Patents by Inventor Craig W. BEMMENT

Craig W. BEMMENT 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).

  • Publication number: 20220205386
    Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.
    Type: Application
    Filed: March 17, 2022
    Publication date: June 30, 2022
    Applicant: ROLLS-ROYCE PLC
    Inventors: Craig W BEMMENT, Pascal DUNNING
  • Patent number: 11326512
    Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.
    Type: Grant
    Filed: June 11, 2021
    Date of Patent: May 10, 2022
    Assignee: ROLLS-ROYCE plc
    Inventors: Craig W Bemment, Pascal Dunning
  • Publication number: 20220099035
    Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A turbine to fan tip temperature change ratio of a low pressure turbine temperature change to a fan tip temperature rise is in the range from 1.46 to 2.0.
    Type: Application
    Filed: September 3, 2021
    Publication date: March 31, 2022
    Applicant: ROLLS-ROYCE PLC
    Inventors: Craig W BEMMENT, Pascal DUNNING
  • Patent number: 11280274
    Abstract: There is provided a gas turbine engine comprising a low pressure shaft and a high pressure shaft; wherein the low pressure shaft connects a fan to a fan drive turbine, and the high pressure shaft connects a high pressure turbine to a compressor section. The low pressure shaft and the high pressure shaft are arranged such that when operating at idle the idle shaft speed ratio is greater than 6.05. The idle shaft speed ratio is the ratio of the speed of the high pressure shaft to the speed of the low pressure shaft at idle conditions.
    Type: Grant
    Filed: August 13, 2020
    Date of Patent: March 22, 2022
    Assignee: ROLLS-ROYCE plc
    Inventor: Craig W Bemment
  • Publication number: 20210388772
    Abstract: A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.
    Type: Application
    Filed: August 25, 2021
    Publication date: December 16, 2021
    Applicant: ROLLS-ROYCE plc
    Inventor: Craig W. BEMMENT
  • Patent number: 11181042
    Abstract: 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: Grant
    Filed: May 14, 2019
    Date of Patent: November 23, 2021
    Assignee: ROLLS-ROYCE plc
    Inventors: Michael O Hales, Craig W Bemment, Stephane M M Baralon, Benjamin J Sellers, Christopher Benson, Benedict R Phelps, Mark J Wilson
  • Patent number: 11149690
    Abstract: 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: Grant
    Filed: August 21, 2018
    Date of Patent: October 19, 2021
    Inventors: 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
  • Publication number: 20210310407
    Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.
    Type: Application
    Filed: June 11, 2021
    Publication date: October 7, 2021
    Applicant: ROLLS-ROYCE PLC
    Inventors: Craig W BEMMENT, Pascal DUNNING
  • Patent number: 11136922
    Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A low pressure turbine temperature change is defined as: the ? ? first ? ? turbine ? ? entrance ? ? temperature the ? ? first ? ? turbine ? ? exit ? ? temperature . A fan tip temperature rise is defined as: the ? ? fan ? ? tip ? ? rotor ? ? exit ? ? temperature the ? ? fan ? ? rotor ? ? entry ? ? temperature .
    Type: Grant
    Filed: August 20, 2019
    Date of Patent: October 5, 2021
    Assignee: ROLLS-ROYCE plc
    Inventors: Craig W Bemment, Pascal Dunning
  • Publication number: 20210301718
    Abstract: A gas turbine engine (10) comprising: a high pressure turbine (17); a low pressure turbine (19); a high pressure compressor (15) coupled to the high pressure turbine (17) by a high pressure shaft (27); a propulsor (23) and a low pressure compressor (14) coupled to the low pressure turbine (19) via a low pressure shaft (26) and a reduction gearbox (30); wherein the low pressure compressor (14) consists of four compressor stages (14) and defines a cruise pressure ratio of between 2.4:1 and 3.3:1; the high pressure compressor (15) defines a cruise pressure ratio of less than 17:1; and the high pressure compressor (15) and low pressure compressor (14) together define a cruise core overall pressure ratio of greater than 36:1.
    Type: Application
    Filed: March 9, 2021
    Publication date: September 30, 2021
    Applicant: ROLLS-ROYCE plc
    Inventors: Michael O. HALES, Craig W. BEMMENT, Benjamin J. SELLERS, Ian J. BOUSFIELD, Amarveer S. MANN
  • Patent number: 11131250
    Abstract: A gas turbine engine for an aircraft and a method of operating a gas turbine engine on an aircraft. Embodiments disclosed include a gas turbine engine for an aircraft including: an engine core has a turbine, a compressor, and a core shaft; a fan located upstream of the engine core, the fan has a plurality of fan blades; a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle; and a gearbox that receives an input from the core shaft and outputs drive to the fan wherein the gas turbine engine is configured such that a jet velocity ratio of a first jet velocity exiting from the bypass exhaust nozzle to a second jet velocity exiting from an exhaust nozzle of the engine core at idle conditions is greater by a factor of 2 or more than the jet velocity ratio at maximum take-off conditions.
    Type: Grant
    Filed: July 30, 2019
    Date of Patent: September 28, 2021
    Assignee: ROLLS-ROYCE plc
    Inventor: Craig W Bemment
  • Patent number: 11085399
    Abstract: 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: Grant
    Filed: August 21, 2018
    Date of Patent: August 10, 2021
    Inventors: 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
  • Patent number: 11078838
    Abstract: A method of operating a gas turbine engine compressor. The engine comprises a compressor having an environmental control system bleed port having an outlet in fluid communication with an aircraft environmental control system air duct, and an air turbine starter configured to rotate a compressor shaft of the gas turbine engine. The air turbine starter has an inlet in fluid communication with the environmental control system air duct via an air turbine valve. The method comprises determining a surge margin of the compressor, and where the surge margin of the compressor is determined to be below a predetermined minimum surge margin, opening the air turbine valve to supply air to the air turbine.
