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: 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
  • Publication number: 20200400080
    Abstract: A gas turbine engine for an aircraft includes an engine core with a turbine, compressor, and core shaft connecting the two; and a fan upstream of the core with a plurality of blades extending from a hub each with a leading and trailing edge, wherein fan tip radius is between the engine centreline and each blade's leading edge outermost tip and hub radius is between the engine centreline and the hub's outer surface at each blade's leading edge radial position, the ratio of hub to tip radius between 0.2 and 0.285. A fan rotor entry temperature is the average temperature of airflow across the leading edge of each blade at cruise conditions and a fan rotor exit temperature is an average temperature of airflow across a radially outer portion of each blade at the trailing edge at cruise conditions, the ratio of entry to exit temperature between 1.11 and 1.05.
    Type: Application
    Filed: August 20, 2019
    Publication date: December 24, 2020
    Applicant: ROLLS-ROYCE PLC
    Inventors: Craig W. BEMMENT, Pascal DUNNING
  • Publication number: 20200400100
    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 comprising a plurality of fan blades extending from a hub. First and second turbine entrance and exit temperatures are defined as average temperature of airflow at the entrance or exit to the respective turbine at cruise conditions. A low pressure turbine temperature change is defined as: the ? ? first ? ? turbine ? ? entrance ? ? temperature the ? ? first ? ? turbine ? ? exit ? ? temperature .
    Type: Application
    Filed: September 17, 2019
    Publication date: December 24, 2020
    Applicant: ROLLS-ROYCE plc
    Inventors: Craig W BEMMENT, Pascal DUNNING
  • Publication number: 20200400101
    Abstract: A gas turbine engine includes an engine core including a turbine, compressor, and core shaft connecting the turbine to compressor, wherein a compressor exit temperature defined as an average temperature of airflow at exit from compressor at cruise conditions and a core entry temperature defined as an average temperature of airflow entering engine core at cruise conditions, and a fan located upstream of the engine core, wherein a fan rotor entry temperature defined as an average temperature of airflow across leading edge each fan blade at cruise conditions and fan tip rotor exit temperature 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 compressor temperature rise defined as: the ? ? compressor ? ? exit ? ? temperature the ? ? core ? ? entry ? ? temperature .
    Type: Application
    Filed: September 17, 2019
    Publication date: December 24, 2020
    Applicant: ROLLS-ROYCE plc
    Inventors: Craig W. BEMMENT, Pascal DUNNING
  • Publication number: 20200400099
    Abstract: A gas turbine engine for an aircraft includes an engine core including a turbine, compressor, and core shaft, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor at cruise conditions and a core entry temperature is defined as an average temperature of airflow entering the engine core at cruise conditions. A fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions. A core compressor temperature rise is defined as the compressor exit temperature divided by the core entry temperature. A fan root temperature rise is defined as the core entry temperature divided by the fan rotor entry temperature. A core compressor to fan root temperature rise ratio is in a specified range.
    Type: Application
    Filed: September 3, 2019
    Publication date: December 24, 2020
    Applicant: ROLLS-ROYCE plc
    Inventors: Craig W. BEMMENT, Pascal DUNNING
  • Publication number: 20200400068
    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: Application
    Filed: September 3, 2019
    Publication date: December 24, 2020
    Applicant: ROLLS-ROYCE plc
    Inventors: Craig W. BEMMENT, Pascal DUNNING
  • Publication number: 20200400081
    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 ? ? exit ? ? temperature the ? ? first ? ? turbine ? ? entrance ? ? temperature . A fan tip temperature rise is defined as: the ? ? fan ? ? tip ? ? rotor ? ? exit ? ? temperature the ? ? fan ? ? rotor ? ? entry ? ? temperature .
