Implementation of additive 3D printing technology in the development of an experimental oxygen-hydrogen low thrust rocket engine

Texto integral

Introduction

One of the consistent trends in developing domestic and foreign rocket and space technology is the widespread use of additive technologies (AF – Additive Fabrication, additivity – addition) [1–3], the main essence of which is the sequential layer-by-layer build-up of metallic or non-metallic material in accordance with the command-application program of 3D modeling. In the rocket engine manufacturing, the most promising method of additive technologies is considered to be SLM (Selective laser melting) technology of 3D printing, implemented in the form of laser melting of metal powder using CAD models for the specific engine parts [4–6].

The selective laser melting technology makes it possible to create complex monolithic mixing heads of rocket engines with nozzles providing minimal hydraulic and gas-dynamic losses, complex-profile parts of turbopump units, cooling ducts of liquid rocket engine (LRE) chambers with artificial roughness and other structural elements of engines.

Therefore, additive technologies open up opportunities to develop rocket engines with improved energy efficiency characteristics while simultaneously reducing the weight-dimension characteristics of products.

Simultaneously, the rocket and space industry should deal with the assignment of forming a technological system for post-printing processing parts with a focus on innovative electrochemical and electrophysical methods [7].

Analyzing the state of the problem

Analyzing the scope of application of additive technologies in the rocket engine manufacturing, it is advisable to consider the possibility of manufacturing low-thrust liquid rocket engines (LTRE) using the SLM printing method.

In terminology, liquid rocket engines are defined by GOST 22396-77 “Low-thrust liquid rocket engines” as actuators of the control system of spacecraft with a thrust from 0.01 to 1600 N. LTRE are divided into separate types according to various characteristics. In particular, according to their purpose, they are usually classified into the following types [8]:

• orientation LTRE;
• stabilization LTRE;
• correction LTRE;
• revers LTRE and others.

The specific operating conditions of low thrust liquid-propellant rocket engines in space, the features of their functional parameters and characteristics cause great difficulties in designing low-thrust rocket engines [9]. The assignment of designing and engineering-technological testing the low thrust liquid-propellant rocket engines is determined by the following factors:

• low propellant consumption;
• few mixing unit elements;
• complexity in creating regenerative cooling;
• difficulty in organizing effective mixture formation and ensuring high combustion efficiency.

The above mentioned factors, in their turn, create special problems in ensuring an acceptable thermal state of the engine chamber, which is determined by the temperature reserve of the wall in the critical section zone, preventing overheating of the mixing unit and boiling of propellant components in heat-stressed pulse modes.

Domestic liquid-propellant rocket engines use two-component and single-component propellant. High-boiling self-igniting components are used as two-component propellant: unsymmetrical dimethylhydrazine (UDMH – propellant) and nitrogen tetroxide (NT – oxidizer). The single-component propellant is usually hydrazine, which is gasified by a catalytic exothermic decomposition reaction. The scientific, technical and educational literature describes the physical, chemical and energy characteristics of these propellants quite fully, and their advantages and disadvantages are known. The previous research highlighted their reliability and trouble-free ignition, stability of the working process in various operating modes of the low thrust liquid-propellant rocket engine with these components.

Simultaneously, the rocket and space industry states new requirements for spacecraft (SC) and their propulsion systems (PS). The industry focuses on the tasks of optimizing the cost and time of design, development and production of engines, creation of propulsion and power plants for spacecraft with minimal mass and dimensional parameters with high power density. The environmental safety factors during ground testing, refueling, storage and operation of products are of primary importance.

Currently, research and design organizations, fields of the industry, with the involvement of scientists and specialists from the Russian higher education system, create the promising propulsion systems for new generation spacecraft, including small spacecraft for various purposes, as well as those using environmentally friendly propellant [12]. The scientific and technical groundwork in the direction of low thrust rocket engines using environmentally friendly propellant components is currently based on the results of research conducted at JSC NIIMash [13], JSC A. M. Isaev Design Bureau of Chemical Machine Building [14], PJSC RSC Energia [15], Bauman Moscow State Technical University [16], Moscow Aviation Institute [17], JSC State Research Center “Keldysh Center” [18], and others.

