Patents by Inventor Stephane M. M. BARALON
Stephane M. M. BARALON has filed for patents to protect the following inventions. This listing includes patent applications that are pending as well as patents that have already been granted by the United States Patent and Trademark Office (USPTO).
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Patent number: 11359493Abstract: A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (?3??1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.Type: GrantFiled: June 21, 2021Date of Patent: June 14, 2022Assignee: ROLLS-ROYCE PLCInventors: Benedict R. Phelps, Stephane M M Baralon, Mark J. Wilson
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Patent number: 11346229Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.Type: GrantFiled: February 22, 2021Date of Patent: May 31, 2022Assignee: Rolls-Royce plcInventors: Benedict Phelps, Stephane M M Baralon, Mark J. Wilson
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Patent number: 11339727Abstract: Gas turbine aircraft engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor; and a fan upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades. Each fan blade comprises a main blade and an auxiliary blade. Over substantially all of the auxiliary blade's radial span, the leading edge of the auxiliary blade is rearwards of the closest point on the trailing edge of the main fan blade, and on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate in tandem.Type: GrantFiled: November 20, 2020Date of Patent: May 24, 2022Assignee: ROLLS-ROYCE PLCInventors: Benedict R. Phelps, Stephane M M Baralon
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Patent number: 11268386Abstract: A gas turbine engine comprises carbon fibre fan blades. At cruise conditions, the fan tip air angle ? in the range: 64 degrees???67 degrees. Additionally or alternatively, the fan blade tip angle ? is in the range of from 62 to 69 degrees. Arrangements in accordance with the present disclosure provide advantages which may include improved bird-strike performance. This may allow advantages associated with carbon fibre fan blades to be better exploited.Type: GrantFiled: August 29, 2019Date of Patent: March 8, 2022Assignee: ROLLS-ROYCE plcInventors: Christopher Benson, Stephane M M Baralon
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Patent number: 11261734Abstract: The present disclosure relates to a fan blade for a gas turbine engine, the fan blade comprising an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span. A maximum thickness of the cross-sections through the aerofoil portion from a suction surface of the aerofoil portion to a pressure surface of the aerofoil portion perpendicular to the camber line decreases along the blade span from the root to the tip. There is a discontinuity in the rate of decrease of maximum thickness between a radius at 30% of the blade span from the aerofoil root and 70% of the blade span from the aerofoil root. The rate of decrease of the maximum thickness before the discontinuity is less than the rate of decrease of the maximum thickness after the discontinuity.Type: GrantFiled: August 20, 2019Date of Patent: March 1, 2022Assignee: ROLLS-ROYCE PLCInventors: Stephane M M Baralon, George Crammond, Stuart Andrews
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Publication number: 20210404342Abstract: A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (?3??1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.Type: ApplicationFiled: June 21, 2021Publication date: December 30, 2021Applicant: ROLLS-ROYCE plcInventors: Benedict R. PHELPS, Stephane M M BARALON, Mark J. WILSON
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Patent number: 11181042Abstract: A gas turbine engine has a cycle operability parameter ? in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined range of cycle operability parameter ? may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.Type: GrantFiled: May 14, 2019Date of Patent: November 23, 2021Assignee: ROLLS-ROYCE plcInventors: Michael O Hales, Craig W Bemment, Stephane M M Baralon, Benjamin J Sellers, Christopher Benson, Benedict R Phelps, Mark J Wilson
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Patent number: 11149690Abstract: A gas turbine engine 10 is provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P102) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P104), is no greater than a certain value. The gas turbine engine 10 may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.Type: GrantFiled: August 21, 2018Date of Patent: October 19, 2021Inventors: Benedict R. Phelps, Mark J. Wilson, Gabriel Gonzalez-Gutierrez, Nigel H S Smith, Marco Barale, Kashmir S. Johal, Stephane M M Baralon, Craig W. Bemment
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Patent number: 11085399Abstract: A gas turbine engine 10 is provided in which a fan having fan blades in which the camber distribution along the span allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.Type: GrantFiled: August 21, 2018Date of Patent: August 10, 2021Inventors: Benedict R. Phelps, Mark J. Wilson, Gabriel Gonzalez-Gutierrez, Nigel H S Smith, Marco Barale, Kashmir S. Johal, Stephane M M Baralon, Craig W. Bemment
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Publication number: 20210172323Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.Type: ApplicationFiled: February 22, 2021Publication date: June 10, 2021Inventors: Benedict Phelps, Stephane M. M. Baralon, Mark J. Wilson
