Johannes Lux

Effect of Recess in High-Pressure Liquid Oxygen/Methane Coaxial Injection and Combustion

The effect of a recessed liquid oxygen tube in shear coaxial injection has been investigated experimentally using an optically accessible subscale rocket combustor operated at pressures between 40 and 60 bar. Different single-shear coaxial injectors have been used to inject liquid oxygen and methane at relevant operating conditions covering sub-, near-, and supercritical pressures with respect to the critical point of oxygen. Liquid oxygen was injected at 120 K and the injection temperature of gaseous methane was about 275 K. Detection of spontaneous OH and CH chemiluminescene has been performed to characterize the flame-anchoring zone near the liquid oxygen post tip. In addition, the influence of the injector geometry on the combustion roughness and stability has been investigated during steady-state operating points. An increased flame expansion was observed with a recessed injector element. At low momentum flux ratio, the pressure drop accross the injector increases with a recessed liquid oxygen tube compared with a flush tube. Furthermore, a recessed liquid oxygen tube led to be a smoother combustion in general; however, this configuration also led to additional resonant frequencies in the chamber acoustics.

Flame Stabilization in High-Pressure Liquid Oxygen/Methane Rocket Engine Combustion

Flame stabilization in the near-injector region of a subscale liquid propellant rocket combustion chamber has been investigated using optical diagnostics. Several different single shear coaxial injectors have been used to inject liquid oxygen (LOX) and methane at relevant operating conditions. Three main operating points cover the range of sub-, near-, and superciritical conditions with respect to the thermodynamic cirtical point of oxygen. Whereas liquid temperatures of about 270 K. Similar to previous investigations performed with LOX/H2 combustion, spontaneous OH chemiluminescene has been detected to characterize the flame anchoring zone near the LOX post tip. Theoretical considerations have indicated that the binary mixture of oxygen and methane shows and entirely different behaviour compared with the oxygen/hydrogen system. This is believed to have an influence on the spray evolution and mixing characteristics at supercritical conditions. The experimental investigation includes both the ignition transient a well as the steady-state operating points. It has been detected that the LOX/CH4 flame shows very similar characteristics in comparison with LOX/H2 flame at similar operating conditions. Critical flame stabilization at the startup transient has been found during all hot runs. At steady-state conditions, the influence of the injection parameters on the flame shape is comparable to previous LOX/Hydrogen investigations.

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.

Porous injectors in cryogenic liquid propellant rocket engines at sub- and supercritical pressures

An alternative injection concept for the application in liquid propellant rocket engines has been successfully tested using two subscale rocket combustion chambers at pressures of up to 80 bar. With the new concept liquid oxygen (LOX) enters the chamber through numerous small diameter tubes in a classical parallel showerhead configuration. The fuel is completely fed through a porous faceplate made from sinter bronze. Two propellant combinations have been used, LOX/H2 and LOX/CH4. The near injector zone has been investigated using an optically accessible combustion chamber. In a second configuration an injector head covering the full faceplate has been used to acquire combustion efficiency as well as axial heat flux distribution data.