Concept of a promising low-thrust rocket engine

This article examines a fundamentally new model of a low-thrust rocket engine (LTRE) with the following design and technological features and characteristics:

• propellant component are environmentally friendly gaseous substances (hydrogen H₂ and oxygen O₂ or methane CH₄ and oxygen O₂);
• engine chamber is made of heat-resistant alloy using SLM printing technology in the form of a monolithic product;
• thermal protection of the chamber is provided by an internal cooling system by creating a wall layer together with regenerative cooling by a gaseous component;
• control system ensures multiple switching on and off to the engine in accordance with the operating cyclogram without changing the thrust value and the ratio of propellant components.

In order to study the possibility of practical implementation of these theoretical assumptions, JSC State Research Center “Keldysh Center” has developed an advanced model of a low thrust rocket engine operating on gaseous propellant components: oxygen and hydrogen. The engine chamber is manufactured using additive SLM printing technology made of Inconel 718 powder material [19; 20] as a single part without welded or soldered joints.

The camera was manufactured due to the Cooperation Agreement among JSC “State Scientific Center of the Russian Federation “M. V. Keldysh Research Center” and the Federal State Budgetary Educational Institution of Higher Education “Reshetnev Siberian State University of Science and Technology” with the participation of the university’s industrial partner OOO “Polychrome” (Krasnoyarsk). The entire printing process is implemented on the ASTRA 420 3D printer, developed and manufactured at this enterprise.

Fig. 1 presents the general view of the engine chamber. The combustion chamber (CC) has a spherical shape to ensures a uniform distribution of heat flows and temperature gradient throughout the entire volume of the sphere if the heat capacity of hydrogen is high. The internal diameter of the combustion chamber is about 33 mm. The critical section diameter is 12 mm. The supersonic part of the nozzle has the shape of a cone with a half-angle of 20°, the expansion ratio of the nozzle F = 11. The regenerative cooling path has 16 channels, the thickness of the fire wall is 0.6 mm, the height of the cooling channels is 1 mm, the thickness of the fins is 0.8 mm. The mixture formation scheme provides for the supply of the entire oxygen consumption through one nozzle with a diameter of 8 mm, located along the central axis of the chamber.

 

Рис. 1. Общий вид камеры РДМТ

Fig. 1. General view of the LTRE chamber

 

The design of the low thrust rocket engine is calculated for operation at a combustion chamber pressure p_cham from 7 to 10 kg/cm², a total propellant consumption ṁ_∑ from 30 to 40 g/s and a ratio of propellant components K_m from 3 to 4.5, the calculated thrust value is 150 N.

Methodology and 3D printing technology of the LTRE chamber

The technical capabilities of the ASTRA 420 3D printer make it possible to achieve relatively high speeds of volumetric construction of parts with the necessary adjustment of the laser power in the dynamic modulation mode of the laser spot [21–23].

Therefore, when developing the methodology and technology to manufacture an experimental model of the LTRE chamber, the main task was to maintain a balance between maintaining optimal energy-physical parameters that ensure the specified quality of the product and reducing the overall printing time according to the program of the scientific experiment.

Based on the results of testing the technology for printing products of similar sizes made of Inconel 718 metal powder, to manufacture the studied sample of the experimental chamber, the operating mode parameters are determined and presented in Table 1.

 

Table 1

Operating mode parameters

Laser power, W	Laser beam speed, mm/s	Laser beam diameter, µm	Shading step, µm	Idle speed, mm/s
300	600	190	150	600

 

To produce the parts with different spatial configurations, the general and individual approaches were analyzed. It was decided that for each printing case it was necessary to use individual patterns, which represented a separate path of the laser beam with the corresponding segmentation into layers. Areas with a thickness of up to 1.5 mm were scanned without changes, with continuous shading. The remaining area was cut off and scanned in a checkerboard pattern, in small sections without changing the overall printing speed (Fig. 2).

 

Рис. 2. Сегментация изделия и паттерны печати

Fig. 2. Product segmentation and printing patterns

 

When preparing the low thrust rocket engine project for printing, more than 40% of the entire time turned out to be spent idling the laser beam through the empty center of the combustion chamber. This happened because all the shading was considered as one element, which was filled track by track, without skipping empty areas without shading. The same also applies to other products that have a similar configuration – a cavity with a volume greater than the surrounding metal wall (Fig. 3).

