Rd 191 and 180 comparison. © State Corporation for Space Activities "Roscosmos"

ANALYSIS OF THE EFFECTIVENESS OF NOZZLE EXTENSION FOR ROCKET ENGINE RD-191

Marat Seydagaliev

Russia, Baikonur

Nikolay Ilyushenko

5th year student of the department “Design and testing of aircraft” branch “Voskhod” of MAI,

Russia, Baikonur

Olga Shestopalova

candidate of science, assistant professor of branch “Voskhod”

of the Moscow aviation institute (national research university),

Russia, Baikonur

ANNOTATION

Modern rocket engines have almost reached the limit of fuel energy capabilities, so increasing the efficiency of a rocket engine even by small values ​​is not an easy task. The paper proposes a solution to this problem by using a sliding nozzle nozzle. For calculations, the most efficient and promising single-chamber liquid-propellant rocket engine RD-191 was taken as an example.

ABSTRACT

Modern rocket engines have almost reached the limit of energy fuel capabilities so increasing the efficiency of rocket engine even for small values ​​is a big problem. There is a solution which suggests to use the nozzle extension. As an example for the calculations was taken RD-191 – the most effective and perspective liquid propellant rocket engine by now.

Keywords: launch vehicle (LV), propulsion system (PS), nozzle nozzles, liquid-propellant rocket engine (LRE), jet thrust, specific impulse.

keywords: launch vehicle, nozzle extension, liquid propellant rocket engine, jet thrust, specific impulse.

To date, the most promising launch vehicle national cosmonautics is the Angara family of launch vehicles, which are based on the universal rocket module - 1 (URM-1). The propulsion system of the URM-1 is the RD-191 liquid-propellant rocket engine. This paper evaluates the effectiveness of using a nozzle nozzle for the RD-191 engine. Nozzle nozzle - the retractable part of the rocket engine nozzle, installation in working position which provides an increase in the output area of ​​the nozzle, as a result, increases the efficiency in rarefied layers of the atmosphere or in vacuum.

The following assumptions were made in the calculation:

• the engine is running normally (with constant mass flow);
• the launch vehicle flies in a straight line, at a constant speed;
• losses due to friction and velocity dissipation at the nozzle outlet are not taken into account.

Required for calculation specifications LRE RD-191 are presented in table 1.

Table 1 .

Characteristics of the RD-191 rocket engine

Characteristic	Designation	Meaning
Thrust (Earth), tf
Thrust (emptiness), tf
Specific impulse (Earth), s
Specific impulse (void), s
Pressure in the combustion chamber, kgf / cm in sq.
Pressure at the nozzle exit, kgf/cm2 in sq.
Temperature in the combustion chamber
Nozzle expansion ratio
Nozzle outlet diameter, mm
Diameter of the minimum section of the nozzle, mm

For calculations, it is proposed to use the jet engine thrust formula under the assumption of one-dimensionality of the gas flow through the nozzle:

where: µ – second mass flow rate; are the pressure, velocity, and cross-sectional area at the nozzle exit, respectively; - pressure environment, (depends on the lift height h).

The flow velocity at the nozzle exit is determined by the relationship known from gas dynamics:

(2)

where: - gas constant of combustion products; - temperature pressure in the combustion chamber, respectively; is the adiabatic index.

The adiabatic index depends on the fuel components used, for a kerosene-oxygen pair; =1.11 .

From expressions (1) and (2) we obtain the final expression for calculating the thrust of a jet engine:

(3)

Obviously, the thrust of the engine changes as it rises to altitude. The reason for this is that the ambient pressure is a continuously changing quantity.

Equation (3) describes the thrust of an engine with a constant geometric expansion ratio. Consider the case in which at each moment of time the calculated mode of operation of the engine () is realized. Then equation (3) will take the form:

(4)

To calculate the average thrust of an engine using a sliding nozzle, it is necessary to determine the geometric characteristics of the nozzle nozzle. Calculations have shown that the optimal radius of the nozzle nozzle, at which the average thrust will be the greatest throughout the entire operation of the engine, exceeds the radius of URM-1 (1.45 m), based on this, we accept the radius of the sliding nozzle equal to 1.20 m, this will allow using nozzle nozzles in a package design and layout scheme (Angara-A3, Angara-A5, Angara-A5V). Based on the nozzle radius, we determine the pressure at the nozzle exit and calculate the engine thrust according to equation (1).

