Dmitry Suslov

Injector-Driven Combustion Instabilities in a Hydrogen/Oxygen Rocket Combustor

Self-excited combustion instabilities of the first tangential mode have been found in a research combustor operated with the cryogenic propellant combination of hydrogen/oxygen. In a series of consecutive test campaigns, the influence of operating conditions on these self-excited combustion instabilities was examined. This included a variation of the combustion chamber pressure, the mixture ratio, and the propellant temperatures. It has been shown how these operating parameters influence the resonance frequencies of the combustion chamber. The analysis of the influence of operating conditions on the oscillation amplitude of the first tangential mode indicated that the instability occurred when the frequency of the first tangential mode of the combustion chamber was shifted into the frequency of the second longitudinal mode of the liquid oxygen injector. With a variation of the injector length, and therefore its longitudinal resonance frequencies, this hypothesis has been tested. Based on the experimental ...

Film Cooling of Accelerated Flow in a Subscale Combustion Chamber

Experimental investigations have been carried out to examine film cooling effectiveness of an accelerated hot gas in a subscale rocket combustion chamber. In support of future first-stage high-performance rocket combustion chambers, a Vulcain2-like test case has been examined with combustion pressure levels up to 12 MPa. The effectiveness of an almost tagentially injected film of hydrogen with an initial temperature of approximately 280 K has been determined. Axial distributions of temperature were measured inside the copper liner as well as on the chamber surface in the convergent and divergent parts of the nozzle segment. An existing film cooling model has been modified for application in a combined convective and filmcooled combustion chamber with an accelerated hot gas. The new model predicts film cooling effectiveness at different combustion-chamber pressures and film blowing rates at sub-, trans-, and supersonic conditions.

Film Cooling in a High-Pressure Subscale Combustion Chamber

By the application of film cooling in addition to regenerative cooling, a considerable reduction in thermal and structural loads of rocket combustion-chamber walls can be reached. This paper discusses important influence parameters on film cooling in terms of efficiency of the injected film and wall temperature reduction. For the experimental investigations a high-pressure subscale combustion chamber operated with the cryogenic propellant combination LOX=GH2 was used. A gaseous film with ambient tempered hydrogen was injected in the axial direction at the face plate. Typical film-cooling parameters such as film blowing rate, velocity ratio between film injection velocity and hot-gas velocity, circumferential slot positioning, and film injection slot height were investigated systematically at the European Research and Technology Test Facility P8.

Investigation of Two Dimensional Thermal Loads in the Region Near the Injector Head of a High Pressure Subscale Combustion Chamber

Optimization of heat transfer management is a key issue in designing a rocket combustion chamber. This paper presents the new test specimen and the measurement technique that has been developed und successfully used by the Institute of Space Propulsion for high resolution investigations of the influence of injector-wall interactions on heat transfer in a subscale combustion chamber. The new measurement method allows obtaining two dimensional heat load distribution on the hot gas side at real rocket engine-like conditions at combustion chamber pressures up to 15 MPa. The presented investigations have been performed at the European Research and Technology Test Facility P8 for cryogenic subscale rocket engines.

Circumferential Film Cooling Effectiveness in a LOX/H2 Subscale Combustion Chamber

In order to study wall temperature distribution, film cooling effectiveness and wall heat flux reduction due to film cooling in the circumferential direction, experimental investigations have been carried out, using gaseous hydrogen (GH_2) as a film coolant with tangential slot injection in close proximity to the injector face plate. A Vulcain 2 like test case has been performed, using a combustion chamber pressure up to 11.5 MPa at a high mixture ratio ROF = 6 with ambient temperatures of hydrogen. The experimental investigations have shown significant variations of wall temperatures due to injector design as well as a distribution of film cooling effectiveness which persists far downstream from the film coolant injection slots. Circumferential variations of wall temperature and film cooling effectiveness are much more pronounced at higher combustion chamber pressures.