 

Рис. 3. Штриховка цилиндрических изделий: 1 – сканирование изделия; 2 – холостой ход

Fig. 3. Shading cylindrical products: 1 – scanning of the product; 2 – idling

 

To optimize operation time, additional segmentation into equal parts around the central axis was used. Since all areas were considered independently of each other, they were scanned one by one without moving through the chamber cavity. Simultaneously, the required scanning angle was maintained over the entire area (Fig. 4).

 

Рис. 4. Сегментированное изделие с указанным ходом лазерного луча

Fig. 4. Segmented product with a specified laser beam path

 

Optimized print time was 46 hours, compared to 75 hours with conventional shading. Upon visual inspection of the product, no obvious tendency of the external relief to porosity was detected (Fig. 5).

 

Рис. 5. Камера РДМТ. Вид сбоку и сверху

Fig. 5. LTRE chamber. Side and top view

 

Preparation of the experimental LTRE to the bench test

The current state of the theory and practice of developing rocket engines does not allow to create new products that differ from previous analogues in many parameters and characteristics without conducting experimental research and bench tests close to the actual conditions of their operation. The main purpose of testing is to obtain reliable information about the condition of the tested product, which is subsequently used to solve scientific, design and technological problems. Therefore, testing is a vital part of programs and projects to develop highly efficient and reliable engines.

Experimental LTRE preparation to the bench test consists of installing the necessary fittings onto the engine to ensure propellant supply, ignition of components, and securing the product on the bench. Fig. 6 demonstrates the assembled LTRE design solution.

 

Рис. 6. Экспериментальный РДМТ:

1 – свеча зажигания; 2 – корпус; 3 – измерение давления в камере сгорания; 4 – камера сгорания с соплом; 5 – подача водорода; 6 – форсунка кислорода; 7 – подача кислорода

Fig. 6. Experimental LTRE:

1 – spark plug; 2 – housing; 3 – the combustion chamber pressure measurement; 4 – combustion chamber with nozzle; 5 – hydrogen inlet; 6 – oxygen injector; 7 – oxygen inlet

 

Engine chamber 4 has got an M14x1.5 connection for supplying hydrogen to the cooling tract. Hydrogen, passing through the cooling channels, moves along the walls of the combustion chamber to its initial section, leaves the cooling tract, mixing with oxygen.

Housing 2 has a M14x1.5 connection to join the oxygen supply solenoid valve and a threaded M18x1 socket for installing the CД-55AHM spark plug 1. The pressure in the combustion chamber is taken through the M12x1.25 fitting 3.

Test unit

Tests of the developed experimental unit of the oxygen-hydrogen low thrust rocket engine were conducted at bench 7 of JSC State Research Center “Keldysh Center”. Fig. 7 presents the pneumohydraulic scheme of the experimental unit.

 

Рис. 7. Пневмогидравлическая схема экспериментальной установки

Fig. 7. Pneumohydraulic diagram of the test unit

 

The unit ensures the supply of the required oxygen and hydrogen flow rates to the LTRE and the determination of their actual values by using flow-metering nozzles operating at a supersonic pressure drop.

To supply gaseous propellant components, the pipelines are used with a diameter of d_SI = 10 mm. The tests were carried out in continuous modes according to the cyclogram with simultaneous sending of commands to open the propellant solenoid valves and the spark plug. The spark plug switched off in 0.2 s after the start. The duration of the LTRE fire activations was 1–45 s.

Supplying gaseous propellant components to the engine inlet is carried out from cylinders with pressure reduction to the values provided for in the test program.

During the tests, the recorded data are the values of the pressures of the propellant components in front of the flow-metering nozzles Р_INP “О” and Р_INP “Prop”, after the nozzles (that is the pressure at the entrance to the LTRE) Р_ENG “О” and Р_ENG “Г” and in the combustion chamber p_C, as well as the temperatures of gaseous oxygen and hydrogen.

Using chromel-alumel thermocouples, temperatures were measured on the outer surface of the chamber T₁ – T₄. Fig. 8 demonstrates the LTRE equipped with thermocouples and installed on a test bench.