Below are the results of calculations (Fig. 1) of engine thrust according to equations (3), (4) for three cases:

• engine with fixed nozzle;
• engine with perfectly height-adjustable nozzle;
• engine with a single-stage adjustable nozzle.

Figure 1. Change in engine thrust depending on the flight altitude: 1 - unregulated nozzle, 2 - single-stage adjustable nozzle; 3 - perfectly height-adjustable nozzle

The calculation results showed that the use of a nozzle attachment for the Angara launch vehicles, made in a batch scheme, allows increasing the average thrust of each URM-1 by 9.28 tf, taking into account losses due to friction in the nozzle. When using a sliding nozzle nozzle on light-class launch vehicles made in a tandem scheme (Angara 1.1 and 1.2), the thrust increase will be 17.5 tf due to the absence of a restriction on the radius of the nozzle nozzle. When making structural changes to the RD-191 nozzle (in order to increase the pressure at the nozzle exit), it seems possible to increase the thrust by 24.4 tf for the stack and 35.7 tf for the tandem scheme.

Adjusting the height of the nozzle by using a nozzle nozzle is not a fundamentally new engineering solution, but a practical implementation this decision never found because of the difficulty of providing nozzle cooling. Today, this problem is resolvable due to the emergence of fundamentally new materials that were not available before, having a high swimming temperature, strength, wear resistance, etc. That is why the presented work is relevant and practically realizable.

Bibliography:

1. Alemasov V.E. Theory rocket engines: studies. for universities. - M .: Mashinostroenie, 1980.

2. Grechukh L.I. Designing a liquid rocket engine: guidelines to course and diploma design. - M.: Publishing house of OmGTU, 2011. - 69 p.

3. Dobrovolsky M.V. Liquid rocket engines: textbook. for universities. -M.: MSTU named after N.E. Bauman, 2006. - 269 p.

4. Propulsion system. RD-191 - [ Electronic resource]. - Access mode. – URL: http://ecoruspace.me/%D0%A0%D0%94-191.html (Date of access: 04/08/16).

It is quite obvious that the development of marching propulsion systems for launch vehicles is inextricably linked, and especially in the long term, with the improvement of the launch vehicles themselves. In general, we can say that everything will be determined by the goals of world space activities. According to experts, in 2000-2010. expected: - almost 1000 launches of launch vehicles (LV) of various classes, including about 20% for launch spacecraft(SC) to geostationary orbits (GSO); - every second of the 2000 spacecraft being withdrawn will be commercial; - the cost of spacecraft launched annually will be about $4-5 billion. In addition, the implementation of the large-scale international project ISS Alfa worth tens of billions of dollars will be continued. A manned expedition to Mars, the creation and operation of a base on the Moon, the energy supply of the Earth from space, the fight against the meteorite hazard, the removal of especially hazardous waste and space tourism are projects of the not too distant future. It is noteworthy that the number of countries that became owners of spacecraft for the first time has doubled over the past 15 years (from 15 to 30).

The further development of world space activity is constrained by the high cost of launching a spacecraft ($5,000...10,000 per kilogram for launching into a low circular orbit) and insufficient reliability of launch vehicles. Thus, every 20…30th flight is an emergency one, and in 50% of cases it is due to the fault of propulsion systems (PS). The cost of one heavy-class launch vehicle accident, including the loss of a spacecraft, is $300...700 million, which exceeds the cost of developing a powerful LRE (200...250 tf thrust). Economic losses, for example, as a result of the Space Shuttle disaster, exceeded $2 billion. In addition, accidents lead to a delay in the implementation of programs by up to one and a half to two years and a decrease in competitiveness.

Thus, the priority requirements for advanced launch vehicles (SV) are to increase their reliability and reduce the cost of spacecraft launch.

As shown by studies conducted in Russian research institutes and design bureaus, the main type of engine for promising SVs for the next 20-25 years will remain liquid propellant rocket engines. Other propulsion systems, for example, hypersonic ramjet engines (scramjet engines), which use atmospheric air as an oxidizing agent and promise a significant reduction in launch weight, require the solution of a number of complex problems. These are problems associated primarily with the development of designs of remote control and aircraft in general, operating under conditions of high velocity pressure and aerodynamic heating (1500 K and more). These problems push the implementation of the scramjet to a more distant future.

