The rig stand for testing integrated rocket ramjet engine

• Autorzy: Rowiński A.

Identyfikatory

DOI

10.5604/12314005.1217261

• Języki publikacji: EN

Abstrakty

EN

Integrated rocket ramjet engine is adapted to aircraft propulsion and supersonic missiles moving at the speed of 8 Ma. The engine’s construction enables flexibly benefit from both types of drive depending on the conditions of the flight. The ejector mode of operation applicable to Mach numbers smaller than 2 cooperate with the rocket engine positioned in flow channel. Secondary air stream enters the engine through the convergent divergent nozzle and supplies the air to the ejector and booster. Rocket engine using the ejector effect would be used only in the phase of accelerating an object to the supersonic speed and then the drive would gradually shift to ramjet. The range of speed for the ramjet mode is 2-6 Mach. The prototype of the rocket ramjet engine of over 1300 N is equipped with annular combustion chamber in which phenomena of rotating detonation as well as the aero spike nozzle were used. Both the test stand as well as the engine is adapted to trials suitable to the conditions of a flight at the speed of 1.4 Ma. The test stand is powered by compress air coming from the instalment set up in the earth test bed and by oxygen and methane at a pressure of 10 bar. The rig is designed for functional tests of prototype, areas of the creation of mixture of firearms used to measuring and the range of stable functioning of the engine in the ramjet mode. Moreover, the measured parameters in gas supply installations as well as the temperature and pressure in the combustion chamber and thrust created by integrated rocket ramjet engine are measured.

Słowa kluczowe

EN

turbine engine; combustion chamber; detonation combustion; rotating detonation engine; RDE

• Wydawca: Łukasiewicz Research Network - Institute of Aviation ; European Science Society of Powertrain and Transport KONES Poland ; Wydawnictwo Instytutu Technicznego Wojsk Lotniczych
• Czasopismo: Journal of KONES
• Rocznik: 2016
• Tom: Vol. 23, No. 4
• Strony: 421--428
• Opis fizyczny: Bibliogr. 6 poz., rys.

Twórcy

autor

Rowiński A.

• Institute of Aviation Krakowska Avenue 110/114, 02-256 Warsaw, Poland tel.: [phone], fax: [phone]

Bibliografia

• [1] Bykovskij, F. A., Zhdan, S.A., Vedernikov, E., F., Continuous spin detonation in ducted annular combustors, Application of detonation to Propulsion, edited by Torus Press, 2004.
• [2] Kailasanath, K., Review of Propulsion Applications of Detonation Waves, AIAA Journal, Vol. 38, No. 9, 2000.
• [3] Kindracki, J., Badania eksperymentalne i symulacje numeryczne procesu inicjacji wirującej detonacji gazowej, PhD dissertation, MEiL, Politechnika Warszawska, Warszawa 2008.
• [4] Kindracki, J., Wolanski, P., Gut, Z., Experimental Research on the Rotating Detonation in Gaseous Fuels-Oxygen Mixtures, Shock Waves, Vol. 21, No. 2, pp. 75-84, 2011.
• [5] Łukasik, B., Czyż, S., Irzycki, A., Rowiński, R., The Study of the Continuously Rotating Detonation Combustion Chamber Supplied with Different Types of Fuel, Journal of KONES, Vol. 20, No. 3, Warsaw 2013.
• [6] Wolański, P., Koncepcja nowego silnika odrzutowego o spalaniu detonacyjnym, Progress in Astronautics, T. 29, Nr 1, 2006.

Uwagi

PL

Opracowanie ze środków MNiSW w ramach umowy 812/P-DUN/2016 na działalność upowszechniającą naukę.