Experience in the design and experimental testing of large-scale engines and units operating at pressures up to 60 MPa, as well as mastered manufacturing technologies for such units, were used when working on the 11D520 and 11D521 engines.

The engine is made according to a closed circuit with afterburning of oxidizing generator gas after the turbine. Fuel components: oxidizer - liquid oxygen, fuel - kerosene. The engine consists of four combustion chambers, a turbopump unit (TPU), a fuel booster pump unit (FBU), an oxidizer booster pump unit (BNAO), two gas generators, an automation control unit, a cylinder unit, an automatic drive system (SPA), and a steering drive system ( PSA), a fuel flow regulator in the gas generator, two oxidizer throttles, a fuel throttle, oxidizer and fuel start-and-shut-off valves, four ampoules with starting fuel, a starting tank, an engine frame, a bottom screen, emergency protection system sensors, two heat exchangers for heating helium on pressurization of the oxidizer tank. One of the main design features of this engine is the presence of four chambers, swinging in two planes, and two gas generators operating on one turbine. Four combustion chambers made it possible to have chamber parameters in terms of thrust close to the mastered range: 185 tons of thrust, compared with 150 tons achieved in other developments. In addition, the presence of four chambers and two combustion engines allows for autonomous testing of these units.

Fig.1. Engine RD-170 (without steering gears; image enlarges when pressed)

The turbopump unit is located between the chambers, and its axis is parallel to the axis of the chambers. This solution makes it possible to optimally place the engine in the limited dimensions of the LV tail section.

To ensure maintainability of the structure, detachable flange connections are widely used. Self-sealing double-barrier seals with metal gaskets are used to ensure the tightness of large diameter stressed flanges.

During the development of the engine, provision was made to ensure the possibility of using it at least twenty times as part of the carrier, including inter-flight fire checks as part of the unit. Guaranteed engine performance reserves in terms of service life and number of starts in excess of those required in operation (before the last use) must be at least 5, required for one flight.

At the end of the 80s, the maximum number of tests on one engine was 21.

Table 1. Engine technical parameters

Parameter	Meaning		Units
Traction
near the Earth	740 000		kg
	7256		kN
in the void	806 000		kg
	7904		kN
Thrust throttling limits	100-40		%
Specific thrust impulse
in a vacuum	337		With
at sea level	309		With
Combustion chamber pressure	24.5	MPa
Consumption of fuel components through the engine	2393		kg/s
Component ratio	2.63		m(ok)/m(g)
Adjusting the ratio of components	±7		%
Working hours	140-150		With
Engine weight
dry	9755		kg
flooded	10750		kg
Dimensions
height	4015		mm
diameter in the cutting plane of the nozzles	3565		mm

The engine contains a combustion chamber 1, a turbopump unit 2, consisting of a turbine 3, a two-stage fuel pump 4 and a single-stage oxidizer pump 5, two gas generators 6, a fuel booster pump 7, driven by a hydraulic turbine 8, and an oxidizer booster pump 9, driven by is the gas turbine 10.

The oxidizer booster pump (BNAO) 9 is connected through pipeline 11 to the inlet of the oxidizer pump 5, the output of which is connected through the shut-off valve 12 to the manifold cavity 13 of the mixing head 14 of the gas generator 6. An oxidizer filter is installed at the BNAO inlet.

The fuel booster pump (BNAG) 7 is connected through pipeline 15 to the input of the first stage 16 of the fuel pump 4. The first stage of the fuel pump 16 is connected to the input of the second stage 17 of the fuel pump and through pipeline 18, in which the throttle 19 with electric drive 20 is installed, is connected to the manifold 21 combustion chamber 1, from which fuel is distributed through channels 22 of regenerative cooling of combustion chamber 1. A fuel filter is installed at the inlet of the BNAG.

The regenerative cooling channels 22 of the nozzle 23 are connected through a manifold 24 to a shut-off valve 25. The output of this valve is connected to a manifold 26 located on the cylindrical part of the combustion chamber. The output of the collector 26 through the regenerative channels 27 for cooling the cylindrical part of the combustion chamber is connected to the fuel cavity 28 of the mixing head 29 of the combustion chamber 1.

The second stage 17 of the fuel pump 4 (through which 20% of the total fuel consumption passes) is connected through a pipeline 30 to the main input 31 of the draft regulator 32, controlled by an electric drive 33 and having a check valve 34 at the input. The output 35 of the draft regulator 32 is connected to ampoules 36 ( 2 pcs.), filled with starting fuel triethylaluminum Al (C 2 H 5) h. The outputs from these ampoules through start-up shut-off valves 37 are connected to the cavity of the fuel 38 mixing heads 39 of the gas generators 6. The output of the gas generators 40 is connected to the turbine 3, the output of which is connected through pipelines 41 to the cavity 42 of the mixing heads 29 of the combustion chambers 1.

In addition, the outlet from the turbine 3 through a pipeline 43, in which a heat exchanger 44 and a pressure valve 45 are installed, is connected to the manifold of the turbine 46 driving the oxidizer booster pump 9.

The pneumohydraulic circuit of the liquid-propellant rocket engine also contains a starting system, which includes a starting tank 47 with a dividing membrane 48, a high-pressure gas supply pipe 49 and an outlet pipe 50. The outlet pipe 50 of the starting tank 47 is connected through a filling valve 51 to a fuel supply pipeline 15 from a fuel booster pump 7 In addition, the outlet pipe 50, on one side, through a pipeline 52, in which a check valve 53 is installed, is connected to the second input 54 of the draft regulator 32, through which the engine is started, and on the other hand, through a check valve 55, it is connected to the ampoule 56 , filled with starting fuel (hypergole), the output of which is connected through valve 57 to line 58 for supplying starting fuel to the ignition nozzles 59 of the combustion chamber. A nozzle 60 is installed in line 58, providing a dosed supply of starting fuel to the ignition injectors.

To reduce the aftereffect impulse, fuel shut-off valves are installed between the cooling paths of the nozzle and combustion chamber (valves 25), as well as in front of the manifold of the second and third curtain belts (shown in Fig. 2.2).

Pneumatic valves are actuated by helium from a block of high-pressure cylinders using solenoid valves.

Engine operation

The engine starts according to the “self-start” scheme. First, the drives 20 and 33 are installed in positions that ensure the initial installation of the thrust regulator 32 and the throttle 19. Then the rocket's tank valves are opened (not shown in the diagram) and, under the influence of hydrostatic head and boost pressure, the fuel components fill the cavities of the oxidizer and fuel pumps to the start-off valves 12 and 25 and check valve 34 of the draft regulator 32, respectively. The engine cavities are filled with fuel up to the starting ampoules 36 and 56 through the filling valve 51, check valves 53 and 55. The starting tank 47 is also filled with the main fuel. This condition is considered the initial state for starting the engine.

When the engine starts, the tank 47 is pressurized and fuel is forced out of it, the pressure of which breaks through the membranes (not shown) of the starting ampoules 36 and 56. At the same time, the start-cutting valves 12 and 37 and 25 are opened, respectively. As a result, starting fuel from ampoules 36 and 56, under the influence of pressure created by the starting tank, enters gas generators (through open valves 37) and chambers (through check valves 57). The starting fuel entering the gas generators is ignited with oxygen, which also enters the gas generators due to the pre-launch pressurization of the rocket tanks and the hydrostatic pressure in them. The fuel, having passed through the cooled path of the combustion chambers, after a fixed time enters the mixing heads of combustion chambers 1. During this delay time, the combustion process has time to begin in the gas generators and the generated generator gas spins turbine 3 of THA 2. After the turbine, the oxidizing gas flows through four cooled gas pipes 41 into the mixing heads 29 of the four combustion chambers, where it is ignited with the starting fuel coming from the ignition nozzles 59 and subsequently burned with the fuel entering the chambers. The time of entry of both components into the combustion chambers is selected so that THA 2 manages to reach operating mode while back pressure has not yet been established in chambers 1.

As the pressure behind the fuel pump 17 increases, the starting tank 47 is automatically switched off by closing the check valves 53 and 55, and the fuel supply to the gas generators 6 is switched to the pump 17 due to the programmatic opening of the draft regulator throttle 32.

Part of the oxidizing gas from the turbine outlet is taken to drive the two-stage gas turbine 10 of the booster prepump 9. This gas, passing through the heat exchanger 44, heats the gas used to pressurize the rocket tanks. After the turbine 10, the gas is discharged into the outlet manifold 11, where it is mixed with the main oxidizer flow and condensed. The use of gas taken from the output of the turbocharger turbine as a working fluid to drive the turbine of the oxidizer booster pump makes it possible to reduce the temperature in the gas generator and, accordingly, reduce the power of the turbocharger turbine.

Part of the fuel from the output of pump 4 is supplied to the drive of the single-stage hydraulic turbine 8 of the fuel booster pump 7.