 

Рис. 8. Экспериментальный РДМТ, установленный на стенде

Fig. 8. Experimental LTRE mounted on a bench

 

The parameters obtained as a result of the LTRE tests are used to determine the consumption complex β using the relationship:

$\mathrm{\beta }=\frac{{Р}_{С}\cdot F_{с}}{{\dot{m}}_{\Sigma }},$

where F_c is the critical cross-sectional area of the engine nozzle; ṁ_∑ is the total mass flow rate of propellant, which consists of the oxidizer flow rate ṁ_о and the propellant flow rate ṁ_prop.

The coefficient of the consumption complex φ_β is defined as the ratio of the experimentally obtained value β to the theoretical β_Т, obtained as a result of the thermodynamic calculation for the values of K_m and р_C corresponding to the mode of the test performed:

${\mathrm{\phi }}_{\mathrm{\beta }}=\frac{\mathrm{\beta }}{{\mathrm{\beta }}_{Т}}.$

The coefficient φ_β is determined by the quality of mixing and combustion and characterizes the perfection of the operation process in the combustion chamber.

With the maximum values of measurement errors of pressure and temperature sensors, the total error in determining the mass flow rate of each of the propellant components does not exceed 0.7 %, the maximum error in determining the experimental value of the flow complex, taking into account the accuracy of determining the diameter and flow rate coefficient of the nozzles, is no more than 2.2 %.

Results of experimental research

Table 2 summarizes the data on all tests of the low thrust rocket engine. In total, 12 fire tests were conducted within the framework of the experimental program.

 

Table 2

Test results

Test №	Injector О2	t ign.	PINP “O”	РENG “О”	РINP “Prop”	РENG “Prop”	pC	ṁо	ṁprop	ṁ∑	Km	α	β	φβ
		s	105 Pa	105 Pa	105 Pa	105 Pa	105 Pa	g/s	g/s	g/s	–	–	m/s	–
1	№ 1	1	30.1	40.3	7.8	19.7	6.4	29.6	10.1	39.6	2.94	0.37	1806	0.71
2		5	30.6	40.3	7.8	19.6	6.5	30.0	10.1	40.1	2.98	0.38	1812	0.71
3		45	34.2	42.7	8.6	20.7	7.1	33.5	10.7	44.1	3.14	0.40	1800	0.71
4		20	31.2	40.4	8.4	19.4	6.4	30.8	10.1	40.9	3.04	0.38	1762	0.69
5		20	35.1	32.9	8.8	16.1	6.5	34.6	8.3	42.9	4.19	0.53	1714	0.70
6	№ 2	7	31.4	40.3	10.7	19.8	9.1	31.0	10.1	41.1	3.06	0.39	2500	0.98
7	№ 3	5	31.6	25.2	8.0	12.7	5.8	31.4	6.4	37.8	4.90	0.62	1729	0.72
8		11	30.7	40.2	8.8	19.4	6.9	30.5	10.2	40.7	3.00	0.38	1900	0.74
9		30	32.2	37.1	8.8	18.0	6.8	31.8	9.4	41.2	3.40	0.43	1859	0.74
10		30	34.9	34.0	9.1	16.7	6.9	34.6	8.6	43.2	4.03	0.51	1802	0.73
11		30	36.2	31.4	9.2	15.6	6.9	35.9	7.9	43.9	4.52	0.57	1768	0.73
12		27	37.9	30.1	9.2	15.1	7.0	37.6	7.6	45.2	4.95	0.62	1736	0.73

 

During the tests of the experimental LTRE, three variants of the oxygen injector were investigated (Fig. 9).

 

Рис. 9. Схемы вариантов форсунки кислорода

Fig. 9. Diagrams of oxygen injector options

 

Injector option № 1 is a single 8mm diameter hole located on the chamber axis. Option № 2 has 6 holes with a diameter of 2.6 mm, directed at an angle of 45° to the camera axis. Option № 3 is derived from option № 2 by adding a 6.5 mm diameter hole aligned with the camera axis.

Fig. 10 shows the dependence of the coefficient of the consumption complex on the ratio of propellant components, indicating a significant influence of the type of oxygen injector on the perfection of the operation process.

 

Рис. 10. Зависимость коэффициента расходного комплекса от соотношения компонентов топлива

Fig. 10. The Combustion efficiency dependence from the mixture ratio

 

The first 5 tests were carried out with injector option № 1.