Currently, active purposeful work is being carried out abroad to create new disposable launch systems (Arian-5, the Delta-4, Atlas-5 and H-2A launch vehicle families) based on liquid-propellant rocket engines. The salient features of most of them are:

• - creation of central stage propulsion systems on new LREs using highly efficient oxygen-hydrogen fuel, with special attention being paid to reducing the cost and increasing the reliability of LREs (IHPRT program in the USA). As a rule, the composition of the control unit includes one high-thrust engine (RS-68 with a thrust of 294 tf; RS-76 with a thrust of 373 tf; LRE for VA-1 with a thrust of 635 tf);
• - widespread use of cheap and reliable booster rocket engines solid fuel(RDTT), the number of which varies from 0 to 6, which allows minimal cost get a family of carriers of different carrying capacity;
• - formation of heavy launch vehicles from two or three central blocks.

Recognized worldwide the highest level Russian rocket engine building. This is confirmed by the NPO Energomash developed in 1975-1985. the RD-170 engine, which runs on oxygen-kerosene fuel and has no equal in the world in terms of the achieved parameters and energy-mass characteristics. It is not for nothing that foreign firms have intensified their activity in using Russian engines on modifications of US launch vehicles. Thus, the RD-180 engine, developed by NPO Energomash and being further development RD-170 is intended for use on the Atlas-2AR launch vehicle manufactured by Lockheed Martin. The use of RD-180 will significantly increase the energy capabilities of the carrier. In the United States, it is also planned to use the NK-33 and NK-43 engines, which were developed in the early 1970s. for the Soviet lunar rocket H1. After concluding an agreement with the Aerojet corporation, these rocket engines are being finalized for subsequent installation on the K-1 reusable launch vehicle of the Kistler Aerospace company. Widespread use of these cheap (at world prices) and highly efficient liquid-propellant rocket engines, created in Russia, will significantly reduce the cost of spacecraft launch.

Much attention is paid to reducing the cost of manufacturing launch vehicle stages, the cost of preparing and conducting launches. As a result, an approximately 1.5-2-fold reduction in the cost of launching and an increase in the reliability of foreign launch vehicles to the level of such Russian missiles, like Soyuz and Proton (Fig. 1).

In a somewhat more distant future, it is envisaged to replace booster solid propellant rocket engines with reusable rocket-propelled rocket boosters, as well as reusable one- and two-stage systems (Venture Star, etc.). Their use should reduce the cost of launching another 5-10 times.

characteristic feature A similar way of developing disposable launch systems is to increase the number of required zones for the fall of spent stages. Each of the options with additional boosters results in two additional oversized zones for the booster and first stages to drop. As a result, for a family of launch vehicles based on a two-stage carrier, instead of one zone, two to six zones are required, depending on the number of booster accelerators.

With a coastal location of the launch complex, which is typical for foreign spaceports, this does not matter; for the intracontinental location of Russian spaceports, this is practically unacceptable, especially if we take into account the requirements of launching LVs at different azimuths.

As for the high-latitude (62.8 °) Russian Plesetsk cosmodrome, then (with equal energy parameters of the LRE, which is becoming characteristic of the current stage), in order to launch spacecraft of the same mass into geostationary orbit (GSO), it is required to increase the power of the propulsion systems of domestic SVs by 30 ...40% compared to foreign ones, located mainly near the equator. Previously, this unfavorable factor was parried by a significantly higher efficiency of domestic liquid-propellant rocket engines (RD-170, etc.) compared to foreign engines (the specific impulse is higher by 30…35 s). However, the widespread use of oxygen-hydrogen fuel in modern foreign launch vehicles (Ariane-5, Delta-4, H-2A) and its absence in domestic projects significantly worsened the comparative picture.

Thus, for promising domestic single-use SVs, especially medium and heavy classes, in case of refusal to use booster solid propellant rocket engines and to parry an unfavorable geographical factor, it is required to develop sustainer propulsion systems with significantly higher thrust or use several engines in propulsion, i.e. transition to the use of multi-engine units based on modular rocket engines.

Based on the above, the Keldysh Center and TsNIIMash proposed a "Concept for the development of a launch vehicle system Russian Federation for the period after 2005."

The "Concept" is based on the following basic principles:

• - unconditional provision of guaranteed and independent access to outer space from the territory of the Russian Federation;
• - ensuring, in the long term, the high competitiveness of domestic satellites in the world market of space services.