A small part of liquid oxygen is taken from the gas generator collectors and enters the cooling path of the turbine housing and gas ducts.

At the entire stage of engine starting, the opening of the throttle of the draft regulator 32 and the fuel throttle 19 is programmed from the initial installation positions to the positions corresponding to the nominal engine mode using the corresponding drives 33 and 20.

In this way, the engine starts smoothly and returns to the main mode in 3 seconds.

Before switching off, the engines are transferred to the final stage mode, which is 50% of the nominal.

Table 1a. Simplified cyclogram of the operation of the 11D521 engine as part of block "A" of the Energia launch vehicle
(according to the flight program November 15, 1988)

№	Time (s) from the start command ("lift contact")	Description (Condition)
1	-3.2	Launch, starting thrust software set.
2	-0.2	Exit to the main thrust stage.
3	38	Start of program throttling to reduce the speed pressure.
4	74	End of program throttling to reduce speed pressure.
5	108.5	Start of software throttling to limit longitudinal overload to 2.95 units.
6	130	Switching the engine to the final thrust stage mode 49.5%.
7	142	Turning off the engines.

The chamber is a soldered-welded one-piece unit and consists of a mixing head, combustion chamber and nozzle. The chamber is attached to the gas path using a flange connection.

Table 2. Camera technical parameters

Parameter	Meaning		Units
Reduced length of KS	1079.6		mm
Diameter KS	380		mm
Minimum nozzle diameter	235.5		mm
Degree of contraction subsonic nozzle parts	2.6
Nozzle outlet diameter	1430		mm
Supersonic expansion ratio nozzle parts	36.87
Chamber length	2261		mm
Temperature in KS	3676		K
Pressure in the CS	24.5		MPa
Pressure at the nozzle outlet	0.072		MPa
Thrust coefficient
in a vacuum	1.86
at sea level	1.71
Camera Deflection Angle	8		degrees

Fig.4. Diagram of fuel supply to the cooling tract of the chamber: gas pipeline middle bottom of the mixing head front (fire) bottom of the mixing head nozzles forming anti-pulsation partitions (total 54 pcs.) main injectors supply of ignition mixture (4 nozzles fed from a separate manifold) air curtain top manifold fuel supply manifold for cooling the cylindrical part of the compressor station collector of the middle 26 and lower 27 curtain belts main manifold for supplying fuel to the compressor station external load-bearing wall of the compressor station manifold for removing fuel from the nozzle cooling path internal wall of the CS fuel supply manifold for cooling the nozzle outlet nozzle fuel moves to the nozzle exit through even (conventionally) and returns through odd channels fuel supply for cooling the nozzle outlet fuel supply from the pump supply of fuel to the middle and lower belts of the curtain septum in channels cylindrical part of the KS mixing head central nozzle gas cavity of the mixing head perforated rear bottom of the mixing head middle belt of the curtain lower belt of the curtain

The chamber body consists of a combustion chamber and a nozzle. The chamber body includes an outer power shell 11 and an internal fire wall 13 with milled channels forming an external regenerative cooling path for the chamber, which has three cooler inlets. The first input is connected to the cooling path of the critical section of the nozzle, the second input is connected to the cooling path of the nozzle outlet, and the third is connected to the cooling path of the combustion chamber. In this case, the first output is connected to the third input, and the first input, the second input and the supply to the two lower belts of the slot curtains are united by a common pipe, branched and located outside the chamber.

Internal cooling is provided by three belts of slotted curtains in the subcritical part of the combustion chamber. Through them, about 2% of the fuel is supplied to the wall in the form of films that evaporate and protect it from heat flows, which in the critical section of the nozzle reach values ​​of the order of 50 MW/m 2.

The ignition means are made of four jet nozzles 6 equally spaced around the circumference, installed behind the front (fire) bottom 3 in the power body of the chamber 11. The axes of the flow holes of the jet nozzles are located at an acute angle to the exit from the power body and are deflected in a circle in the transverse plane from the longitudinal axis power housing in the same direction, and the axis of the flow hole of each jet nozzle is crossed with respect to the axes of the flow holes of the nozzles adjacent to it. The injectors are hydraulically connected by a common manifold.

All nozzles are two-component with an axial supply of oxidizing gas and a tangential supply of fuel. The nozzles located near the fire (inner) wall of the chamber are made with increased hydraulic resistance along the fuel line compared to other nozzles due to the reduction in the diameters of the fuel supply holes, i.e. providing reduced fuel consumption compared to other injectors.

To suppress pressure pulsations, the initial zone of mixture formation and combustion, in which, as a rule, high-frequency oscillations arise, is divided into seven approximately equal volumes using anti-pulsation partitions consisting of nozzles protruding beyond the fire bottom, which do not fit tightly together along their cylindrical generatrices. Due to this, the natural frequencies of vibrations in the volumes between the partitions sharply increase, moving far from the resonant frequencies of the combustion chamber structure. In addition, protruding nozzles stretch the combustion zone, which also reduces the possibility of high-frequency phenomena. The gaps between protruding nozzles that are loosely adjacent to each other have an additional damping effect.

The part of the nozzle protruding beyond the fire bottom is cooled by fuel passing through spiral channels (screw swirler) 6 of the inner sleeve.

The remaining nozzles are recessed into the fire bottom (their outlet cavities 4 extend into conical bores 5 in the fire bottom 7) and are made with different hydraulic resistance when supplying fuel, divided according to the mass flow of fuel into three groups with the possibility of ensuring a difference in fuel consumption between each group of 3% up to 10% at nominal mode. In this case, the nozzles (except for those located near the fire wall of the chamber) are fixed in the fire bottom and middle bottom so that nozzles from different groups are adjacent to each other by cyclically sequentially repeating the arrangement of nozzles from the first to the last group.
The introduction of injectors with different flow rates is necessary in order to reduce the effects of high-frequency vibrations on engine operating conditions.

Fig.6.2 Location of nozzles on the mixing head (images enlarge),

Each of the four chambers is equipped with a swing unit. The traction force is transmitted from the camera to the power frame through a gimbal. The supply of generator gas fired at the turbine to the compressor station is carried out through a 12-layer composite bellows located inside the gimbal. The bellows is armored with special rings and is cooled by a small amount of cold oxygen flowing between the inner surface of the bellows and the thin inner wall.

Fig.8. Swing unit diagram

The swing unit consists of support rings 9 and 10, which are respectively hermetically connected to the combustion chamber and the gas duct (exit from the turbine), which contain consumable elements of external flow cooling 11 and 12, also shown in the view A. The bellows 13 is located inside the cardan ring 14. The cardan ring 14 is connected through hinges 15, forming two rotary axes, by power brackets 16 and 17 with support rings 9 and 10. Inside the bellows 13 there are two shells 18 and 19, each of which is a body of rotation and cantilevered, respectively, to one of the mentioned support rings, and the free end of the shell 18 is made in the form of a nipple with a spherical end 20 and is installed with a gap A in the shell 19. The center of the nipple sphere with a spherical end 20 is located on the swing axis of the chamber. The size of the specified gap is chosen to ensure the flow of the cooling working fluid (oxidizer) necessary for reliable cooling of the bellows 13.

The bellows 13 is made multilayer and is equipped with protective rings 21, inserted between the corrugations 22 of the bellows 13. Outside the protective rings 21, a tightly fitting casing 23 is installed, made of layers of cylindrical spirals 24 connected at the ends to the support rings 9 and 10 of the bellows assembly. Adjacent layers of spirals are adjacent to each other, and their turns are wound in opposite directions.

Installation of a metal power casing in the form of a metal cylindrical spiral outside the protective rings 21 of the bellows 13 increases its strength properties and at the same time limits the spontaneous bending of the bellows 13 when the engine chamber is rotated at relatively large angles (10-12°), thereby increasing its stability.

The turbopump unit is made according to a single-shaft design and consists of an axial single-stage jet turbine, a single-stage centrifugal screw oxidizer pump and a two-stage centrifugal screw fuel pump (the second stage is used to supply part of the fuel to gas generators).

Table 3. TNA Parameter Meaning Units oxidizer fuel Pump outlet pressure 60.2 50.6 MPa Component flow through the pump 1792 732 kg/s Impeller diameter 409 405 mm Efficiency pump 0.74 0.74 Shaft power 175 600 77 760 hp 129.2 57.2 MW Shaft rotation speed 13 850 min -1 Turbine power 257 360 hp 189.3 MW Turbine inlet pressure 50.9 MPa Number of turbine stages 1 Turbine pressure reduction ratio 1.94 Turbine inlet temperature 772 TO Efficiency turbines 0.79

On the main shaft with the turbine there is an oxidizer pump, coaxially with which two stages of the fuel pump are located on another shaft. The shafts of the oxidizer and fuel pumps are connected by a gear spring to unload the shaft from thermal deformations that arise as a result of the large temperature difference between the working bodies of the pumps, as well as to prevent freezing of the fuel.

