# US 2015/0176501 A1 · Bleed air systems for use with aircrafts and related methods > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 2015/0176501 A1 - Document ID: US20150176501A1 - Status: Published application - Kind: A1 - Application number: 14/632,322 - Filed: 2015-02-26 - Published: 2015-06-25 - Earliest claimed: 2012-01-24 - Assignees: The Boeing Company - Applicants: The Boeing Company - Inventors: Steve G. Mackin; George A. Ludlow - Web page: https://patentgrep.com/patent/US20150176501A1 ## CPC - F02C 9/18: by bleeding, bypassing or acting on variable working fluid interconnections between turbines or compressors or their stages {(F02C3/113 takes precedence)} - F02C 3/13: having variable working fluid interconnections between turbines or compressors or stages of different rotors {(controlling flow ratio between different flows of multi-flow jet-propulsion plant, e.g. ducted fan F02K3/075)} - F02C 6/04: Gas-turbine plants providing heated or pressurised working fluid for other apparatus, e.g. without mechanical power output (F02C6/18 takes precedence {; for a fluidised-bed combustor F02C3/205}) - F02C 6/08: the gas being bled from the gas-turbine compressor - F02C 7/32: Arrangement, mounting, or driving, of auxiliaries - B64D 2013/0618: {with arrangements for reducing or managing bleed air, using another air source, e.g. ram air} - Y02T 50/50: On board measures aiming to increase energy efficiency - Y02T 50/60: Efficient propulsion technologies, e.g. for aircraft - F02C: GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS (construction of turbines F01D; jet-propulsion plants F02K; construction of compressors or fans F04; generating combustion products of high pressure or high velocity F23R; using gas turbines in compression refrigeration plants F25B11/00) - B64D: EQUIPMENT FOR FITTING IN OR TO AIRCRAFT; FLIGHT SUITS; PARACHUTES; ARRANGEMENT OR MOUNTING OF POWER PLANTS OR PROPULSION TRANSMISSIONS IN AIRCRAFT - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## Abstract Bleed air systems for use with aircrafts and related methods are disclosed. An example apparatus includes a turbo-compressor including a compressor having a compressor inlet fluidly coupled to a low-pressure compressor of the aircraft engine and a compressor outlet fluidly coupled to a first system of an aircraft. The turbo-compressor also includes a turbine inlet fluidly coupled to a high-pressure compressor of the aircraft engine and a turbine outlet fluidly coupled to a second system of the aircraft. ## Claims ### Claim 1 An apparatus comprising: a turbo-compressor including: a compressor having a compressor inlet fluidly coupled to a low-pressure compressor of an aircraft engine and a compressor outlet fluidly coupled to a first system of an aircraft; and a turbine having a turbine inlet fluidly coupled to a high-pressure turbine of the aircraft engine and a turbine outlet fluidly coupled to a second system of the aircraft. ### Claim 2 (depends on 1) The apparatus of claim 1, wherein the second system is a low-pressure turbine of the aircraft engine. ### Claim 3 (depends on 2) The apparatus of claim 2, wherein the turbine outlet is fluidly coupled to a casing of the low-pressure turbine of the aircraft engine to provide cooled bleed air to the casing and blades within the casing. ### Claim 4 (depends on 1) The apparatus of claim 1, wherein the first system includes at least one of a thermal anti-icing system or an environmental control system. ### Claim 5 (depends on 1) The apparatus of claim 1, wherein the compressor of the turbo-compressor is to increase a pressure of bleed air received at the compressor inlet to a higher pressure at the compressor outlet. ### Claim 6 (depends on 1) The apparatus of claim 1, wherein the turbine inlet is fluidly coupled to a first stage of the high-pressure turbine of the aircraft engine and to a second stage of the high-pressure turbine of the aircraft engine, the first stage to provide higher pressure bleed air than the second stage. ### Claim 7 (depends on 1) The apparatus of claim 1, wherein the turbine inlet is fluidly coupled to a high-pressure compressor of the aircraft engine, the high-pressure turbine of the aircraft engine is to provide higher pressure bleed air than the high-pressure compressor of the aircraft engine. ### Claim 8 An aircraft comprising: a turbo-compressor comprising: a compressor and a turbine, the compressor having a compressor inlet and a compressor outlet, and the turbine having a turbine inlet and a turbine outlet; a first passageway to fluidly couple a first bleed air port from a compressor of an engine of the aircraft to the compressor inlet, the first passageway to provide bleed air at a first pressure from the compressor of the engine to the compressor inlet; a second passageway to fluidly couple a second bleed air port from a turbine of the engine to the turbine inlet, the second passageway to provide bleed air at a second pressure from the turbine of the engine to the turbine inlet, the second pressure higher than the first pressure; and a third passageway to fluidly couple the compressor outlet to a system of the aircraft that is to receive compressed air. ### Claim 9 (depends on 8) The aircraft of claim 8, wherein the turbine of the engine is a high-pressure turbine. ### Claim 10 (depends on 9) The aircraft of claim 9, wherein the system is a first system, the turbo-compressor further comprising a fourth passageway to fluidly couple the turbine outlet to a second system of the aircraft for cooling the second system. ### Claim 11 (depends on 10) The aircraft of claim 10, wherein the fourth passageway is to provide cooled bleed air to a casing of the low-pressure turbine of the engine. ### Claim 12 (depends on 10) The aircraft of claim 10 further comprising at least one of a thermal anti-icing system or an environmental control system, wherein the third passageway is to fluidly couple the compressor outlet to at least one of the thermal anti-icing system or the environmental control system. ### Claim 13 (depends on 8) The aircraft of claim 8, wherein the compressor is to increase a pressure of the bleed air received at the compressor inlet to a third pressure at the compressor outlet, the third pressure higher than the first pressure. ### Claim 14 A method comprising: receiving, at a compressor inlet of a turbo-compressor, a first bleed air from a compressor of an aircraft engine, the turbo-compressor comprising a turbine operatively coupled to a compressor; receiving, at a turbine inlet of the turbo-compressor, a second bleed air from a high-pres sure turbine of the aircraft engine to drive the turbine of the turbo-compressor; increasing a pressure of the first bleed air in the compressor of the turbo-compressor; and after increasing the pressure of the first bleed air, discharging the first bleed air through a compressor outlet of the turbo-compressor to a system of the aircraft. ### Claim 15 (depends on 14) The method of claim 14 further comprising discharging cooled bleed air through a turbine outlet of the turbo-compressor to a low-pressure turbine of the aircraft engine. ### Claim 16 (depends on 14) The method of claim 14 further comprising receiving, at the turbine inlet, a third bleed air having lower pressure than a pressure of the second bleed air. ### Claim 17 (depends on 14) The method of claim 14 further comprising receiving the second bleed air from the high-pressure turbine during a first operating state of the aircraft engine and receiving, at the turbine inlet, a third bleed air from a high-pressure compressor of the aircraft engine during a second operating state of the aircraft engine. ### Claim 18 (depends on 14) The method of claim 14 further comprising receiving the second bleed air from a first stage of the high-pressure turbine during a first operating state of the aircraft engine and receiving, at the turbine inlet, a third bleed air from a second stage of the high-pressure turbine during a second operating state of the aircraft engine, the second stage located downstream of the first stage. ### Claim 19 (depends on 14) The method of claim 14, wherein receiving the first bleed air comprises receiving bleed air from a low pressure compressor of the aircraft engine during a first period of operation, the method further comprising receiving, at the compressor inlet, bleed air from a high pressure compressor of the aircraft engine during a second period of operation.