The decisive step in this case is the development and subsequent widespread use of two-stage carriers with the first reusable winged stage (Fig. 2), which can provide:

• - reduction of costs for removal by ~2 times;
• - an almost complete solution to the problems with the allocation of zones for the fall of spent stages and the removal of severe restrictions on flight routes, which will make it possible to transfer carrier launches from the Plesetsk cosmodrome to Kapustin Yar and at the same time provide a 15 ... 20 percent increase in their energy capabilities.

It should be noted that the creation of reusable first stages does not require the solution of new scientific and technical problems and can be successfully solved at the current level of development of domestic aviation and rocket technology.

The groundwork accumulated during the development of the Buran orbiter, the development of returnable winged stages in aviation design bureaus, the MMKS system at RSC Energia, where the modified Buran OK was considered as the first reusable stage, as well as the latest developments of the GKNPTs im. M.V. Khrunichev on the Angara launch vehicle of a light class with a reusable first stage showed the reality of solving the problem.

The determining link in this case will be the creation of a reusable and reliable propulsion system based on a liquid-propellant rocket engine, and this task cannot yet be considered solved. The only reusable rocket engine SSME of the Space Shuttle system in operation in the world is far from fulfilling the requirements of the TOR in terms of service life (almost 10 times) and the cost of inter-flight maintenance. It is not for nothing that the IHPRT program in the United States provides for the creation of a demonstration sample of an oxygen-hydrogen rocket engine with a multiplicity of use up to 100 times and a tenfold reduction in maintenance costs while simultaneously reducing the cost of development and manufacture (Fig. 3).

The transition to a reusable first stage will lead to an increase in the launch mass of the launch vehicle by ~30%, which will require an increase in the propulsion thrust of this stage. A transition to a multi-engine plant is required. Thus, the development of a redundant multi-engine plant using a reusable liquid-propellant rocket engine should be considered the most important task of the domestic rocket and space engine building at the present stage. The requirements that must be met by such a control system include the following:

• - the failure of one engine should not lead to the disruption of the flight program;
• - the frequency of use of remote control at the first stage should be 10-15, in the subsequent - 50-100;
• - the cost of inter-flight maintenance of PS should not exceed 3% of the cost of PS, with a subsequent reduction to 0.5% or less.

One of the possible ways to solve this problem is the development of a new generation of liquid-propellant rocket engines according to the scheme with a reducing gas generator. This scheme is characterized by a fairly high probability of non-intensive development of emergency processes (the development period of which exceeds 0.1 ... 0.5 s). In such accidents, as a rule, there is no external destruction of the gas path (Table 1). All this can provide efficient work emergency protection systems with a simultaneous increase in the coverage factor of emergency situations up to 0.9 ... 0.95. The way is opened for the creation and successful operation of redundant PS, which is confirmed, in particular, by the experience of operating the Saturn-V launch vehicle.

The specified feature of a liquid-propellant rocket engine with a reduction gas generator, especially in combination with the use of an open engine scheme with the release of gas generator gas or bypassing it into a nozzle, is especially important for the development of new carriers intended for delivering crews to an international space station and the launch of promising manned vehicles for various purposes.

Emergency processes that develop at a high intensity and have an explosive nature (t = 0.001…0.002 s) practically completely exclude the possibility of rescuing cosmonauts, since it is impossible to carry out emergency separation of the compartment with the crew under these conditions.

Currently, there are only two carriers in the world that provide the launching of crews into space: this is the domestic Soyuz launch vehicle and the American Space Shuttle. The sluggish development of emergency processes in the engines of the Soyuz launch vehicle, which are characterized by a low intensity of parameters, an open engine circuit and the use of a gas generator with a reducing generator gas, made it possible to implement an effective crew emergency rescue system, which was repeatedly confirmed during the 30-year operation of this launch vehicle. and its prototypes. The explosive nature of the accident of the Space Shuttle carrier with the Challenger ship led to the death of the entire crew.

The use of LRE with a recovery scheme of gas generation can significantly reduce the severity of the problem of ignition of structural materials in a generator gas environment with a high oxidizing potential. This creates the prerequisites for abandoning the use of more expensive structural materials and technological processes and opens up the possibility of reducing the launch cost by 10...15%, despite the reduction in the energy parameters of the LRE and LV due to the transition to a less energy efficient scheme of the LRE.