Investigation Of Sub- And Supercritical LOX/Methane Injection Using Optical Diagnostics

The DLR Institute of Space Propulsion at Lampoldshausen has been working on dierent aspects of oxygen/methane combustion for a couple of years. Within this framework, the European High Pressure Research and Technology Test Facility P8 was equipped with a methane fluid system as a first step to enable single injector investigations of high pressure methane injection and combustion processes. The fluid system provides ambient temperature methane at flow rates up to 1 kg/s and at combustion chamber pressures up to 10 MPa. The measurement system has been adapted to the new fluid system to provide a precise determination of injection conditions. Hence, the P8 now allows investigations of LOX/CH4 combustion chamber processes under realistic rocket engine conditions. For the first operational tests, a single injector windowed rocket combustion chamber was operated at various steady-state conditions in the sub-, near- and supercritical regime with respect to the critical pressure of oxygen. The propellants were injected through a single shear coaxial injector element at temperatures of about 120 K and 275 K, respectively. High speed optical diagnostic techniques such as flame emission spectroscopy and shadowgraphy have been applied in the near injector region to assess injection and atomization behavior as well as flame anchoring and stabilization. The paper describes the modifications of fluid and MCC systems at P8 and presents preliminary results of the first test campaign.

Experimental Investigation of Porous Injectors for Liquid Propellant Rocket Engines

Three different injector heads have been tested using a LOX/H 2 subscale combustion chamber at the European High Pressure Research and Technology Test Facility P8. For reference purposes, a first configuration based on a classical shear coaxial injector head with 13 elements has been investigated. The second configuration consists of a planar porous faceplate made from sinter bronze and 68 implemented single LOX tubes. Hydrogen was injected through the porous faceplate whereas liquid oxygen enters the combustion chamber in a classical parallel showerhead configuration at moderate injection velocities. With the third configuration, a hemispherical shaped faceplate has been used in order to increase both the mechanical strength and the hydrogen injection area. In addition, the LOX tubes have been modified from a parallel to a far downstream impinging arrangement in order to enhance the combustor wall compatibility. Liquid oxygen was injected at 120 K whereas the hydrogen injection temperature was about 50 K with the porous configurations and about 90 K with the coaxial injector head. c*-efficiency data have been gathered as well as some basic information on the axial heat flux distribution during steady state operation. All three configurations have been operated at combustion chamber pressures up to 80 bar and propellant mixture ratios between 1.5 and 6. In addition to steady states, ramp-tests have been performed in order to investigate the throttling capabilities of the porous injectors.

Influence Parameters on Film Cooling Effectiveness in a High Pressure Subscale Combustion Chamber

In typical high performance rocket engines like the European Vulcain 2 or the Space Shuttle Main Engine (SSME) a significant reduction of thermal and structural loads is achieved by the application of film cooling with an acceptable loss in overall engine performance. Within the presented paper, the effectiveness of film cooling has been studied in a LOX/GH2 subscale rocket combustion chamber. Aiming at high performance first stage applications, combustion chamber pressures have been varied up to 12 MPa. Using GH2 with ambient temperatures for film cooling, typical influence parameters on film cooling effectiveness like film blowing rate, velocity ratio between film injection velocity and hot gas velocity, and film injection slot height have been investigated systematically at the European Research and Technology Test Facility P8.

Circumferential Behavior of Tangential Film Cooling and Injector Wall Compatibility in a High Pressure LOX/GH2 Subscale Combustion Chamber

Within the present investigations of high pressure combustion at DLR Lampoldshausen, circumferential film cooling effectiveness in combination with injector wall compatibility have been studied in a LOX/GH2 fired subscale combustion chamber using gaseous hydrogen as a film coolant, injected in tangential direction in close proximity of the injector head. With the aim of future first stage high performance rocket combustion chambers, the combustion pressure levels have been varied up to 11.5 MPa with real engine like conditions. Experimental investigations have shown remarkable circumferential variations of thermal head load due to injector design and film effectiveness which remain existent far downstream from the film injection point and injector face plate, and which are more pronounced at higher chamber pressure levels.

Combustion efficiency sensitivity studies of the API injector concept

Since a few years DLR Lampoldshausen has been engaged in the development of a new injection concept for cryogenic liquid propellant rocket engines. This concept bases on the injection of the liquid oxygen through a large number of simple tubes while the entire hydrogen is injected through the porous face plate.3, 4 The key feature of the API concept is the independency of the atomization, vaporization, mixing and combustion performance of initial injection conditions such as velocity ratio, momentum fluxratio, Weber number or other non-dimensional parameters which are generally deemed important. This statement will be supported by data optained with a 50mm subscale combustion chamber. The influence of the velocity ratio at injection will be shown through a variation of the propellant mixture ratio and/or the chamber pressure.