Abstract: In a gas turbine engine internal cooling system a high pressure turbine disc is cooled by an air flow directed at the face of the disc by nozzles formed in the inner platforms of the downstream guide vane annulus. A proportion of the cooling air collected in a plenum adjacent the disc is inducted into a second part of the vane internal cooling system is via a multi-pass system of passages is used to cool the vane internally and is exhausted through trailing edge apertures into the gas stream. Because of the variation of temperature and static pressure across the radial distance of the annular gas stream the air flow is metered internally and distribute between exit apertures along the height of the vane trailing edge to provide required cooling and exit flow.