As calculation-theoretical, experimental and design studies show, the service life of turbomachines to a decisive extent depends on the level of their energy intensity. Therefore, the high power density of the units, primarily HPP of the modern most energy-efficient LRE RD-170, RD-180, RD-191, casts doubt on the possibility of achieving a high multiplicity (up to 25-30) of using such engines and low cost (less than 1 ... 2 % of manufacturing cost) inter-flight engine maintenance. This is evidenced by the operating experience of the world's only reusable rocket engine SSME. The most important role in limiting the frequency of use of LRE is played by cyclic fatigue of the material. It is known that under high-cycle loading, the limiting number of cycles (respectively, the operating time) of a structural element depends, in particular, on the level of dynamic stresses in a power law (Weller equation). Therefore, reducing the energy intensity by a factor of 2 makes it possible, in principle, to increase the duration of the LRE by more than an order of magnitude (Fig. 4).

In table. 2 shows that with the transition to an open LRE scheme with a pressure level in the combustion chamber of 140 ... 150 kgf / cm2, it becomes possible to reduce the pressure behind the pumps and the required turbine power by 2 ... 2.5 times compared to the parameters of the RD-191 engine from the RD family -170. those. to create a rocket engine with a very high service life and a multiplicity of use up to 30 ... 40. In combination with the use of cryogenic propellant components (liquid oxygen and liquid methane), which create conditions for minimal inter-flight maintenance of LRE, it becomes possible to reduce costs (in the line of propulsion) for one flight by 20 ... 30 times (see Fig. 3).

It should be noted that when developing engines for promising reusable launch vehicles, US developers are following almost the same path (reducing the tension level of units and creating a reusable rocket engine with a bulkhead after 30-40 flights).

The results of computational-theoretical and experimental studies carried out by the Keldysh Center taking into account experience design developments KBKhA and NPO Energomash, led to the conclusion that the task with the greatest effect can be solved by developing a new generation of liquid-propellant rocket engines using the fuel pair "oxygen and liquefied natural gas" (LNG), and LNG should contain 98% methane.

The use of this fuel pair provides:

• - the possibility of developing a highly efficient liquid-propellant rocket engine according to the scheme with a reducing gas generator;
• - creation of reusable engines with a minimum amount of inter-flight maintenance.

The pair "oxygen and LNG" has a low cost and wide prospects for use in other industries (aviation, railway and road transport). Although in Russia there is still practically no infrastructure for the use of liquefied natural gas, the existing practice of operating cryogenic components (oxygen, hydrogen), as well as rich world experience in the production and transportation of LNG, allow us to conclude that it is possible to create the necessary infrastructure at relatively low costs.

The most expedient scheme of a propulsion rocket engine for launch vehicles of a new generation is an open, open circuit with a reducing generator gas (Fig. 5). To reduce the losses of the specific impulse of thrust, it is advisable to apply the bypass of the exhaust gas generator into the nozzle. As a result of the implementation of research, theoretical and experimental work, including on specially designed model engines, recommendations were received on the organization of the working process in the gas generator and combustion chamber, the possibility of achieving a high degree of perfection of processes and effective cooling of the combustion chamber, long-term resource of work and repeated use. In general, the full reality of creating a highly efficient liquid-propellant rocket engine of a new generation and the possibility of moving to full-scale design developments are shown.

The creation of a redundant first-stage PS based on highly reliable liquid-propellant rocket engines will ensure a guaranteed and cost-effective launch of both manned objects and unique expensive large-mass spacecraft.

In conclusion, it should be noted that the use of the main provisions of the "Concept" developed by the Keldysh Center opens up prospects for creating a new generation of sustainer reusable engines that provide:

• - high reliability;
• - ease of inter-flight maintenance and multiple use;
• - formation of multi-engine redundant PS.

On the basis of such rocket engines (Table 3), new, environmentally friendly, not requiring exclusion zones, reliable and cost-effective launch vehicles with the first reusable stage, which reduce the cost of spacecraft launch by almost half, can be developed.

