Bleed flow extraction system for a gas turbine engine

A gas turbine engine assembly for an aircraft, the gas turbine engine assembly comprising:

• a core engine comprising a compressor section, the compressor section defining a low pressure bleed port for extracting a first stream of bleed air and a high pressure bleed port for extracting a second stream of bleed air;
• and
• an air cycle machine configured for providing bleed air to an accessory system of the aircraft, the air cycle machine comprising:
• a turbine in fluid communication with the high pressure bleed port for receiving the second stream of bleed air, the turbine expanding the second stream of bleed air such that the second stream of bleed air rotates the turbine;
• a compressor in fluid communication with the low pressure bleed port for receiving the first stream of bleed air, the compressor being mechanically coupled to the turbine by a shaft such that the turbine drives the compressor and increases a pressure of the first stream of bleed air;
• and
• a junction in fluid communication with both the turbine and the compressor, the junction being configured to combine the first stream of bleed air from the compressor and the second stream of bleed air from the turbine into a combined bleed air stream to be supplied to the accessory system.