Fig. 10. Shaft with turbine, centrifugal screw wheel of oxidizer pump,
bearings and impeller seals

To protect angular contact shaft bearings from excessive loads, effective auto-unloading devices have been developed.

In an engine of a closed oxidation circuit, the protection of the oxygen path units of the fuel pump from fire when exposed to accidental fire initiators is of particular importance. Due to the exceptionally high pressure in the duct of the 11D520 and 11D521 engines, as well as the high mechanical loads characteristic of a powerful engine, the problem of combustion protection during their creation was especially acute.

To prevent fire due to breakdowns of structural elements or friction of rotating parts against stationary ones (due to the selection of gaps from deformations or work hardening on the mating surfaces from vibration), the gap between the blades of the nozzle apparatus and the rotor is made relatively large, and the edges of the blades are made relatively thick.

To prevent fire and destruction of turbine gas path parts, nickel alloys are used in the design, including heat-resistant ones for hot gas lines. The turbine stator and exhaust tract are forcibly cooled with cold oxygen. In areas of small radial or end clearances, various types of heat-protective coatings are used (nickel for the rotor and stator blades, metal-ceramic for the rotor), as well as silver or bronze elements, which prevent fire even if there is a possible contact with the rotating and stationary parts of the turbopump unit.

To reduce the size and mass of foreign particles that could lead to a fire in the gas path of the turbine, a filter with a cell of 0.16x0.16 mm was installed at the engine inlet.

The high pressure of liquid oxygen and, as a consequence, increased fire risk determined the design features of the oxidizer pump.

Thus, instead of floating sealing rings on the impeller collars (usually used on less powerful pumps), fixed gap seals with a silver lining are used, since the process of “floating” of the rings is accompanied by friction at the points of contact of the impeller with the housing and can lead to fire of the pump.

The screw, impeller and torus outlet require particularly careful profiling, and the rotor as a whole requires special measures to ensure dynamic balance during operation. Otherwise, due to large pulsations and vibrations, destruction of pipelines and fires at joints occur due to the mutual movement of parts, friction and hardening.

To prevent fire due to breakdowns of structural elements (screw, impeller and guide vanes) under conditions of dynamic loading with subsequent fire due to rubbing of debris, means were used such as increasing structural perfection and strength due to geometry, materials and cleanliness of mining, and also the introduction of new technologies: isostatic pressing of cast billets, the use of granular technology and other types.

The oxidizer booster pump consists of a high-pressure screw and a two-stage gas turbine, which is driven by oxidizing gas taken after the main turbine with its subsequent bypass to the inlet of the main pump.

Fig. 11a. Simplified diagram of an oxidizer booster pump unit (image enlarges).

The composite housing, consisting of housings 1 and 2 connected by a flange, has a bushing 4 fixed to the power ribs 3, the internal cavity of which is closed by a fairing 5. Inside the bushing 4 there is a ball bearing 6 mounted on the pump impeller, made in the form of a screw 7. Fairing 5 The liner 8 installed in the sleeve 4 is pressed in. The liner 8 has holes 9 connecting the cavity of the liner 8 with the high-pressure channel 10. Housing 2 contains a fairing 11, fixed in it using straightening blades 12. This fairing contains a ball bearing 13, secured with a nut 14 on the auger 7. The auger has blades 15. Along these blades, the auger is inserted into the impeller of the turbine 16 (which actually consists of two stages, and not one, as shown in the simplified diagram) and welded with it, i.e. The turbine impeller is fixed to the peripheral part of the pump impeller. The turbine impeller has profiled blades 17, the inter-blade spaces of which are connected by nozzles in the nozzle apparatus with the inlet manifold. The combustion products with excess oxygen are supplied through the inlet pipe 18. The turbine outlet cavity, made in the housing 2 in the form of an annular cylindrical cavity, communicates through channels 19 with a conical annular pipe 20, which through holes 21 communicates with a cylindrical outlet 22.

When the BNAO is operating, liquid oxygen is supplied to the pump inlet (shown by an arrow), and combustion products with excess oxygen taken from the gas duct after the turbine of the main HPU (see ASG in Fig. 2) are supplied to the turbine inlet (shown by an arrow). The combustion products then fall onto the profiled blades 17 of the turbine, providing liquid oxygen supply by the screw 7. Behind the turbine, the combustion products through the holes 19 enter the cavity of the pipe 20, and then through the holes 21 to the pump outlet, where they mix with liquid oxygen and condense. To solve the problem of the occurrence of low-frequency pulsations during gas condensation, fragmentation of the flow discharging gas was used.

Unloading the screw 7 from the actions of axial forces is ensured by supplying high-pressure liquid oxygen (see Fig. 2.2) through the high-pressure channel 10 into the high-pressure cavity of the auto-unloading device. In the place of small gap between the impeller and the housing in the high-pressure cavity of the auto-unloader, a silver lining is used to prevent fire from possible contact.

In the line for supplying combustion products to the BNAO turbine, a newly developed “hot gas” valve (45 in Fig. 2.1) is installed, which operates under conditions of oxygen generator gas with high temperature and high pressure.

The fuel booster pump consists of a high-pressure auger and a single-stage hydraulic turbine operating on kerosene taken after the main pump.

Structurally, the fuel booster pump is similar to the oxidizer booster pump with the following differences:

• a single-stage hydraulic turbine operates on fuel taken from the outlet of the fuel pump of the main HPU;
• High pressure fuel is removed to relieve the auger from axial actions from the inlet manifold of the BNAG hydraulic turbine.

Fig. 12. Fuel booster pump unit

Fig. 13. Gas generator

A single-zone gas generator producing gas with an excess of oxidizer to drive the turbine consists of a solder-welded housing with a spherical outer shell and an outlet pipe rigidly connected to it, a cylindrical fire chamber with a diameter of 300 mm and a mixing head equipped with two-component and two-stage oxidizer nozzles, design which is made with a combustion zone and a gas ballasting zone inside the nozzles. In fact, each nozzle, together with the channel of the thick-walled fire bottom in which it is located, forms an individual two-zone gas generator. As a result, the uniformity of the temperature field across the cross section of the total gas flow formed by such nozzles is ensured at high flow rates.

Fig. 14a. Gas generator diagram:
1 - spherical power shell; 2 - outlet pipe; 3 - cover; 4 - bushing; 5 - fire bottom; 6 - through chambers in the fire bottom; 7 - oxidizer cavity; 8 - spacer (outer wall of the fire chamber); 9 - annular cavity; 10 - shell (inner wall) of the fire chamber; 11 - fire chamber; 12 - mixing module (nozzle); 13 - mixing module housing; 14 - fuel channel; 15 - ring channel of the oxidizer; 16 - mixing chamber; 17 - fuel supply pipe; 18 - fuel cavity; 19 - oxidizer supply pipe; 20 - windows in sleeve 4; 21 - tangential holes for supplying oxidizer; 22 - grooves on the outer surface of the nozzle body; 23 - calibrated fuel supply channels; 25 - tangential holes for fuel supply; 26 - conical borings; 27 - cooling cavity; 28 - channels forming the cooling cavity; 29 - holes for supplying oxidizer to the cooling cavity; 30 - annular slot for the exit of the oxidizer from the cooling cavity.

When the gas generator is operating, fuel from pipe 17 fills cavity 18 and is supplied through calibrated channels 23 and tangential holes 25 into channels 14 and then into mixing chambers 16. The oxidizer through pipe 19 is fed into the annular cavity 9, and through windows 20 fills cavity 7. Part of the oxidizer through tangential holes 21 enters the mixing chamber 16, where, mixing with fuel, it causes fire. Through grooves 22, the oxidizer is also supplied to chamber 6, ensuring mixing of high-temperature combustion products. Next, in the fire chamber 11, the high-temperature combustion products are cooled with simultaneous evaporation of the liquid and heating of the gaseous oxidizer. At the exit from the gas generator, an oxidizing agent is mixed with the gas generation products, supplied through the annular slot 30.

Fig. 14b. TNA with gas generators

The gas generator provides oxidizing gas at the output in a wide temperature range (from 190 to 600°C), which allows you to regulate engine thrust from 30 to 105% of the nominal value.

The connection between the housing and the mixing head is carried out using a split flange. To ensure tightness, a seal with metal gaskets is used.

To ensure an acceptable level of temperature stress in the load-bearing body parts, the gas ducts between the gas generators, turbine and chambers are cooled with oxygen.

To prevent fire in the gas ducts, the swing units of the chamber mixing head, and the oxidizer valve, increased (compared to less powerful engines) requirements for the cleanliness of gas paths and the prevention of the presence of organic substances are installed.