Some characteristics of the rocket engine
Characteristic	Closed circuit with oxidative HG		Closed circuit with recovery GG
Fuel components	O2 +RG-1		O2 +CH4
Producer gas composition	O2 - 91% H2 O - 4% CO2 - 5%	O2 - 6% NO2 - 73% N2 - 6% H2O - 4% N2 O4 - 2% CO2 - 6% HNO3 - 4%	O2 - 0% CH4 - 55% H2O - 6% H2 - 24% CO2 - 3% CO - 12%
Oxidation potential	O2 - is in a free state	O2 - is in a bound state	O2 - absent
Tank cleanliness requirements	0.05...0.1 mg/m2	5.0...7.0 mg/m2	3.0...5.2 mg/m2
TNT equivalent
The time of the accident of the gas path until the loss of tightness, s	<0,06 (~40% аварий)		0.1 (without opening the gas path)
The speed of the automatic control system for cutting off fuel lines	**0.8...0.1 s,	**0.8...0.1 s, in perspective - 0.06...0.08 s	**0.8...0.1 s, later - 0.06...0.08 s
BACS coverage rate
Consequences of accidents (after shutdown of the SAZ fuel lines)	SAZ does not work, destruction of the remote control, compartment and block (in the presence of initiators)	Destruction of the engine, burning of the compartment structure and attenuation of the process

LIQUID ROCKET ENGINE RD-191

14.06.2016

The Russian NPO Energomash plans to double the production of RD-191 engines for Angara launch vehicles in 2017, General Director of the enterprise Igor Arbuzov said.
“The testing phase of the Angara launch vehicle has begun, the number of orders for the RD-191 has increased. Consequently, NPO Energomash should double its production volumes (in 2016 - 22 engines, in 2017 - 40), ”the Energomash NPO corporate publication quotes him as saying.
According to him, in order to fulfill the order, the company will have to increase the number of workers in production by 250-300 people.
TASS

16.04.2019
The problems with low-frequency vibrations of the RD191 engines for the Russian Angara launch vehicle have been resolved, Petr Levochkin, chief designer of NPO Energomash, said in an interview with Interfax.
“We have introduced a number of solutions to the engine design to suppress these low-frequency vibrations and achieved with the Khrunichev Center that these measures allow the engine to operate normally and meet the technical requirements,” said P. Levochkin.
So he commented on the media reports that appeared in January of this year that the vibrations of the RD191 engine that occur during the launch of the Angara rocket can lead to its destruction.
The interlocutor of the agency explained that low-frequency vibrations arose due to an extremely difficult regime for the power plant of the rocket, when the engine of the central block of the first stage operates at only 30% power to save fuel.
“RD191 is unique. On the Angara-A5, the engine of the central block, while the sides are working, should work in a sparing mode, saving fuel. A 30% deep throttling mode was chosen for this rocket,” P.Levochkin said.
Interfax-AVN

LIQUID ROCKET ENGINE RD-191

The development of the RD-191 engine began at the end of 1998. This engine with afterburning of oxidizing gas is designed for the Angara family of domestic launch vehicles. The design of the engine is based on the design of the RD-170/171 engines.
RD-191 is a single-chamber liquid-propellant rocket engine with a vertically located turbopump unit. During 1999, design documentation was issued, in 2000, autonomous testing of the RD-191 engine units began and pre-production was completed. In May 2001, the first finishing engine was assembled. The first firing test of the RD-191 was carried out in July 2001.
As of June 2011, 120 fire tests of the engine were carried out with a total operating time of 26892.4 seconds, including three fire tests of the RD-191 as part of the URM-1 (the first stage module of the Angara launch vehicle) were successfully carried out in the summer-autumn of 2009 ) at the Research Center of the Russian Communist Party (Peresvet, Moscow Region).

CHARACTERISTICS

Liquid propellant rocket engine with afterburning of oxidizing gas
Fuel - oxygen + kerosene
Traction, ground/empty, tf 196/212.6
Specific impulse, earth / void, s 311.2 / 337.5
Pressure in the combustion chamber, kgf/cm2 262.6
Weight, dry/flooded, kg 2290/2520
Dimensions, height/diameter, mm 3780/2100
Development period 1999–2011
Purpose For the first stage of the Angara launch vehicle family