Fig. 11 shows the change in the recorded parameters during test № 3 with a duration of 45 s. Temperature records Т₁ – Т₄ indicate that the combustion chamber reaches a steady-state thermal regime within ~15 s.

 

Рис. 11. Изменение параметров работы РДМТ в процессе испытания № 3 (форсунка № 1, ṁ_∑ = 44 г/с, K_m= 3,1)

Fig. 11. The operating parameters changes of LTRE during test № 3 (injector № 1, ṁ_∑ = 44 g/s, K_m = 3.1)

 

It is noted that when using the oxygen injector option № 1, low completeness of propellant combustion is achieved, characterized by values of the consumption complex coefficient of no more than φ_β = 0.71. The measured temperatures were also low and did not exceed 200 C. The characteristics obtained using injector № 1 are not acceptable for oxygen-hydrogen low thrust rocket engines. The low completeness of propellant combustion is probably explained by the stabilization of the flame at some distance from the outlet section of the oxygen injector, which, given the short length of the combustion chamber, significantly reduces the volume in which propellant combustion occurs.

In order to improve the characteristics of the low thrust rocket engine, oxygen injector option № 2 was installed in test № 6.

The new injector configuration has significantly changed the operation process in the LTRE chamber. The coefficient of the consumption complex increased to a value of φ_β = 0.98, which indicates a practically uniform mixing of the propellant components and their combustion throughout the entire volume of the combustion chamber. Oxygen directed through 6 holes actually destroys the hydrogen curtain, resulting in an intense increase in the measured temperatures on the outer surface of the combustion chamber during test № 6 (Fig. 12). Test № 6 was stopped at the 8th second, when the Т₄ temperature at the initial section of the chamber reached 1,100 C, which is close to the limit for the Inconel 718 material. Further tests with injector № 2 were not performed, since longer firing would inevitably result in burnout of the chamber.

 

Рис. 12. Изменение параметров работы РДМТ в процессе испытания № 6 (форсунка № 2, ṁ_∑ = 41 г/с, K_m = 3,1)

Fig. 12. The operating parameters changes of LTRE during test № 6 (injector № 2, ṁ_∑ = 41 g/s, K_m = 3.1)

 

For further testing, a central hole with a diameter of 6.5 mm was additionally made in injector № 2. The resulting injector № 3 was installed in the LTRE design during tests № 7–12.

The injector № 3 configuration made it possible to obtain a more acceptable combination of complete propellant combustion and thermal state of the structure. Tests of the LTRE with oxygen injector № 3 were carried out with a successive increase in the ratio of propellant components (Table 2). The duration of tests No. 9–12 was set equal to 30 s, which ensured that the measured temperatures of the combustion chamber reached steady-state values.

An increase in the ratio of the propellant components was accompanied by a corresponding increase in combustion chamber temperatures (Fig. 13).

 

Рис. 13. Зависимость максимальных температур на наружной поверхности камеры сгорания от соотношения компонентов топлива

Fig. 13. The maximum temperature dependence on the outer surface of the combustion chamber on the mixture ratio

 

Test No. 11 (Fig. 14, 15) demonstrated the design operability with a propellant component ratio of K_m = 4.5, which allows to achieve high energy-mass perfection of the PS). A further increase in the ratio of propellant components to K_m= 4.95 in test № 12 (Fig. 16) resulted in burnout of the combustion chamber wall at the 24th second. The burnout (Fig. 17) occurred in a section that had only boundary cooling with hydrogen.

 

Рис. 14. Изменение параметров работы РДМТ в процессе испытания № 11 (форсунка № 3, ṁ_∑ = 44 г/с, K_m = 4,5)

Fig. 14. The operating parameters changes of LTRE during test No. 11 (injector No. 3, ṁ_∑ = 44 g/s, K_m = 4.5)

 

Рис. 15. Огневая работа РДМТ на испытании № 11 (ṁ_∑ = 44 г/с, K_m = 4,5)

Fig. 15. Hot fire of LTRE on test № 11 (ṁ_∑ = 44 g/s, K_m = 4.5)

 

Рис. 16. Изменение параметров работы РДМТ в процессе испытания № 12 (форсунка № 3, ṁ_∑ = 45 г/с, K_m = 4,95)