		The ampoule contains a housing 1 with inlet 2 and outlet 3 pipes of membrane units 4 and 5 installed inside the housing 1, and a means for filling the housing with starting fuel 6. Each membrane unit 4, 5 contains a piston 7, which can be made integral with the membrane 8 or in which the membrane 8 is sealed with its outer surface. Piston 7 is installed in the guide 9 of the housing in a sliding fit. The peripheral section of the membrane 8 is hermetically sealed with the body 1 under the guide 9. The piston 7 is connected to the shank 10, which can be cylindrical or of any other shape and placed in the sleeve 11. The sleeve 11 is attached to the body 1 of the ampoule on brackets 12. The sleeve 11 has a spring retainer 13, for example, made in the form of a spring ring, and the shank 10 is made with an annular groove 14. When the membrane unit is activated, the spring clamp 13 limits the movement of the shank 10. The shank 10 is made with holes 15 for bleeding gas from the stagnant zone when filling the ampoule. The membrane 8 on the side of the inlet 2 is made thin in the form of an annular jumper 16, which breaks when interacting with the working medium at diameter D. Dimension D is slightly smaller than the diameter of the piston 7. At the junction of the membrane 8 with the piston 7 it is made with a smaller thickness in order to exclude scuffing when piston 7 moves in guide 9 of housing 1.
Fig. 16. Diagram of an ampoule with starting fuel (image enlarges).

The design includes a means for filling the housing with starting fuel 6, which is installed in the partition 17 of the housing 1 and consists of two plugs - a filling plug 18 and a drain plug 19, which are installed respectively in the filling 20 and drain 21 channels. Each of the plugs has a screw plug 22, a sealing plug 23, a sealing gasket 24 and a nut 25. The screw plug 22 has a flow hole 26.

Filling the ampoule with starting fuel is carried out as follows. On the assembled ampoule, before installing the nuts 25 and sealing plugs 23, the threaded plugs 22 are not fully screwed in, so as to ensure the opening of the flow area of ​​the filling 20 and drain 21 channels through hole 26. Filling with starting fuel is carried out by feeding it through the filling channel 20 into the internal housing cavity 1 between membrane units 4 and 5, and then through the drain channel to the drain. After filling the ampoule, screw the screw plugs 22 until they stop, then drain the starting fuel before the threaded plug 22 of the filling plug 18 and after the threaded plug 22 of the drain plug 19. After this, install the sealing plugs 23, sealing gaskets 24 and nuts 25. After this, the ampoule is ready for installation on a rocket engine. In the internal cavity of the ampoule in body 1 between the membranes 8, a gas cushion is formed as a result of assembling and filling the ampoule. The presence of a gas cushion helps ensure the reliability of the ampoule during storage and the effective movement with acceleration of the piston 8 when medium pressure is applied to the inlet of the ampoule.

The device works as follows. When a high-pressure component from the input side of the membrane assembly 4 is exposed to deformation of the membrane 8, and then destruction along the circumference D. If the membrane 8 is unevenly destroyed, with the appearance of leakage, the pressure in front of the piston 7 does not drop, due to the operation of the throttling gap formed by the housing guide 9 and piston 7, piston 7 continues to move, and after complete destruction of membrane 8 it accelerates. The accelerated movement of the piston 7 is ensured due to the presence of force from the pressure difference acting on the surface area determined by the diameter D.

The length “A” over which the piston moves with acceleration and the gap between the piston 7 and the guide 9 are selected to ensure guaranteed cutting of the membrane 8 along the entire perimeter, the required delay in opening the flow area of ​​the line after cutting the membrane 8, the acceleration of the piston 7 required for operation spring retainer 13. The dimensions of the membrane jumpers 8 are determined based on the specified pressure, ensuring the destruction of the jumper.

Next, the moving shank 10 along the flow is fixed using a spring clamp 13, while the hydraulic characteristics of the open membrane unit 4 are reproduced with high accuracy, since there are no structural elements with an uncertain position in the flow of the component.

After opening the membrane unit 4, due to the increased pressure of the starting fuel, the membrane unit 5 opens in a similar way.

The RD-170 and RD-171 engines use different options for swinging the chambers and controls for their deflection.

The chambers of the RD-170 engine as part of block A of the Energia rocket swing in two planes: in a radial plane passing through the longitudinal axis of the engine and the axis of the chamber, and in a tangential plane perpendicular to it. This control scheme is more effective in the structure of the Energia rocket package, but requires more powerful steering machines that overcome the load created by the oncoming aerodynamic flow on the protruding part of the combustion chamber nozzle beyond the parameter of the outer contour of the block when it is deflected in the radial direction.

The combustion chambers of the RD-171 engine of the first stage of Zenit are deflected when controlled only in the tangential rolling plane. The chamber nozzles do not extend into the aerodynamic flow around the stage and do not experience its load. Steering gears are significantly less powerful. The control efficiency of this option is sufficient for the Zenit missile.

The remaining engine systems are unified.

At the final stage of engine testing V.P. Glushko initiated the development of a more advanced engine design, which, compared to the RD-170 (RD-171) engine, provided higher thrust (5% boost) and in which measures should be implemented to reduce the dynamic stress of the feed units. The corresponding design documentation was developed and the engine was eventually named RD-173.

Until 1996, 28 engines were manufactured and underwent various testing. The RD-173 engines use a more advanced design of feed units, primarily the main pump. The RD-170 engine control system has undergone major modifications. In the process of testing the RD-173, it was confirmed that engine startup and its operation in all envisaged modes are characterized by stable operation of all units and systems, ensuring the required startup character and accuracy of maintaining parameters without using the oxidizer chokes. The exclusion of oxidizer chokes and, accordingly, two drives from the engine structure simplified its design, increased reliability and reduced engine weight. The design of the bellows of the swing unit made of nickel alloy was introduced, which also increased the reliability of the engine.

The accumulated experience in setting up the engine control system in the process of technological control tests using external feedback made it possible, in the process of testing the RD-173 engine, to switch to a significantly simpler control system, consisting of two digital drives that directly control the traction control and the SOB throttle. Simplification of the control system increased the reliability of the engine and reduced its weight.

In the RD-173 engine, taking into account the large positive statistics of the operation of gas generators, the mixing heads are welded, in contrast to the flange connection in the RD-170 (RD-171) engines, where the possibility of quickly replacing the head after a technological control test was provided. This, as well as other solutions obtained during testing of the RD-173 engine, were used in the development of the RD-180 engine.

Orders for the production of RD-171 engines ceased in 1995. At the same time, NPO Energomash continued to produce a more advanced modification of the RD-170 (RD-171) engines - the RD-173 engine. Since 1995, NPO Energomash has supplied RD-171 engines for the Sea Launch program, which were modified from RD-170 engines previously manufactured for the first stages of the Energia launch vehicle. These engines created the basis for the implementation of the program until 2004. For further development of the program, it became necessary to resume engine production at NPO Energomash. Taking into account the accumulated experience in testing the RD-173 and RD-180 engines, in which solutions were introduced aimed at increasing reliability and ensuring boost by 5%, NPO Energomash proposed to manufacture RD-173 engines for the Sea Launch program. This proposal was supported by the lead developer of the Zenit launch vehicle, the Yuzhnoye State Design Bureau, and approved by the customer of the launch vehicle. The engine received the designation RD-171M. Certification of the RD-171M engine was completed on July 5, 2004. 8 tests lasting 1093.6 seconds were carried out on the certification engine, with the last test (above plan) at 105% mode. The first commercial RD-171M engine was delivered to Ukraine on March 25, 2004 after a technical test lasting 140 seconds.

In 2006, the RD-171M engine was certified for use in the Zenit-M launch vehicle during the implementation of government programs of the Russian Federation.

The technical diagnostic system was developed in parallel with the creation of the engine as a means of assessing the technical condition of the engine and predicting its performance. In addition, it was used to analyze failures and defects, since it made it possible to more deeply study the interrelation of parameters and their statistical characteristics.

The system is a set of technical means, diagnostic methods and diagnostic object, as well as organizational and technical measures for collecting, converting, storing, analyzing information and making decisions about the condition of the engine. The system must provide identification of the location and causes of malfunctions.

The technical diagnostic system has the following subsystems:

• information and measurement;
• functional diagnostics;
• test diagnostics as a non-destructive method of condition monitoring.

