Abstract: The invention relates to a burner for combustors of gas turbines, the burner having an atomizer nozzle for atomisation of fuel in the combustion air and having a primary and secondary flow channel. In the burners, mostly used in aircraft engines, an extensively homogenous distribution of the air-fuel mixture is to be achieved in the combustor to reduce emissions. The flow channels opening into the combustor are separated from a first component arranged concentrically in relation to the burner axis and having a sleeve-shaped atomizer lip extending cylindrically or conically, and the external secondary flow channel is bounded externally and radially by a second annular component arranged concentrically and having an internal wall extending to converge and diverge. The second component forms an area with the most narrow flow cross-section, and the first component is arranged radially inwards and ends with the atomizer lip at the axial height of the cross-section or upstream thereof.

Abstract: A gas turbine combustor has at least one fuel nozzle located at a head end of the combustor. At least one swirl device is adjustable as a function of a load for supplying combustion air. The swirl device includes radial/tangential apertures formed between profiles of an annular body arranged coaxially with respect to the nozzle. The apertures are uniformly distributed around the circumference of the annular body and have aperture cross-sections that are constant along their entire length. The swirl device further includes a sleeve having fingers which extend inwardly to engage the apertures. The sleeve is adjustable with respect to the annular body. The fingers are axially displaceable within the apertures via the sleeve. The fingers are each arranged parallel to aperture walls spaced from the fingers to bound a maximal adjusting path of the fingers to form duct walls of the apertures.

Abstract: A fuel feed method and apparatus for a turboramjet engine having a combustion chamber operable as a turbine engine afterburner and as ramjet engine combustion chamber is disclosed which features the concurrent injection of kerosene and hydrogen. At low Mach numbers, kerosene is the predominant fuel injected, while at growing speeds, hydrogen is injected in increasing proportions. This provides for a short combustion chamber length while minimizing the on-board fuel load and benefits especially small-size hypersonic aircraft.

Abstract: A fuel injection device having three concentric streams of air for the atomization of fuel sprayed onto a sleeve by an injector facing into the sleeve. The inner and outer streams of air are imparted with swirl in opposite direction, while the central stream of air is free of swirl. The two inner streams of air atomize the fuel as a result of shear forces, while the outer stream of air forms a stable recirculation region in a combustion chamber into which the fuel injector device extends.

Abstract: A reverse flow combustion chamber includes an annular chamber enclosed between flame tube wall sections to which cooling air is so supplied from an outer annular channel acted upon with secondary air opposite the main flow direction in the flame tube in such a manner that the cooling air that it is blown out in the opposite flow direction film-like against an adjoining flame tube wall.