Results of experimental studies of parameters of low-thrust rocket engines operating on gaseous oxygen-hydrogen fuel

Abstract

The article presents the results of an experimental study of the thermal parameters and with-standing experimental samples rocket engines gaseous oxygen with thrust 25 N and 100 N. In experimental models rocket engines thrust 25 N and 100 N ignition fuel components arranged in the discharge chamber spark plugs. Scheme carburetion of engines characterized by the interaction of coaxial swirling flows of fuel and oxidant is carried out in two stages. At the same time, realized highly turbulent flow, contributing to the efficient mixing of fuel and oxidizer in a limited volume of the combustion chamber. Cooling chamber traction motor 25 H organized using a gas curtain of fuel from the mixing head and the application of high temperature structural material – boronsiliconized graphite, which is made of a combustion chamber and a nozzle. In rocket engines thrust 100N further organized subsonic curtain fuel located at the end of the cylindrical portion of the combustion chamber, but at the same time as the construction material used stainless steel type 12X18H10T. Experimental studies have been tested two structural variants of organization of the ignition process, evaluated the effectiveness of the scheme of mixing hydrogen and oxygen gases. In this case, the following values of specific impulse engines: for rocket engines thrust 25H with the geometric expansion ratio of the nozzle Fа = 45 - 3846 m/s; for rocket engines thrust 100N with Fа = 45 and Fа = 250 – respectively 3855 m/s and 4100 m/s. From the point of view of the thermal state in the study design rocket engines propellants, promising is the use of new construction materials, such as ceramics, graphite-based materials with the development of the technology for their production, as well as interfacing to the mixing head, made, usually made of stainless steel. In support of the above, in the use of the camera rocket engines thrust of 25N boron-siliconized graphite allowed during the test the engine for 100 seconds to get the maximum temperature of the outer surface of the chamber at ~ 1045 ° C.

About the authors

Y. I. Ageenko

Isaev Chemical Machinery Design Bureau – Branch of Federal State Unitary Enterprise «Khrunichev State Research and Production Space Center»

Author for correspondence.
Email: kbhimmash@korolev-net.ru

Candidate of Science (Engineering)

Chief designer

Russian Federation

E. A. Lapshin

Samara State Aerospace University

Email: ke_src@ssau.ru

Engineer Research Center of Space Energy

Russian Federation

I. I. Morozov

Samara State Aerospace University

Email: ke_src@ssau.ru

Researcher Research Centre for Space Energy

Russian Federation

I. V. Pegin

Isaev Chemical Machinery Design Bureau – Branch of Federal State Unitary Enterprise «Khrunichev State Research and Production Space Center»

Email: kbhimmash@korolev-net.ru

Deputy of the chief of the department liquid rocket engines of the small thrust

Russian Federation

V. V. Ryzhkov

Samara State Aerospace University

Email: ke_src@ssau.ru

Candidate of Science (Engineering)

Research supervisor of the Research center of space power

Russian Federation

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