During the development of the diagnostic system, the following were created:

• technique for monitoring the stability of startup characteristics, main mode and final stage mode. The technique was intended to estimate the values ​​of slowly changing parameters and their rates obtained during fire tests, taking into account the field of permissible boundaries;
• methodology for tolerance control of parameters in the main mode and final stage mode; it was intended to assess the compliance of engine parameters measured during fire tests with calculated values ​​​​obtained using mathematical models and model characteristics of units based on their autonomous tests, which is determined by finding the parameters in the tolerance field;
• technique of contour linking of slowly changing parameters; was intended to evaluate the functioning of the engine as a whole and its circuits in stationary modes by comparing the measured and calculated values ​​of slowly changing parameters at characteristic points;
• methodology for assessing stability and determining vibroacoustic characteristics; was intended to control the level of pulsation and vibration for compliance with statistical tolerances and assess the stability of the combustion chamber and gas generator, with an analysis of the physical nature of the spectra and determination of oscillation damping decrements;
• methodology for estimating the amount of exhausted life of assembly units; it is based on the theory of high-cycle fatigue of materials and takes into account dynamic loads caused by pulsations and vibrations; the integral value of fatigue damage was assessed during technological control tests, its value during operation was predicted, and their sum was compared with the limit value determined based on the results of multi-resource tests;
• parametric control technique - used for diagnostics in stationary modes in order to localize faults; the analysis is based on assessments of the functional characteristics of the units;
• a set of non-destructive testing methods.

In mass production, each engine, after manufacturing and a full control cycle, undergoes autonomous control technological tests, which are carried out at the manufacturer's firing stand with the engine starting according to the full flight program or a slightly accelerated one. After fire bench tests, the engine may undergo a rebuild. This means that in order to ensure that the quality of the design is maintained after fire tests, partial disassembly of individual units is carried out.

1. Gubanov B.I. Triumph and tragedy of "Energy"
2. George P. Sutton. Rocket Propulsion Elements, 7th edition
3. Katorgin B.I. Prospects for creating powerful liquid rocket engines
4. George P. Sutton "History of Liquid Propellant Rocket Engines"
5. Prospect NPO "Energomash"
6. Description of the invention to the patent of the Russian Federation RU 2159351. Gas generator ( US Patent 6244040).
7. Description of the invention to the patent of the Russian Federation RU 2159349. Gas generator module ( US Patent 6212878).
8. Description of the invention to the patent of the Russian Federation RU 2158841. Liquid rocket engine chamber and its housing ( US Patent 6244041).
9. Dobrovolsky M.V. Liquid rocket engines. - M.: MSTU, 2005.
10. Description of the invention to the patent of the Russian Federation RU 2159352. Swing unit of the LRE chamber with afterburning.
11. Description of the invention to the patent of the Russian Federation RU 2158839. Liquid rocket engine with afterburning turbogas ( US Patent 6226980
12. NPO "Energomash" named after academician V.P. Glushko. The path in rocket technology. Ed. B.I. Katorgina. M., Mechanical Engineering-Polyot, 2004.

One of the most important parts of the engine is the turbopump unit for supplying oxygen and kerosene to the combustion chamber.

Shaft with turbine, centrifugal screw wheel of the oxidizer pump, bearings and impeller seals.
More details http://www.lpre.de/energomash/RD-170/index.htm

In the 50s, Soviet and American specialists, almost independently of each other, found a way out of the impasse. (By the way, it was after this that the era of space rockets began.) The nozzle shell was made of two layers, between which the coolant flowed: the inner thin wall transferred the heat of hot gases well to it, the outer thick one absorbed power loads. Behind the apparent simplicity was the titanic work of technologists; it is not so easy to combine three components into a single whole...

RD-170 at the stand.

By our time, the production of double-layer casings had been brought to perfection, and in order to increase engine power, something fundamentally new was required. This is what was embodied in the RD-170. It artificially creates conditions under which the region of maximum temperatures is located along the axis of the combustion chamber, and at its periphery it is much “cooler”. The latter is achieved by changing the optimal ratio of fuel (kerosene) and oxidizer (oxygen).

Excess kerosene is injected into the peripheral area through additional nozzles. In addition, part of the kerosene, which played the role of a coolant, seeps through the capillary holes on the inside of the nozzle. That is, a fire raging near the walls is partially extinguished... with fuel! This made it possible to increase the temperature in the core of the chamber, and consequently, the engine power.

It is growing thanks to one more feature. The fact is that it is not so easy to achieve complete combustion of the entire fuel mixture inside the chamber; some of it, although small, is usually carried out of the nozzle. Therefore, a “cocktail” of fuel and oxidizer must be prepared very quickly and efficiently. Designers tried all kinds of mixers and nozzles: jet, slot, lattice, vortex, centrifugal... And in the 60s, the RD-253 (it launches Protons into space) used something that would make any fire specialist flinch safety: self-igniting components were mixed directly in the pipeline, up to the combustion chamber! Of course, we had to take into account a lot of subtleties, but the main thing is that the engine worked successfully. However, for almost 30 years no one dared to repeat such a scheme. Before the advent of the RD-170.

The figure shows that already in the turbopump unit all the oxygen supplied to it and part of the kerosene are mixed. The designers made a fire in the pipeline the design mode of the engine - due to an excess of oxygen, the temperature of the mixture (and here its composition is not optimal for combustion) rises to only 400 ° C. However, what is a hot mixture with an excess of oxygen? A very aggressive environment, fatal to any metal. The walls of the pipeline, of course, can be made very thick, but in the path of the corrosive flow there is a thin and flexible bellows pipe. You can’t do it differently - when controlling a rocket, the engine must rotate in two planes by 6 - 8 degrees. Here, chemists have already tried their best and created a unique nickel alloy for the pipe (the composition of which, of course, is classified), capable of withstanding an aggressive mixture with a pressure of 270 - 300 atm.

In the combustion chamber it combines with heated kerosene that has passed through the cooling jacket, and now the flame is raging with might and main: although the pressure drops to 250 atm, the temperature reaches 3500°C! At the same time, on the walls (we already know why) it is approximately 2800°C lower. The gas escapes from the nozzle with a specific impulse of 330 s and creates a thrust of 800 t/s (with an engine mass of about 11 t).

Much in the RD-170 aroused admiration among American specialists. But for NPO Energomash this stage has already been passed. On the desk of the General Director, Doctor of Technical Sciences Boris Katorgin, there are already drawings of the world's first three-component (oxygen, hydrogen, kerosene) liquid propellant rocket engine. For now it was called RD-701. The engine weight will be 1.8 tons, and it will develop a maximum thrust of 200 t/s. It will operate in two modes, consuming 6% hydrogen, 12.6% kerosene and 81.4% oxygen after start, and with further acceleration - no kerosene at all. The application of the new engine has already been determined - space shuttles taking off from Mriya-type aircraft.
http://epizodsspace.no-ip.org/bibl/tm/1993/6/rd-170.html
The key moment in the international activities of NPO Energomash should be considered 1992, when on October 26, an “Agreement on Joint Marketing and Technology Licensing” was signed with Pratt & Whitney of United Technologies Corporation, in which NPO Energomash appointed United Technologies Corporation as its exclusive marketing representative with respect to the manufacture, use or sale of propulsion systems and licensed technologies in the United States.

In accordance with the signed agreement, NPO Energomash and Pratt & Whitney carried out active and successful marketing activities. In January 1994, in the published report of NASA headquarters “Access to Space,” the possibility of using engines developed by NPO Energomash as the main propulsion engines of American space launch vehicles was officially mentioned for the first time. Such an engine could be the RD-180 engine, a two-chamber derivative of the RD-170 engine used in the first stages of the Zenit and Energia launch vehicles.

In addition, as part of one of the contracts, on October 11-25, 1995 in West Palm Beach, Florida, three bench launches of the RD-120 rocket engine developed by NPO Energomash were successfully carried out at the Pratt & Whitney company's firing stand. In a short time, a large complex of work was completed in the United States to prepare the American test base for fire tests of the Russian serial liquid propellant rocket engine. The success of this program served as strong evidence of the real feasibility of fruitful cooperation between Russian and American specialists.

Also in 1995, Lockheed Martin announced a competition for an engine for its new Atlas IIAR launch vehicle. At the first stage, two American companies competed for the right to present the new RD-180 engine developed by NPO Energomash for Atlas IIAR - Pratt & Whitney and Rocketdyne. In August 1995, the choice was made in favor of Pratt & Whitney. In addition to the RD-180 engine project, the NK-33 engine from the Russian enterprise Trud named after. N.D. Kuznetsov from Samara and a version of the MA-5 engine from Rocketdyne. On January 12, 1996, in Denver, Colorado, Lockheed Martin announced the selection of the RD-180 liquid rocket engine as the engine for the first stage of the Atlas IIAR launch vehicle.

In a very short time, NPO Energomash carried out a large amount of work on the development of the engine, including fire tests at the NPO Energomash stand. In 1998, four successful demonstration fire tests of the RD-180 No. 4A engine were carried out in the United States. As a result, a new RD-180 rocket engine was developed, which in March 1999 was certified for use in the Atlas III launch vehicle.

