Abstract: A gas turbine engine annular combustor has a bulkhead at its upstream end which is protected by an annular array of heatshield segments. The segments are spaced apart from the bulkhead by means of a plurality of flanges. Cooling air apertures provided in the bulkhead direct cooling air into the space between the heatshield and the bulkhead. A portion of the cooling air is directed on to the downstream face of heatshield by means of a plurality of radially extending slots disposed around a fuel nozzle aperture in the centre of each heatshield segment. The slots provide for unrestrained thermal expansion of the heatshield region immediately surrounding the fuel nozzle aperture, thereby to prevent excessive thermal hoop stresses developing in the heatshield during engine operation.

Abstract: Combustion apparatus for gas turbine engines has a combustion chamber (4,5) provided with fuel/air mixture injection devices (11) in an upstream end wall (9,10) of the chamber. Each device has a passage (17) having inlets (12,15) for air and fuel which mix in the passage and discharge through an outlet (19) initially in a direction along the adjacent surface of said chamber end wall (9,10). The latter wall has a secondary inlet (20) admitting a stream of air transversely to the flow from the passage outlet (19) thereby to generate an oblique flow (19A) having a component movement (Y) away from the combustion chamber wall. This directs inflammable mixture away from the latter wall and simultaneously assists mixing of air and fuel. A cooling film (24,24A) passing along the chamber wall passes over the passage wall (16) and, in certain examples, between the oblique flow and the chamber wall.