High temperature anti-ice systems for aircraft engines
Aircraft engines and high temperature anti-ice systems for aircraft engines are disclosed herein. An example aircraft engine includes: a fan including a plurality of fan blades; a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; a supply duct to accept bleed air from the compressor section; and a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft.
1 . An aircraft engine comprising:
a fan including a plurality of fan blades;
a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section;
a first supply duct to accept bleed air from the compressor section;
a second supply duct to accept the bleed air from the compressor section; and
a heat exchange system to capture waste heat from the turbine section and convey the waste heat to the bleed air, the bleed air with the waste heat to be conveyed to at least one of an environmental control system or a wing of an aircraft, the heat exchange system including a thermal transport bus that includes:
a closed flowpath;
a fluid in the closed flowpath; and
a pump to circulate the fluid in the closed flowpath, wherein the first supply duct is to provide the bleed air with waste heat to a first portion of the aircraft and the second supply duct is to provide the bleed air to a second portion of the aircraft.
2 . The aircraft engine of claim 1 , wherein the heat exchange system is to increase the bleed air to at least one of: (a) over 210 degrees Celsius or (b) over an autoignition temperature of jet fuel.
3 . The aircraft engine of claim 1 , wherein the heat exchange system is to increase the bleed air to over 232 degrees Celsius.
4 . The aircraft engine of claim 1 , wherein a portion of the closed flowpath is adjacent the turbine section.
5 . The aircraft engine of claim 1 , wherein a portion of the closed flowpath is in a post-combustion exhaust path in the turbine section.
6 . The aircraft engine of claim 1 , wherein the pump is to control a rate of flow of the fluid in the closed flowpath to control an amount of waste heat conveyed by the heat exchange system to the bleed air.
7 . The aircraft engine of claim 1 , wherein a portion of the heat exchange system is wrapped about at least a portion of the turbine section.
8 . The aircraft engine of claim 1 , wherein the bleed air with waste heat in the first supply duct has a greater temperature than the bleed air in the second supply duct.
9 . The aircraft engine of claim 1 , wherein the first portion of the aircraft is a first distance from a fuel tank and the second portion of the aircraft is a second distance from the fuel tank, the second distance less than the first distance.
10 . The aircraft engine of claim 9 , wherein the first portion includes at least one of a wing of the aircraft or a first portion of a truss near the wing and the second portion includes a second portion of the truss near fuselage.
11 . An aircraft engine comprising:
a fan including a plurality of fan blades;
a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section;
a heat exchange system including:
a closed flowpath;
a fluid in the closed flowpath; and
a pump to circulate the fluid in the closed flowpath, wherein a portion of the closed flowpath is at least one of (1) adjacent the turbine section or (2) in a post-combustion exhaust path in the turbine section, the heat exchange system to convey waste heat to the fluid; and
a return duct, the heat exchange system is to provide the fluid from the aircraft engine to an aircraft at a first temperature and to receive the fluid from the return duct at a second temperature, the first temperature greater than the second temperature.
12 . An aircraft engine comprising:
a fan including a plurality of fan blades;
a turbomachine operably coupled to the fan for driving the fan, the turbomachine including a compressor section, a combustion section, and a turbine section; and
a heat exchange system including:
a closed flowpath;
a fluid in the closed flowpath; and
a pump to circulate the fluid in the closed flowpath, wherein a portion of the closed flowpath is at least one of (1) adjacent the turbine section or (2) in a post-combustion exhaust path in the turbine section, the heat exchange system to convey waste heat to the fluid; and
a supply duct to accept bleed air from the compressor section, wherein the heat exchange system is to provide the fluid with waste heat to a first portion of an aircraft and the supply duct is to provide the bleed air to a second portion of the aircraft.
13 . The aircraft engine of claim 12 , wherein the fluid with waste heat has a greater temperature than the bleed air.
14 . The aircraft engine of claim 12 , wherein the first portion of the aircraft is a first distance from a fuel tank and the second portion of the aircraft is a second distance from the fuel tank, the second distance less than the first distance.
15 . The aircraft engine of claim 12 , wherein the first portion of the aircraft includes a wing and the second portion of the aircraft includes an interior cabin.