A lot of work was done by the Foreign Economic Activity Service to obtain government support for the Russian-American project to develop and sell the RD-180 engine. The Ministry of Defense of the Russian Federation and the Russian Space Agency provided great assistance in this. In close cooperation with these organizations, in 1997, a decree of the President of the Russian Federation was prepared and signed, authorizing NPO Energomash to sell the RD-180 engine on the American market and organize parallel production of this engine in the USA as part of a joint American-Russian venture.

On January 27, 1997, NPO Energomash and Pratt & Whitney signed an Agreement on the establishment of a limited liability company RD AM ROSS, LLC. The joint venture was created for marketing, sales and organization of a production base in the USA for the parallel production of RD-180 engines and their modifications.

On May 16, 1997, a five-party Agreement was signed on the use of RD-180 engines produced by NPO Energomash and on supporting parallel production of RD-180 in the USA, in which the Russian Space Agency, NPO Energomash, Lockheed Martin, RD AMROSS and Pratt & Whitney stipulated mutual obligations in the event that Lockheed Martin Astronautics wins the final stage of the EELV competition. In this document, Lockheed Martin guaranteed the purchase of 101 commercial RD-180 engines.

The peculiarity of the Russian-American project, in which NPO Energomash participates, is that the main contractor, the American company Lockheed Martin, almost simultaneously developed two new launch vehicles, one of which (Atlas III) was intended primarily for launching commercial payloads, and the other (Atlas V) was developed under the EELV (Enhanced Explosive Launch Vehicle) program and should become the basis of a whole family of medium and heavy launch vehicles used in space launches in the interests of both the US government and commercial customers .

Currently, the RD-180 engine is certified for use in Atlas V launch vehicles (EELV) of both medium and heavy class.

On March 28, 1997, a contract was signed for the supply of RD-180 rocket engines to the USA between NPO Energomash and RD AMROSS, LLC.

The first commercial RD-180 engine was delivered to the USA on January 2, 1999. At the beginning of September 2011, 55 commercial engines had already been delivered to the USA. Six launches of the Atlas III launch vehicle with RD-180 engines were carried out (the first was on May 24, 2000). All launches took place without any comments on the operation of the engines.

Among the major flights commissioned by NASA are the launches of the Lunar Surface Orbiter and Lunar Crater Survey System (LRO/LCROSS), the spacecraft to explore the surface of Mars, the spacecraft to explore Pluto and its moon Charon as part of the Flight to New horizons for Pluto”, “Solar Dynamics Observatory” to obtain qualitatively new scientific data on the study of the Sun. By the end of 2011, the Mars Science Laboratory is planned to be launched on the Atlas 5 launch vehicle.

All major inventions used in the development and production of the RD-180 engine are protected by international patents. Obtained 20 US patents and 13 patents from the European Patent Office.

All information about Russian-American cooperation in the field of RD-180, which is available in Russian, represents the worst kind of lie - half-truths. Where individual completely truthful facts are intertwined with concealment of key information and sealed with targeted, unnoticeable lies.

Yesterday, as soon as I wrote an article about Russian space fakes, I was immediately bombarded with an “example” of the US lagging behind Russia in the space sector. They say that American rockets fly on Russian RD-180 engines, and without these Russian engines the American space program will immediately stall. With a bunch of links. So, they say, Americans won’t go anywhere without their mother’s dissipation.

Clicking on the links sent to me showed that all the information about Russian-American cooperation in the field of RD-180, which is available in Russian, represents the worst kind of lie - half-truths. Where individual, completely truthful facts (engine production is entirely concentrated in Russia) are intertwined with the concealment of key information and sealed with targeted, unnoticeable lies.

Let's start with the fact that there is no “Russian RD-180 engine” in nature. There is the RD-180 engine, created as part of Russian-American cooperation, which was developed in Russia by order of the United States, and which is currently produced by the American company Pratt & Whitney at Russian production facilities. Therefore, the very presentation of the material in the Russian media, which writes that “the United States buys engines in Russia” is 100% a complete lie. This is the same as writing that “Apple buys its iPhones in China” only on the grounds that all of their production is concentrated there.

However, let me tell you everything in order, because the story there is very interesting.

At the end of the 50s, the United States had several hundred Atlas ballistic missiles in service. When the Caribbean crisis happened, the Americans considered that these missiles were not effective enough to withstand the Soviet threat, they were removed from service, but were not thrown away or disposed of. According to the concept that was then adopted in the United States, and which is still in effect, all military ballistic missiles should be able to be used as launch vehicles to launch cargo into orbit.

Therefore, with the decommissioning of the Atlases, the US space department received about a hundred ready-made space rockets for launching satellites and spacecraft into space. Moreover, I note - this is very important - that they are actually free, free missiles, since the Pentagon has already paid for them earlier.

Atlases were widely used in the first years of space exploration as the main carrier (it was on the Atlas that the first American astronaut John Glenn took off), and then as a “backup” rocket. When, for example, Challenger exploded, until the causes of the disaster were clarified, the Shuttle program was suspended, and all space launches were made on Atalas.

Meanwhile, in the 90s, it became clear that the Titan rockets, on which all American “medium” launches were made, should be discontinued - the negative consequences of using poisonous aerosol as fuel were too strong.

And there were still hundreds of free Atlases left for conservation. It was decided to equip these Atlases with new, more powerful engines and replace the Titans with them. The American company General Dynamic, which was in charge of the Atlases, announced a tender in 1995 for the development of a new engine, and this tender was unconditionally won by a large margin by the Russian company NPO Energomash, which offered a price several times lower than its competitors.

Times were hard in Russia, dumping had to be done. But most importantly, Energomash had a good start. In order to obtain an engine with the characteristics the Americans needed, it was only necessary to “halve” the existing engine from the Energia rocket, making only two instead of four chambers.

As a result, Energomash “developed” the required engine, which was named RD-180, transferred all rights and all documentation for its production to the Americans, and they, in accordance with the terms of the tender, placed engine production in Russia at Energomash factories, since all the equipment was already there necessary technological equipment.

It should be noted that this contract later backfired heavily on the Russian military-industrial complex, because when Russia itself needed a “half” engine for the Rus-M and Angara missiles, it turned out that, according to the terms of the contract, it could not manufacture the RD-180 for its own purposes, but must purchase it from the American company Pratt & Whitney.

As a result, for Rus-M it was necessary to make an “alternative” development, the RD-180V (which was never completed), and to install not a “half”, but a “quarter” RD-191 engine on the Angara.

Well, as for the American Atlases, the missiles equipped with the RD-180 first received the index R (this is not a “Russian engine”, as they say here, but just another index, coincidentally), and then were completely modernized for the RD-180 . And they received the designation Atlas-5.

So all American Atlas 5s now have a first stage equipped with the Pratt & Whitney RD-180 engine, which is assembled in Russia.

Therefore, when Russia came under sanctions, this production also came under sanctions. At first, it was decided to move the production of the RD-180 from Russia to the United States.

But then Elon Musk showed up with his company SpaceX and said: “I can do it better and cheaper.” We figured it out, it turned out to be much cheaper and it would be better to outsource

In Russia, of course, they would be happy with this situation, but in the United States, more than anything else, they are afraid of market monopolization. All relevant regulatory authorities immediately issued a conclusion that the transfer of the contract to SpaceXv would lead to the formation of an unacceptable monopoly.

But as a result of these discussions, it became clear along the way that there was no longer any reason to transfer the production of the RD-180 to the USA. What was “cheap” in 1995 has now become “expensive”.

The RD-180 is a very good engine, but it is already very outdated; to produce it, it will be necessary to revive technologies that have long been abandoned throughout the world. Science and technology do not stand still, and in the USA itself there are a lot of companies that can do what is required much better, much faster and, most importantly, much cheaper compared to Energomash.

In short, it turned out that the RD-180 was no longer needed.

Therefore, General Dynamic held a new tender, which was won by two American companies. United Launch Services which, starting in 2019, will begin supplying the Vulcan BE-4 engine, which will replace the RD-180. And Aerojet Rocketdyne, which will develop the next generation of fundamentally new engines, which in turn will replace the Vulcan BE-4.

Well, to make it clear what happened, I will mention only one detail - the entire contract with United Launch Services is worth 46 million dollars - this is the cost of only five RD-180s.

And the US Congress, in order to hedge its bets and create a reserve for the transition period, allowed Energomash to produce 18 more RD-180s. The last RD-180 in history.

This is, in fact, what lies behind the Russian media headlines “America cannot do without Russian engines.”

There is a very simple way to understand which opponent writes a comment on your article, when he does it sincerely, due to his own convictions, and when “as part of his official assignment.”

When an opponent is “sincere,” his comment can appear at any moment, it is usually “single,” and it usually contains some original maxims, even if they were picked up a couple of minutes ago on Wikipedia.