Fig. 16. The operating parameters changes of LTRE during test No. 12 (injector No. 3, ṁ_∑ = 45 g/s, K_m = 4.95)

 

Рис. 17. Элементы РДМТ после испытаний

Fig. 17. The LTRE elements after testing

 

The pressure drop along the hydrogen line Р_inp “Prop” – p_C (Fig. 18) in the studied range of hydrogen consumption from ṁ_Prop = 6.4–10.7 g/s amounted to 6.9–13.7 × 10⁵ Pa. The main contribution to the hydraulic resistance of the propellant line is made by friction losses in the cooling channels. High hydraulic resistance of the cooling tract in oxygen-hydrogen low-thrust rocket engines is undesirable, since it may require an increase in the supply pressure of the propellant components to the engine. Further, in order to limit pressure losses along the fuel line, it is necessary to rationally profile the flow part of the cooling tract, as well as reduce the roughness of the walls of the cooling channels after manufacturing using the 3D printing method.

 

Рис. 18. Зависимость перепада давлений по линии водорода от расхода водорода

Fig. 18. The pressure drop dependence along the hydrogen line on hydrogen flow rate

 

Conclusion

As a result of the tests, the researched basic design of the oxygen-hydrogen low-thrust rocket engine, manufactured using the additive 3D printing technology, with features such as a combination of regenerative and boundary cooling, mixture formation with counter-current supply of oxygen and hydrogen and a spherical combustion chamber, demonstrated operability at a sufficiently high ratio of propellant components .

The main disadvantage of the studied scheme is the insufficient completeness of fuel combustion, characterized by the values of the coefficient of the consumption complex of no more than φ_β = 0.74, whereas to obtain a specific impulse of thrust J_SI = 3950 m/s, this indicator should be no less than φ_β = 0.88.

This fact, as well as the need to have reserves for the permissible ratio of propellant components, do not allow us to recommend the scheme without changes for further development.

Nevertheless, the conducted research made it possible to study the properties of the LTRE chamber manufactured using additive technology. The obtained experimental data will be used to refine previously developed calculation models of oxygen-hydrogen low-thrust rocket engines and in the development of an improved design of an oxygen-hydrogen low-thrust rocket engine.

It is noted that in the design of an oxygen-hydrogen low-thrust rocket engine with regenerative cooling, when the coolant – hydrogen – flows from the nozzle to the initial section, the maximum temperatures are realized not in the region of maximum heat flows in the critical section of the nozzle, but at the exit from the cooling tract, where the coolant temperature reaches its highest values.

Therefore, it could be concluded that the use of additive technologies opens up opportunities for the implementation of promising design and engineering solutions in the field of space propulsion engineering.

Sobre autores

Vladimir Koshlakov

JSC “Keldysh State Research Center”

Email: kerc@elnet.msk.ru

Dr. Sc., General Director

Rússia, 8, Onezhskaya St., Moscow, 125438

Sergey Mosolov

JSC “Keldysh State Research Center”

Email: mosolov@list.ru

Cand. Sc., Head of Liquid Rocket Engines Department

Rússia, 8, Onezhskaya St., Moscow, 125438

Alexander Klimenko

JSC “Keldysh State Research Center”

Email: klimenkokerc@mail.ru

Cand. Sc., Senior Researcher of Liquid Rocket Engines Department

Rússia, 8, Onezhskaya St., Moscow, 125438

Edkham Akbulatov

Reshetnev Siberian State University of Science and Technology

Email: rector@sibsau.ru

Cand. Sc., Associate professor, rector

Rússia, 31, Krasnoyarskii rabochii prospekt, Krasnoyarsk, 660037

Vladimir Nazarov

Reshetnev Siberian State University of Science and Technology

Autor responsável pela correspondência
Email: nazarov@sibsau.ru

Cand. Sc., Professor of the Department of Aircraft Engines

Rússia, 31, Krasnoyarskii rabochii prospekt, Krasnoyarsk, 660037

Evgeny Gerasimov

LLC “Polychrome”

Email: gerasimov24rus@mail.ru

Design engineer

Rússia, 58, Dubrovinskogo St., Krasnoyarsk, 660049

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