    Type: Application
    Filed: August 20, 2019
    Publication date: December 24, 2020
    Applicant: ROLLS-ROYCE plc
    Inventors: Craig W. BEMMENT, Pascal DUNNING
  • Publication number: 20200370435
    Abstract: A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; and a fan located upstream of the engine core, the fan including a plurality of fan blades, wherein the gas turbine engine is configured such that a flow velocity ratio between a first flow velocity at an exit of the engine core and a second flow velocity at an inlet of the engine core is within a range from around 0.82 to around 1.1 at cruise conditions.
    Type: Application
    Filed: July 16, 2019
    Publication date: November 26, 2020
    Applicant: ROLLS-ROYCE plc
    Inventor: Craig W. BEMMENT
  • Publication number: 20200370511
    Abstract: A gas turbine engine for an aircraft includes: an engine core with a turbine, a compressor, and a core shaft connecting the turbine and compressor, the engine core having an inlet upstream of the compressor and an outlet downstream of the turbine; a fan upstream of the engine core, the fan including a plurality of fan blades; a gearbox receiving an input from the core shaft and outputs drive to the fan to drive the fan at a lower rotational speed than the core shaft; and a nacelle surrounding the engine core defining a bypass duct and a bypass exhaust nozzle, wherein the gas turbine engine is configured such that an axial Mach number at the engine core inlet (which is less than around 0.7) multiplied by an axial Mach number of an exhaust airflow from the bypass exhaust nozzle is between around 0.30 to 0.56 at maximum take-off conditions.
    Type: Application
    Filed: August 13, 2019
    Publication date: November 26, 2020
    Applicant: ROLLS-ROYCE PLC
    Inventor: Craig W BEMMENT
  • Publication number: 20200370512
    Abstract: A gas turbine engine for an aircraft, the gas turbine engine comprising: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; and a nacelle surrounding the engine core and defining a bypass duct and bypass exhaust nozzle, wherein the gas turbine engine is configured such that a first velocity ratio between an axial exhaust flow velocity from the turbine and a fully expanded axial exhaust flow velocity from the bypass exhaust nozzle is greater than around 0.655 under maximum take-off conditions.
    Type: Application
    Filed: July 30, 2019
    Publication date: November 26, 2020
    Applicant: ROLLS-ROYCE plc
    Inventor: Craig W. BEMMENT
  • Publication number: 20200370481
    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: July 30, 2019
    Publication date: November 26, 2020
    Applicant: ROLLS-ROYCE plc
    Inventor: Craig W. BEMMENT
  • Patent number: 10794294
    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, OPR, 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 to OPR ratio is in a range between 4.7 m/s and 7.7 m/s.
    Type: Grant
    Filed: November 26, 2019
    Date of Patent: October 6, 2020
    Assignee: ROLLS-ROYCE plc
    Inventors: Pascal Dunning, Craig W Bemment
  • Publication number: 20200277901
    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: June 18, 2019
    Publication date: September 3, 2020
    Applicant: ROLLS-ROYCE plc
    Inventor: Craig W BEMMENT
  • Publication number: 20200263635
    Abstract: A gas turbine engine generates noise during use, and one particularly important flight condition for noise generation is take-off. A gas turbine engine has high efficiency together with low noise, in particular from the fan. The fan tip relative Mach Number at take-off at a take-off lateral reference point, defined as the point on a line parallel to and 450 m from the runway centre line where the EPNL is a maximum during take-off, is below 1.09. This results in low fan noise, along with optionally enabling a reduction in noise attenuation material.
    Type: Application
    Filed: November 14, 2019
    Publication date: August 20, 2020
    Applicant: ROLLS-ROYCE plc
    Inventors: Craig W BEMMENT, Alastair D MOORE, Robert J TELLING
  • Patent number: 10641182
    Abstract: A geared gas turbine engine for an aircraft includes: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including 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 3.6 or lower, 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 1 to around 1.3 at cruise conditions.
    Type: Grant
    Filed: June 11, 2019
    Date of Patent: May 5, 2020
    Assignee: ROLLS-ROYCE plc
    Inventor: Craig W Bemment