But when it’s “within the framework of a job assignment,” the picture will be different. Such comments never appear immediately. After all, some time must pass before this “official assignment” is formed and “methodological instructions” are given on it. In this case, the “commentators” always appear with a delay of half to one and a half days; they appear immediately in a crowd and each repeats the same “arguments” received during the briefing. And everyone likes each other’s comments in a circle. In short, the picture is obvious and does not require any special investigation.

I usually enter into dialogue with opponents of the first type, unless they try to tell me a Wikipedia article. For obvious reasons, I block opponents of the second type as they approach. After which topics inevitably appear somewhere on third-party resources about how Shipilov is afraid to enter into discussions and silences his opponents. But nothing can be done about it, these are the usual costs of living a person with an active lifestyle.

Why am I telling this?

With the article that the famous “Russian RD-180 engines,” which “America cannot do without,” are in fact American engines, albeit produced in Russia and developed in Russia by order of the United States, it seems that I stepped on someone else’s footsteps. that is a very painful callus. After chatting about the topic neither on Facebook nor on my website, it didn’t work, many discussions were created on other sites and social networks, where numerous “experts” argued with references to the “primary sources” of parallel reality they created, they say to the general public that “Shipilov is lying,” “Shipilov is illiterate.” And even the Laughnews channel devoted several stories to defaming the “illiterate Shipilov.”

In short, they were deeply hooked.

I never pay attention to such things. But this is precisely the case when the pinching achieved its goal. In recent days, several seemingly reasonable and adequate friends began to give me advice that if I “lied,” then it would be better for me to repent myself and admit my mistakes, so that my reputation would not suffer.

And I thought, since such powerful counter-propaganda began to blur even the brains of thinking and reasonable people, then what to say about everyone else.

In short, we need to work on the mistakes. Not over my mistakes, of course, which simply do not exist. And over the mistakes of the Kremlin propagandists.

Below is the argument they use and my comments on that argument.

“The fact that all rights to the engine were registered to the American company Pratt & Whitney and that it is their official manufacturer is a purely legal ploy to circumvent export restriction laws.”

If I ask you to describe in detail what specific “export restrictions” this “legal trick” circumvents, you will not be able to do this. Is not it?

And what do “export restrictions” have to do with it if the engines are imported? You also can’t explain?

The fact that the manufacturer of the RD-180 engines is the American company Pratt & Whitney is precisely a fact. And no matter what “justifications” you come up with for this fact, they do not cancel this fact.

“So what if the engine was ordered by Shatami and is being made specifically for the States! It was developed in Russia, is manufactured in Russia, which means it is a Russian, not an American engine."

If you bought potatoes at the market, then these will be your potatoes, and not the one who grew them and sold them to you.

What are you saying? Are potatoes a bad example? Is there a big difference between potatoes and high technology? OK! Here's another example from the field of high technology.

You need a website, you ordered it from a programmer and then hired the same programmer for maintenance and technical support of the site. Whose website will this be? Yours or the programmer you hired?

“The engine was not made specifically for the States from scratch, it was a ready-made, still Soviet engine from Energia, which was simply remade to meet the requirements of the Americans. This means that this is not an American, but a Russian engine."

Yeah, but if the programmer you hired to make a website for you didn’t write the code from scratch, but used his earlier work, does this somehow change your rights to your own website?

“Pratt & Whitney owns the rights to the engine only in the United States, while global rights remain with Russia. So the RD-180 is a Russian engine"

Ah-ah-ah, there you go!

Well then, name me at least one Russian missile that would use this Russian engine.

You can not? Do you know why?

Yes, because now all the key elements of the RD-180 are protected by US patents! Well, offhand, so as not to be unfounded: US Patent 6244041, US Patent 6226980, USPatent 6442931. Moreover, although the “basic basis” of the engine is taken from the Soviet RD-170, all the fine control mechanics and automation: pumps, valves, control circuits - this is all American, real American developments owned by Lockheed and Martin.

Therefore, when Russia needed exactly such an engine as the RD-180 for the Rus-M missiles, it had to start developing a complete Russian analogue - the RD-180V, which would not use American patents and American developments. It was not possible to solve this problem: by that time there were still specialists in engine production in Russia, but there were no specialists left in their development.

“The USA does not have the technology to make engines like the RD-180, but Russia has them”

In general, this is true. But the meaning of this truth is still different.

I believe that the technology for making steam locomotives in the USA has also been lost. But it does not at all follow from this that they do not know how to make diesel and electric locomotives.

The reality is that in Russia over the past thirty years not a single truly new rocket engine has been developed or appeared. All the “newest” Russian engines: RD-181, RD-191, RD-193 - under these names a single chamber from the four-chamber RD-170 engine developed in the 80s is produced. Therefore, all modern Russian missile technologies are from the 80s of the last century.

The USA really doesn’t have such technologies anymore. There, new developments in the field of rocket engines appear every year. There are completely different principles, goals and methods of implementation.

“The USA cannot do without Russian rocket engines, this is a fact”

If by “Russian rocket engines” we mean “American RD-180 engines produced in Russia,” then yes - right here and now - they cannot. They can’t get by with “little bloodshed.”

The “little loss” clause is made here because both the United States and the European Space Agency have enough alternative carriers to replace the Atlas-5, on which the RD-180 is installed. But it will be expensive and wrong.

Therefore, after the sanctions, the United States ordered 20 more RD-180s to create a “backlog” for the transition period until the RD-180s begin to be replaced in the United States. The current state of rocket technology in the United States allows it to be completed within three years from the start of engine development to its launch into mass production.

“And if this engine is so outdated, then why are the states using it and not their modern developments?”

Yes, simply because it does everything that is required of it, performs its tasks perfectly, and most importantly, at the time of the tender it was extremely cheap.

You, too, would probably prefer to buy a Zhiguli rather than a Mitsubishi Pajero to transport potatoes from the dacha. Another thing is that times go by, and in our time the RD-180 is by no means as cheap compared to its analogues as it was in the 90s. So the question of replacing him was already raised; sanctions only pushed this process forward.

Russian RD-180 rocket engines have been a bone of contention between the two United Launch Alliance (ULA) and rival Orbital Sciences. The first does not allow the second to purchase engines for its Antares rockets.

Clash of the Titans

The reason for this was the participation of Orbital Sciences in government tenders. ULA is illegally preventing competitors from purchasing RD-180 engines from the two companies. This is the contractor of RD Amross - SPNPO Energomash - and the American intermediary Pratt & Whitney Rocketdyne. The first produces the required RD-180 liquid-propellant rocket engine. The other one supplies components to the United States.

The only liquid engine, the RD-180, is optimally suited to the tenders announced by the American government. According to experts, the characteristics of these components are ideal for heavy launch vehicles and the needs of NASA.

What is the RD-180 rocket engine?

RD-180 is a two-chamber derivative of the four-chamber RD-170 used on Zenit. The RD-180 closed-cycle liquid rocket engines with afterburning combine the high performance, convenience and reusability of the RD-170 in dimensions to meet the engine requirements of the Atlas V Evolved Expendable Launch Vehicle.

RD-180 - hydraulic motor for actuating the control valve and deflection thrust vector in the gimbal, with pneumatics for actuating the valve and purge system: the thrust frame for load distribution is self-sufficient as part of the engine. The motor at the start uses LOX lead, with afterburning of generator gas and LOX rich gas turbine drive. Thus, established a 10 percent increase in productivity compared to the operational acceleration of US engines and assuming a clean, reusable operation.
Only in the main assembly, the turbo pump and booster pump required development to the scale of the RD-120 and RD-170. All other components were taken directly from the RD-170.

The RD-180 was developed in 42 months at a fraction of the cost of a typical US new engine design. The motor runs on an intermediate Atlas III and a standard Atlas V launch vehicle.

RD-180 is equipped with two pairs of combustion chambers and nozzles. The engine is developed and produced by the Russian research and production association Energomash. Kerosene is used as fuel and liquid oxygen is the oxidizing agent. The cost of the RD-180 rocket engine in 2010 was $9 million.

Description of design

1. LOX/Kerosene
2. Two thrust chambers (gimbals +8 degrees).
3. One oxygen-rich block in front of the burner.
4. TNA high pressure assembly.
5. Two stage fuel pump.
6. Single stage oxygen pump.
7. Single turbine.
8. Self-igniting ignition.
9. Autonomous (valves, TVC) powered by kerosene from the fuel pump.
10. Health monitoring and life forecasting system.
11. Automated flight preparation (once installed on the vehicle, all operations are automated through launch).
12. Minimization of interface from launch pad and vehicles (pneumatic and hydraulic systems, autonomous, electrical integrated panels, thrust frame to simplify mechanical interface).
13. Environmentally friendly operations with enriched oxidizer triggering preburner ignition, as well as oxidizer startup and shutdown modes that eliminate coking and unburned potentially contaminated kerosene.
14. 50-100% continuous throttling subject to potential real-time testing of trajectory and motor matching on site prior to triggering lock.
15. 80% RD-170 parts.
16. Pressure chamber: 256.6 bar.
17. Area ratio: 36.4.
18. Thrust-to-weight ratio: 77.26.
19. Oxidizer-fuel ratio: 2.72.

