A gas turbine engine, comprising: a first compressor having a first overall pressure ratio; a second compressor having a second overall pressure ratio, wherein a ratio of said first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0; a first rotor disk supporting a first array of rotor blades; and a second rotor disk supporting a second array of rotor blades, wherein: a section of the gas turbine engine includes a thermally isolated area bounded at a fore location, an aft location, a radially inner location, and a radially outer location; the thermally isolated area is radially inward of a core airflow path of the gas turbine engine; the thermally isolated area is provided with a flow of cooling fluid from a single inlet; the single inlet is a tangential onboard injector (TOBI); the first compressor is a low pressure compressor; the second compressor is a high pressure compressor; and the section is a high pressure turbine section, characterised in that: the thermally isolated area is bounded at the fore location by at least one fore wall extending from the tangential onboard injector (TOBI) and a fore seal provided between the at least one fore wall and a fore-extending flange of the first rotor disk; the thermally isolated area is bounded at the radially outer location by a first outer seal extending from the first rotor disk, a second outer seal between the first and second rotor disk (70,72), and a third outer seal extending from the second rotor disk; the thermally isolated area is bounded at the aft location by an aft wall between a core airflow path boundary structure and an aft seal between the aft wall and an aft-extending flange of the second rotor disk; the first and second rotor disks (70, 72) are provided within the thermally isolated area; and the first and second arrays of rotor blades (64, 68) include internal passageways (132, 134, 160) configured to direct the flow of cooling fluid into the rotor blades (64, 68) and then into a core air flow path. a first compressor having a first overall pressure ratio; a second compressor having a second overall pressure ratio, wherein a ratio of said first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0; a first rotor disk supporting a first array of rotor blades; and a second rotor disk supporting a second array of rotor blades, wherein: a section of the gas turbine engine includes a thermally isolated area bounded at a fore location, an aft location, a radially inner location, and a radially outer location; the thermally isolated area is radially inward of a core airflow path of the gas turbine engine; the thermally isolated area is provided with a flow of cooling fluid from a single inlet; the single inlet is a tangential onboard injector (TOBI); the first compressor is a low pressure compressor; the second compressor is a high pressure compressor; and the section is a high pressure turbine section, characterised in that: the thermally isolated area is bounded at the fore location by at least one fore wall extending from the tangential onboard injector (TOBI) and a fore seal provided between the at least one fore wall and a fore-extending flange of the first rotor disk; the thermally isolated area is bounded at the radially outer location by a first outer seal extending from the first rotor disk, a second outer seal between the first and second rotor disk (70,72), and a third outer seal extending from the second rotor disk; the thermally isolated area is bounded at the aft location by an aft wall between a core airflow path boundary structure and an aft seal between the aft wall and an aft-extending flange of the second rotor disk; the first and second rotor disks (70, 72) are provided within the thermally isolated area; and the first and second arrays of rotor blades (64, 68) include internal passageways (132, 134, 160) configured to direct the flow of cooling fluid into the rotor blades (64, 68) and then into a core air flow path. a section of the gas turbine engine includes a thermally isolated area bounded at a fore location, an aft location, a radially inner location, and a radially outer location; the thermally isolated area is radially inward of a core airflow path of the gas turbine engine; the thermally isolated area is provided with a flow of cooling fluid from a single inlet; the single inlet is a tangential onboard injector (TOBI); the first compressor is a low pressure compressor; the second compressor is a high pressure compressor; and the section is a high pressure turbine section, the thermally isolated area is bounded at the fore location by at least one fore wall extending from the tangential onboard injector (TOBI) and a fore seal provided between the at least one fore wall and a fore-extending flange of the first rotor disk; the thermally isolated area is bounded at the radially outer location by a first outer seal extending from the first rotor disk, a second outer seal between the first and second rotor disk (70,72), and a third outer seal extending from the second rotor disk; the thermally isolated area is bounded at the aft location by an aft wall between a core airflow path boundary structure and an aft seal between the aft wall and an aft-extending flange of the second rotor disk; the first and second rotor disks (70, 72) are provided within the thermally isolated area; and the first and second arrays of rotor blades (64, 68) include internal passageways (132, 134, 160) configured to direct the flow of cooling fluid into the rotor blades (64, 68) and then into a core air flow path.