    Type: Grant
    Filed: April 29, 2019
    Date of Patent: August 3, 2021
    Inventors: Craig W. Bemment, David P. Scothern
  • Publication number: 20210231060
    Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.
    Type: Application
    Filed: April 15, 2021
    Publication date: July 29, 2021
    Applicant: ROLLS-ROYCE plc
    Inventor: Craig W BEMMENT
  • Publication number: 20210222632
    Abstract: A gas turbine engine comprises a fan, a compressor, a low pressure turbine and a high pressure turbine. The fan diameter is greater than 250 cm and less than 381 cm; and the gas turbine engine has a first thrust at sea level static conditions and a second thrust at end of runway conditions; and a thrust take-off ratio greater than 1.32; wherein the thrust take-off ratio is the ratio of the first thrust to the second thrust.
    Type: Application
    Filed: January 7, 2021
    Publication date: July 22, 2021
    Applicant: ROLLS-ROYCE PLC
    Inventors: Craig W. BEMMENT, Pascal DUNNING
  • Patent number: 11053842
    Abstract: An engine core including turbine, compressor, and core shaft connecting the turbine to the compressor, wherein a compressor exit temperature has an average airflow; and a fan upstream including a plurality of fan blades extending from a hub, each fan blade having a leading and trailing edge, wherein a fan rotor entry temperature has an average airflow across the leading edge of each blade at cruise conditions and fan tip rotor exit temperature has an average temperature of airflow across a radially outer portion of each blade at the trailing edge cruise conditions. A fan tip temperature rise as: the ? ? fan ? ? tip ? ? rotor ? ? exit ? ? temperature the ? ? fan ? ? rotor ? ? entry ? ? temperature .
    Type: Grant
    Filed: September 3, 2019
    Date of Patent: July 6, 2021
    Assignee: ROLLS-ROYCE plc
    Inventors: Craig W Bemment, Pascal Dunning
  • Patent number: 11002194
    Abstract: The present disclosure relates to a geared gas turbine engine for an aircraft. Example embodiments include a gas turbine engine for an aircraft including: an engine core having a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan having a plurality of fan blades; and a gearbox that receives an input from the core shaft to drive the fan at a lower rotational speed than the core shaft, the gearbox having a gear ratio of around 3.4 or higher, wherein the gas turbine engine is configured such that a jet velocity ratio between a first jet velocity exiting from a bypass duct of the engine and a second jet velocity exiting from an exhaust nozzle of the engine core is within a range from around 0.75 to around 0.82.
    Type: Grant
    Filed: June 18, 2019
    Date of Patent: May 11, 2021
    Assignee: ROLLS-ROYCE plc
    Inventor: Craig W Bemment
  • Publication number: 20210108570
    Abstract: There is provided a gas turbine engine comprising a low pressure shaft and a high pressure shaft; wherein the low pressure shaft connects a fan to a fan drive turbine, and the high pressure shaft connects a high pressure turbine to a compressor section. The low pressure shaft and the high pressure shaft are arranged such that when operating at idle the idle shaft speed ratio is greater than 6.05. The idle shaft speed ratio is the ratio of the speed of the high pressure shaft to the speed of the low pressure shaft at idle conditions.
    Type: Application
    Filed: August 13, 2020
    Publication date: April 15, 2021
    Applicant: ROLLS-ROYCE PLC
    Inventor: Craig W. BEMMENT
  • Publication number: 20210071586
    Abstract: A gas turbine engine includes: an engine core, compressor system, and core shaft. A compressor exit pressure is defined as an average airflow pressure at the exit of the highest pressure compressor at cruise conditions. The core has an annular splitter and bypass flow. Stagnation streamlines around the engine circumference form a streamsurface. A fan is upstream the core with blades having leading and trailing edges, and a radially inner portion within the streamtube. A fan root entry pressure is an average airflow pressure across the radially inner portion leading edge of each fan blade at cruise conditions. An overall pressure ratio is defined as the compressor exit pressure divided by the fan root entry pressure. A bypass jet velocity is defined as the jet velocity of air flow exiting the bypass exhaust nozzle at cruise conditions. A jet velocity is in a range between 4.7 m/s and 7.7 m/s.
    Type: Application
    Filed: October 1, 2020
    Publication date: March 11, 2021
    Applicant: ROLLS-ROYCE plc
    Inventors: Pascal DUNNING, Craig W. BEMMENT
  • Publication number: 20210071572
    Abstract: A gas turbine engine (10) for an aircraft comprising: an engine core (11) comprising a turbine system comprising one or more turbines (17, 19), a compressor system comprising one or more compressors (14,15), and a core shaft (26) connecting the turbine system to the compressor system, wherein a compressor exit pressure (P30) is defined as an average pressure of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions, the engine core (11) further comprises an annular splitter (70) at which flow is divided between a core flow (A) that flows through the engine core and a bypass flow (B) that flows along a bypass duct (22), wherein stagnation streamlines (110) around the circumference of the engine (10), stagnating on a leading edge of the annular splitter (70), form a streamsurface (110) forming a radially outer boundary of a streamtube that contains all of the core flow (A); a fan (23) located upstream of the engine core (11), the fan comprising a plurality of fan bl
    Type: Application
    Filed: November 12, 2019
    Publication date: March 11, 2021
    Applicant: ROLLS-ROYCE plc
    Inventors: Pascal DUNNING, Craig W. BEMMENT