Aerospace engine with augmenting turbojet

A method for use with a gas turbine engine system, the method comprising:

• moving fluid with a fan section;
• compressing fluid from the fan section;
• passing at least a first portion of the compressed fluid through a first engine core to generate combustion products in the first engine core using at least some of the first portion of the compressed fluid;
• controlling a flowpath control mechanism to selectively direct a second portion of the compressed fluid to a second engine core or the first engine core;
• when the flowpath control mechanism allows the second portion of the compressed fluid to flow into the second engine core, generating combustion products in the second engine core using at least some of the second portion of the compressed fluid;
• and
• when the flowpath control mechanism blocks the second portion of the compressed fluid from flowing into the second engine core, generating combustion products in the first engine core using at least some of the second portion of the compressed fluid.

A gas turbine engine system comprising:

• a fan section;
• a bypass duct downstream of the fan section;
• a low pressure compressor section downstream of the fan section;
• a first engine core downstream from the low pressure compressor section, wherein the first engine core has a reverse flow configuration relative to fluid flow through the bypass duct, the first engine core comprising:
• a first engine core compressor section;
• a first engine core combustor downstream of the first engine core compressor section;
• and
• a first engine core turbine section downstream of the first engine core combustor;
• a low pressure turbine section located downstream of the first engine core, wherein only exhaust fluid from the first engine core passes through the low pressure turbine section;
• an exhaust pipe configured to direct exhaust fluid flow from the low pressure turbine section to the bypass duct;
• a second engine core downstream from the low pressure compressor section, the second engine core comprising:
• a second engine core compressor section;
• a second engine core combustor downstream of the second engine core compressor section;
• and
• a second engine core turbine section downstream of the second engine core combustor;
• and
• a flowpath control mechanism configured to selectively restrict fluid flow through the second engine core.