Engine RD-180. Characteristics

• Specific gravity: 5,480 kg (12,080 lb).
• Height: 3.6 m (11.67 ft).
• Diameter: 3.2 m (10.33 ft).
• Specific impulse: 337.8 s.
• at sea level: 311.3 s.
• Recording time: 270 s.
• First launch: 2000

Story

In November 1996, the first test of the RD-180 was carried out at the Energomash production association. The engine was recognized as the winner in the tenders for installation in the Atlas launch vehicle of the American Corporation. This was necessary for the launch of promising manned spacecraft. It was from then on that RD-180 engines became the most popular.

The engine is reusable. Thoughtful management provided NPO Energomash with almost legendary reliable transactions with the United States. In December 2012, a contract was delivered providing the company with a guarantee for the production of engines until 2019. All production is concentrated in Russia.

Creation of a replacement for the RD-180 in the USA

The Ukrainian events led to sanctions limiting the ability of the United States to use Russian rocket engines. The RD-180 must be replaced with an American-made analogue. In December 2014, the House adopted an amendment. It prohibited the use of Russian RD-180s. The engine will continue to be purchased under the existing supply agreement until 2019 between Energomash and ULA.

Despite the continuation of cooperation and deliveries of the RD-180 under existing agreements, the US Secretary of Defense ordered the termination of cooperation with Russia and the transition to American components. America is obliged to get rid of Russian dependence in the military-political spheres.

To this, Frank Kendall (Secretary of Defense for Procurement) responded that the Pentagon has nothing to replace the Russian RD-180 engines with. As an alternative to the current situation, America announced a tender for the production of its own engines with similar characteristics on its territory.

Russian Deputy Prime Minister Dmitry Rogozin said he was ready to stop the supply of RD-180 and K-33 rocket engines to America.

How much does the RD-180 rocket engine cost for the USA?

Let's talk about prices. SpaceNews reported that the RD-180 engine needs to be replaced. In the US, such a whim will cost $1.5 billion. Not a small amount.

How much does the RD-180 engine cost? The entire project to implement the prototype will last at least six years. According to experts, the United States does not have the opportunity to completely abandon the use of RD-180 engines. It is impossible to solve the problem that has arisen in a short time, since the motors will only be ready in 2022.
Despite assurances from the US Air Force that RD-180s are in stock in the required quantities, there is still a shortage. Therefore, many launches will need to be postponed. Spending in this area could increase to $5 billion.
While the United States is competing and applying sanctions, China is already in line to produce the RD-180.

Perspective

The Pentagon has awarded at least $162 million to Aerojet Rocketdyne and the United Launch Alliance to work on the development of the AR1 and BE-4 rocket engines, candidates to replace the Russian-made engine that currently powers the Atlas V rocket.

The U.S. Air Force is completing initial investments in new rocket engines as the military seeks to move away from its dependence on the Russian RD-180 engine used on the Atlas V, which launches most of the U.S. government's satellites for secure communications, navigation and intelligence-gathering systems.
The Air Force is part of a public-private partnership with Aerojet Rocketdyne and ULA, providing corporate funds to co-finance engine development.

ULA's President and CEO continue to meet to achieve the goal of providing the most reliable launch systems for the United States at the most affordable price, while developing a new engine that will enable entirely new capabilities for space applications.

The agreement with Aerojet Rocketdyne covers the development and testing of the AR1 rocket engine. This is a power unit that burns a kerosene mixture and liquid oxygen. These are the same propellant components found in the RD-180 engine for the Atlas V.

Aerojet Rocketdyne aims to have the engine certified airworthy by 2019, but first launch is not expected until 2020.
The Air Force is committing a minimum of $115.3 million to the AR1 development program, while Aerojet Rocketdyne and ULA are jointly investing $57.7 million, Aerojet Rocketdyne said in a statement.

Before testing, the government's decision to continue supporting the AR1 engine program has a maximum value of $804 million - with $536 million from the Air Force and $236 million from Aerojet Rocketdyne and ULA.

“The AR1 will return the United States to the forefront of kerosene nuclear rocket engine production,” Drake said in a press release. “We are introducing the latest advances in modern manufacturing while leveraging our wealth of knowledge in producing the next generation of rocket engines to deliver engines that will end our dependence on a foreign supplier to launch our nation's national security assets.”

The AR1 engine will include 3D printed parts and run on enriched oxygen with afterburning generator gas. This is a more efficient engine cycle than other liquid hydrocarbons currently used in US rocket engines.

The BE-4 engine is a major focus for the Air Force. Cash injections are allocated for its implementation. The Air Force is committing to pay at least $46.6 million to the United Launch Alliance for the next generation of the Vulcan rocket. ULA also agreed to add $40.8 million under the terms of the government award.

The lion's share of the initial funding - $45,800,000 - will go towards developing the BE-4 engine, which will generate 550,000 pounds of thrust and consume a cryogenic combination of liquefied natural gas and liquid oxygen.

Two BE-4 engines will boost the first stage of the Vulcan rocket. Officials say BE-4 is fully funded by the company with help from the United Launch Alliance. The Air Force funding will support the company's progress integrating the BE-4 engine with the Vulcan launch vehicle.

Aerojet Rocketdyne touts the AR1 as the easiest replacement for the RD-180 due to its propellant mixture and size. Two AR1 engines are required to meet the performance of the Atlas V's single dual-nozzle RD-180 engine.

ULA executives say the BE-4 engine from Blue Origin, the entrepreneurial space firm founded by Amazon.com, will be ready faster and ultimately be easier to refurbish for reuse.

While the RD-180 engine had the advantage of more than 60 successful launches, the time had come for American investment in domestic production of similar engines.

The BE-4 is scheduled to complete its certification in 2017, and ULA is targeting the first flight of the Vulcan rocket by the end of 2019.

The Air Force also funds construction in outer space for astronaut habitat for deep space exploration and satellite services.

ULA continues to work with both Blue Origin and Aerojet Rocketdyne. It accompanies two options for the next generation of American engines, which is why the company is teaming up with two of the world's leading space companies.

ULA keeps the AR1 engine from Aerojet Rocketdyne as a backup option. The final selection is expected at the end of 2016.

The Air Force's financial commitments to Aerojet Rocketdyne and ULA opened on February 29, 2016, following similar agreements with SpaceX and Orbital ATK.
A new solid rocket booster project made by Orbital ATK for ULA's Vulcan rocket and for its own launcher will also receive Orbital ATK funding.

Shadow in space from earthly “clouds”

Russian RD-180 engines have no alternative in the States. Aerojet Rocketdyne Vice President Jim Meiser believes that the United States is not paying enough attention to the development of its own oxygen-kerosene prototypes.

He said America is definitely behind the Russians and Chinese in creating such propulsion systems. He also mentioned that the United States has already developed an oxygen-kerosene engine, which is in operation Merlin 1D. It is produced by SpaceX. Only, in terms of its characteristics, it does not reach the RD-180.

Of course, this is complete nonsense, because terrestrial clouds cannot cast any shadow into space. But in a political sense, alas, they are discarding it.

USA: industrialists are calm, politicians are worried

A senior U.S. Air Force official said he would stop launching national security satellites aboard United Launch Alliance's Atlas V rocket if the Treasury Department believes importing a Russian engine does not violate U.S. sanctions.

Earlier, Senator John McCain asked the Air Force to prove that Russia's recent reorganization of its rocket and space industry does not make the purchase of RD-180 engines a violation of US sanctions that were imposed against Russian officials in 2014.

US government agencies, led by the Department of the Treasury, are taking a fresh look at RD-180 deliveries. And they are ready not to adhere to sanctions. Grounding the Atlas V would create a bigger obstacle for the Pentagon than fighting.

McCain held a military spaceport hearing where he called on the Air Force to obtain a fresh legal opinion that imports of the RD-18O violated U.S. sanctions imposed on Russian officials in the wake of Ukraine's annexation of the Crimean Peninsula.

McCain singled out two senior Russian officials: Russian Deputy Prime Minister Dmitry Rogozin and Sergei Chemezov, an adviser to Russian President Vladimir Putin. Until recently, they were observers in the space sector. Although they do not benefit financially from the sales of the RD-180, they were subject to sanctions.

On December 28, by order of Putin, the Russian space sector will be reorganized. This restructuring brings adjustments to the Russian space industry and the space agency Roscosmos under a new state corporation, also called Roscosmos.

McCain noted that this organization is currently headed by Rogozin; Chemezov also has something to do with this. Rogozin and Chemezov were among the first Russian officials to receive sanctions during the Ukrainian crisis. Neither can enter the United States. The assets they own were frozen.