# US 7,096,674 B2 · High thrust gas turbine engine with improved core system > Agents: this data is available as JSON from the patentgrep API and MCP server — see https://patentgrep.com/llms.txt - Number: US 7,096,674 B2 - Document ID: US7096674B2 - Status: Granted - Kind: B2 - Application number: 10/941,546 - Filed: 2004-09-15 - Published: 2006-03-16 - Granted: 2006-08-29 - Earliest claimed: 2004-09-15 - Assignees: General Electric Company - Inventors: Robert Joseph Orlando; Kattalaicheri Srinivasan Venkataramani; Ching-Pang Lee; Thomas Ory Moniz; Kurt David Murrow - Web page: https://patentgrep.com/patent/US7096674B2 ## CPC - F02C 7/18: the medium being gaseous, e.g. air {(F02C7/125 takes precedence)} - F02C 5/00: Gas-turbine plants characterised by the working fluid being generated by intermittent combustion - F23R 3/005: {Combined with pressure or heat exchangers} - F23R 3/04: Air inlet arrangements - F23R 7/00: Intermittent or explosive combustion chambers - F05D 2260/20: Heat transfer, e.g. cooling - F05D 2260/96: Preventing, counteracting or reducing vibration or noise - F23R 2900/03044: Impingement cooled combustion chamber walls or subassemblies - F23R 2900/03281: Intermittent fuel injection or supply with plunger pump or other means therefor - 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) - F23R: GENERATING COMBUSTION PRODUCTS OF HIGH PRESSURE OR HIGH VELOCITY, e.g. GAS-TURBINE COMBUSTION CHAMBERS (fluidised bed combustion apparatus specially adapted for operation at superatmospheric pressures F23C10/16) - F05D: INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS - Y02T: CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION ## Abstract A gas turbine engine having a longitudinal centerline axis therethrough, including: a fan section at a forward end of the gas turbine engine including at least a first fan blade row connected to a first drive shaft; a booster compressor positioned downstream of and in at least partial flow communication with the fan section including a plurality of stages, each stage including a stationary compressor blade row and a rotating compressor blade row connected to a drive shaft and interdigitated with the stationary compressor blade row; a core system positioned downstream of the compressor, where the core system further includes an intermediate compressor positioned downstream of and in flow communication with the booster compressor, the intermediate compressor being connected to a second drive shaft, and a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet; and, a low pressure turbine positioned downstream of and in flow communication with the core system, the low pressure turbine being utilized to power the first drive shaft. The core system may also include an intermediate turbine positioned downstream of the combustion system in flow communication with the working fluid, where the intermediate turbine is utilized to power the second drive shaft. A first source of compressed air having a predetermined pressure is provided to the combustion system inlet and a second source of compressed air having a pressure greater than the first source of compressed air is provided to cool the combustion system. ## Claims ### Claim 1 A gas turbine engine having a longitudinal centerline axis therethrough, comprising: (a) a fan section at a forward end of said gas turbine engine including at least a first fan blade row connected to a first drive shaft; (b) a booster compressor positioned downstream of and in at least partial flow communication with said fan section including a plurality of stages, each said stage including a stationary compressor blade row and a rotating compressor blade row connected to said first drive shaft and interdigitated with said stationary compressor blade row; (c) a core system positioned downstream of said booster compressor, said core system further comprising: (1) an intermediate compressor positioned downstream of and in flow communication with said booster compressor, said intermediate compressor being connected to a second drive shaft; and, (2) a combustion system for producing pulses of gas having increased pressure and temperature from a fluid flow provided to an inlet thereof so as to produce a working fluid at an outlet; and, (d) a low pressure turbine positioned downstream of and in flow communication with said core system, said low pressure turbine being utilized to power said first drive shaft; wherein a first source of compressed air having a predetermined pressure is provided to said combustion system inlet and a second source of compressed air having a pressure greater than said first source of compressed air is provided to cool said combustion system. ### Claim 2 (depends on 1) The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by at least approximately 20%. ### Claim 3 (depends on 1) The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by at least approximately 50%. ### Claim 4 (depends on 1) The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of compressed air from said first source by at least approximately 100%. ### Claim 5 (depends on 1) The gas turbine engine of claim 1, wherein said first source of compressed air originates in said booster compressor. ### Claim 6 (depends on 1) The gas turbine engine of