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Publication number: 20210156317Abstract: Gas turbine aircraft engine comprising an engine core comprising a turbine, a compressor, a core shaft connecting the turbine to the compressor; and a fan upstream of the engine core and driven by the turbine, the fan comprising a circumferential row of tandem fan blades. Each fan blade comprises a main blade and an auxiliary blade. Over substantially all of the auxiliary blade's radial span, the leading edge of the auxiliary blade is rearwards of the closest point on the trailing edge of the main fan blade, and on a given aerofoil chordal section of the main fan blade, the leading edge position of an aerofoil chordal section of the auxiliary fan blade lies on a rearwards extension of the camber line of the aerofoil chordal section of the main fan blade, and the main fan blade and the auxiliary fan blade are arranged to rotate in tandem.Type: ApplicationFiled: November 20, 2020Publication date: May 27, 2021Inventors: Benedict R. Phelps, Stephane M M Baralon
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Patent number: 10954798Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.Type: GrantFiled: October 25, 2018Date of Patent: March 23, 2021Inventors: Benedict Phelps, Stephane M M Baralon, Mark J. Wilson
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Publication number: 20210071672Abstract: A gas turbine engine has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The bypass flow rate at the optimum bypass efficiency is appreciably lower than the maximum bypass flow rate at the given conditions. This results in increased design flexibility and improved overall engine performance.Type: ApplicationFiled: November 12, 2020Publication date: March 11, 2021Applicant: ROLLS-ROYCE PLCInventors: Stephane M M BARALON, Mark J WILSON, Benedict R PHELPS
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Patent number: 10876412Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.Type: GrantFiled: October 25, 2018Date of Patent: December 29, 2020Inventors: Mark J. Wilson, Stephane M M Baralon, Benedict Phelps
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Patent number: 10760427Abstract: A slot is provided in an endwall of a flow passage, for example between two stator vanes or rotor blades of a gas turbine engine. The length direction of the flow passage is aligned substantially with the main flow through the flow passage. The alignment of the slot means that the “over-turned” boundary layer flow can be extracted through the slot but with minimal impact on the mainstream flow.Type: GrantFiled: June 20, 2018Date of Patent: September 1, 2020Assignee: Rolls-Royce plcInventor: Stephane M M Baralon
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Publication number: 20200072058Abstract: A gas turbine engine comprises carbon fibre fan blades. At cruise conditions, the fan tip air angle ? in the range: 64 degrees???67 degrees. Additionally or alternatively, the fan blade tip angle ? is in the range of from 62 to 69 degrees. Arrangements in accordance with the present disclosure provide advantages which may include improved bird-strike performance. This may allow advantages associated with carbon fibre fan blades to be better exploited.Type: ApplicationFiled: August 29, 2019Publication date: March 5, 2020Inventors: Christopher BENSON, Stephane M M BARALON
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Publication number: 20200063569Abstract: The present disclosure relates to a fan blade for a gas turbine engine, the fan blade comprising an aerofoil portion having a leading edge extending from a root to a tip, the radial distance between the leading edge at the root and the leading edge at the tip defining a blade span. A maximum thickness of the cross-sections through the aerofoil portion from a suction surface of the aerofoil portion to a pressure surface of the aerofoil portion perpendicular to the camber line decreases along the blade span from the root to the tip. There is a discontinuity in the rate of decrease of maximum thickness between a radius at 30% of the blade span from the aerofoil root and 70% of the blade span from the aerofoil root. The rate of decrease of the maximum thickness before the discontinuity is less than the rate of decrease of the maximum thickness after the discontinuity.Type: ApplicationFiled: August 20, 2019Publication date: February 27, 2020Inventors: Stephane M. M. BARALON, George CRAMMOND, Stuart ANDREWS
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Publication number: 20200018169Abstract: A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.Type: ApplicationFiled: September 10, 2019Publication date: January 16, 2020Applicant: ROLLS-ROYCE PLCInventors: Stephane M M BARALON, Christopher BENSON, Benedict R. PHELPS
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Publication number: 20200011333Abstract: A gas turbine engine has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The bypass flow rate at the optimum bypass efficiency is appreciably lower than the maximum bypass flow rate at the given conditions. This results in increased design flexibility and improved overall engine performance.Type: ApplicationFiled: April 30, 2019Publication date: January 9, 2020Applicant: ROLLS-ROYCE plcInventors: Stephane M M BARALON, Mark J WILSON, Benedict R PHELPS
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Publication number: 20200011273Abstract: A gas turbine engine system has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The fan bypass inlet mass flow rate at the reference operating point is appreciably higher than the mass flow rate through the bypass duct at the peak bypass efficiency at a given fan reference rotational speed and cruise conditions. This results in increased design flexibility and improved overall engine performance.Type: ApplicationFiled: May 14, 2019Publication date: January 9, 2020Applicant: ROLLS-ROYCE plcInventors: Stephane M. M. BARALON, Mark J. WILSON, Benedict R. PHELPS