2019-07-23. The new site of the Perm production will increase the efficiency of the production of rocket and space products.
In July, at the suburban site of Proton-PM PJSC (part of the integrated structure of NPO Energomash JSC), as part of the reconstruction and technical re-equipment of the enterprise, a sheet cutting and painting section was organized. The amount of investments in the creation of production amounted to more than 76 million rubles.
The new site produces ground-based products: parts and assembly units of gas turbine power plants of the Ural series, as well as equipment. In the near future, the site will be involved in the production of combustion chambers for rocket engines and other space-related nomenclature.
Earlier, the governor of the Kama region, Maxim Reshetnikov, noted that the production of rocket engines is the height of scientific and technological progress and an important factor in the development of the region. According to the head of the region, the Permian rocket and engine building enterprises enjoy great confidence in the country's leadership, and the quality of products is assessed as very high. Everyone understands that Perm enterprises are a guarantee of reliability.
Dmitry Shchenyatsky, Executive Director of PJSC Proton-PM, noted that the creation of a cutting section is the next stage in the organization of a modern full-cycle blank production at the company's country site in Novye Lyady. “This is a step forward that will allow us to optimize the production process, use new capacities in the development of promising rocket and space products and the transition to its serial production. Next year, we plan to ensure 100% utilization of the commissioned equipment,” the top manager emphasized.
At the site of sheet cutting and painting with a total area of ​​​​more than 2 thousand square meters. m housed four units of modern technological equipment: a laser cutting unit and a water jet cutting unit for cutting sheet material, a shot blasting chamber for preparing metal for coating, and a painting and drying chamber. In addition, guillotine shears for cutting and chopping metal are installed on the territory of the site, and a warehouse for sheet material is also located here.
The technical characteristics of the laser machine make it possible to cut a contour detail up to 12 mm thick in just one and a half minutes. In turn, the waterjet cutting unit is capable of cutting various materials with a thickness of up to 300 mm with a jet of water, making a cut at an angle, providing the necessary accuracy and cleanliness of the machined surface. It does not use harmful oils, liquids and gases, which increases productivity and safety.
New procurement facilities are being created as part of the organization in the Perm Territory of a production complex for serial production of RD-191 and other promising liquid engines. This project has the status of a priority regional investment project and includes the reconstruction and optimization of the production facilities of PJSC Proton-PM with their concentration on the territory of Novye Lyady, the development by the enterprise of a full cycle of production of RD-191 engine units in the Perm Territory and other new equipment, the creation quality social, educational and housing infrastructure. The total investment will amount to 10.8 billion rubles, while about 250 jobs will be created. The project started in 2018 and runs until 2025.

PERM, August 27 - RIA Novosti. The head of the Roscosmos state corporation, Dmitry Rogozin, announced his intention to open the production of environmentally friendly RD-191 engines for Angara rockets in the Perm Territory, according to the website of the governor and the government of the region.

Rogozin's statement was made on Tuesday during a working meeting with Governor of the Perm Territory Maxim Reshetnikov, which took place as part of the MAKS-2019 aerospace show in Zhukovsky. According to the regional government, one of the main topics of the meeting was the development of the Novy Zvezdny technopolis in the Perm Territory and the related modernization of the Proton-PM enterprise (part of Roskosmos), where it is planned to launch mass production of RD-191 rocket engines. on environmentally friendly fuel components.

“I hope this will have a beneficial effect on the region. If there are any tests of production in the Perm Territory, then this is the RD-191 under the Angara. And this is an oxygen-jet engine, pure components. We love the Perm Territory, we love Kama, we I want to leave a bad mark in such a beautiful region," the press service of the Perm governor quoted Rogozin as saying.

According to the report, Rogozin clarified that the production of RD-191 engines for Angara launch vehicles will increase by a factor of 2023 with the start of mass production of rockets. In this regard, Rogozin drew attention to the development of the social infrastructure of the New Star cluster. "Here I am very grateful to the governor for all his efforts related to the development of infrastructure. Previously, they came to Perm - the working town was just developing. Now there will be new jobs, specialists, and it is necessary that they have not only a road, but also a good school ", Rogozin said.

Governor Reshetnikov, for his part, noted that Proton-PM PJSC has created a master plan, according to which infrastructure is being developed in the Novye Lyady microdistrict, a territory for the promising development of a technopolis.

According to the government of the Perm Territory, by 2025 it is planned to create a modern sports infrastructure in Novye Lyady and build a swimming pool. The buildings of the local polyclinic for 150 visits per day and technoschools will be renovated. V.P. Savinykh for 1,000 seats. In addition, it is planned to reconstruct the treatment facilities and the local filtering station.

"Angara" is a family of environmentally friendly launch vehicles of various classes. It includes light carriers "Angara-1.2", medium - "Angara-A3", heavy - "Angara-A5" and modernized "Angara-A5M", increased carrying capacity - "Angara-A5V". The RD-191 engine is used as part of the universal rocket module URM-1 of the Angara missiles. The light-class rocket Angara-1.2 uses one URM-1, the medium-sized Angara-A3 - three, and the heavy Angara-A5 - five.

MIA Rossiya Segodnya is the official media partner of the MAKS-2019 Aviation and Space Salon.