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.

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.

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.

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.

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.

Abstract: 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 over all 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.

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.

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.

Abstract: A gas turbine engine 10 is provided in which a fan root pressure ratio is no greater than a given value at cruise conditions. The gas turbine engine may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

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.