claim 1, wherein said first source of compressed air originates between adjacent stages of said intermediate compressor. ### Claim 7 (depends on 1) The gas turbine engine of claim 1, wherein said second source of compressed air originates at an aft end of said intermediate compressor. ### Claim 8 (depends on 1) The gas turbine engine of claim 1, wherein said second source of compressed air originates from said intermediate compressor. ### Claim 9 (depends on 1) The gas turbine engine of claim 1, wherein pressure of compressed air from said second source is greater than pressure of said working fluid at said combustion system outlet. ### Claim 10 (depends on 1) The gas turbine engine of claim 1, wherein said combustion system is a constant volume combustor. ### Claim 11 (depends on 1) The gas turbine engine of claim 1, wherein said combustion system is a pulse detonation device. ### Claim 12 (depends on 1) The gas turbine engine of claim 1, wherein said combustion system includes at least one rotating member for powering said second drive shaft. ### Claim 13 (depends on 12) The gas turbine engine of claim 12, wherein said rotating member of said combustion system powers said second drive shaft. ### Claim 14 (depends on 1) The gas turbine engine of claim 1, wherein said combustion system includes no rotating members. ### Claim 15 (depends on 14) The gas turbine engine of claim 14, further comprising an intermediate turbine positioned downstream of said combustion system in flow communication with said working fluid, said intermediate turbine being utilized to power said second drive shaft. ### Claim 16 (depends on 1) The gas turbine engine of claim 1, wherein said booster compressor is driven by said first drive shaft. ### Claim 17 (depends on 1) The gas turbine engine of claim 1, wherein said booster compressor is driven by said second drive shaft. ### Claim 18 (depends on 1) The gas turbine engine of claim 1, wherein compressed air from said second source is provided to cool fluid entering said intermediate turbine. ### Claim 19 (depends on 1) The gas turbine engine of claim 1, wherein a portion of compressed air from said second source is supplied to said combustion system to assist atomization of fuel therein. ### Claim 20 (depends on 1) The gas turbine engine of claim 1, further comprising a heat exchanger in flow communication with said compressed air of said second source. ### Claim 21 (depends on 1) The gas turbine engine of claim 1, wherein said gas turbine engine is able to generate a maximum of approximately 60,000 pounds of thrust. ### Claim 22 A method of cooling a combustion system of a gas turbine engine including a booster compressor and an intermediate compressor, wherein said combustion system produces pulses of gas having increased pressure and temperature from a fluid flow provided thereto, comprising the following steps: (a) providing a first source of compressed air having a predetermined pressure to said combustion system; and (b) providing a second source of compressed air having a pressure greater than said first source of compressed air to cool said combustion system. ### Claim 23 (depends on 22) The method of claim 22, further comprising the step of originating said first source from said booster compressor. ### Claim 24 (depends on 22) The method of claim 22, further comprising the step of originating said second source from said intermediate compressor. ### Claim 25 (depends on 22) The method of claim 22, further comprising the step of originating said first source between adjacent stages of said intermediate compressor. ### Claim 26 (depends on 25) The method of claim 25, further comprising the step of originating said second source at an aft end of said intermediate compressor. ### Claim 27 (depends on 22) The method of claim 22, further comprising the step of cooling compressed air from said second source. ### Claim 28 A gas turbine engine, comprising: (a) a compressor positioned at a forward end of said gas turbine engine having a plurality of stages, each said stage including a stationary compressor blade row and a rotatable blade row connected to a first drive shaft and interdigitated with said first compressor blade row; (b) a core system positioned downstream of said compressor, said core system further comprising: (1) an intermediate compressor positioned downstream of and in flow communication with said compressor connected to a second drive shaft; (2) a combustion system for producing pulses of gas having increased pressure and temperature from a fluid supplied to an inlet thereof so as to produce a working flow at an outlet thereof; and, (3) an intermediate turbine positioned downstream of said combustion system in flow communication with said working fluid, said intermediate turbine being utilized to power said second drive shaft; (c) a low pressure turbine downstream of and in flow communication with said intermediate turbine for powering said first drive shaft; and, (d) a load connected to said first drive shaft; wherein a first source of compressed air having a predetermined pressure is provided to said combustion system and a second source of compressed air having a pressure greater than compressed air from said first source is provided to cool said combustion system.