Abstract: An aspect includes a system including a high compressor of a gas turbine engine having a ratio of a cold-rotor build clearance to a span between 0.7% and 7%. The cold-rotor build clearance is defined for a plurality of rotor blades of the high compressor with respect to an engine casing assembly interior surface of the high compressor, and the span is defined as a gap between a rotor disk of the high compressor and the engine casing assembly interior surface of the high compressor for at least a last two stages of the high compressor closest to a combustor section of the gas turbine engine. The system also includes at least two bowed rotor management systems for the gas turbine engine to prevent damage to the rotor blades for a bowed rotor condition of the high compressor under a plurality of operating conditions.

Abstract: A bowed rotor start mitigation system for a gas turbine engine is provided. The bowed rotor start mitigation system includes a controller operable to receive a vibration input indicative of an actual vibration level of the gas turbine engine and a speed input indicative of a rotor speed of the gas turbine engine. The controller generates a bowed rotor start mitigation request based on comparing the actual vibration level to a modeled vibration level at the rotor speed.

Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed, where the controller modulates a duty cycle of the discrete starter valve via pulse width modulation.

Abstract: A system for controlling a start sequence of a gas turbine engine includes an electronic engine control system, a thermal model, memory, a model for determining a time period (fmotoring), and a controller. The thermal model synthesizes a heat state of the gas turbine engine. The memory records the current heat state at shutdown and a shutdown time of the gas turbine engine. The model for determining the time period is for motoring the gas turbine engine at a predetermined speed Ntarget that is less than a speed to start the gas turbine engine, where tmotoring is a function of the heat state recorded at engine shutdown and an elapsed time of an engine start request relative to a previous shutdown time. The controller modulates a starter valve to maintain the gas turbine engine within a predetermined speed range of NtargetMin to NtargetMax for homogenizing engine temperatures.

Abstract: An aspect includes a system including a high compressor of a gas turbine engine having a ratio of a cold-rotor build clearance to a span between 0.7% and 7%. The cold-rotor build clearance is defined for a plurality of rotor blades of the high compressor with respect to an engine casing assembly interior surface of the high compressor, and the span is defined as a gap between a rotor disk of the high compressor and the engine casing assembly interior surface of the high compressor for at least a last two stages of the high compressor closest to a combustor section of the gas turbine engine. The system also includes at least two bowed rotor management systems for the gas turbine engine to prevent damage to the rotor blades for a bowed rotor condition of the high compressor under a plurality of operating conditions.

Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed.

Abstract: A bowed rotor start mitigation system for a gas turbine engine is provided. The bowed rotor start mitigation system includes a controller operable to receive a vibration input indicative of an actual vibration level of the gas turbine engine and a speed input indicative of a rotor speed of the gas turbine engine. The controller generates a bowed rotor start mitigation request based on comparing the actual vibration level to a modeled vibration level at the rotor speed.

Abstract: A system for starting a gas turbine engine of an aircraft is provided. The system includes a pneumatic starter motor, a discrete starter valve switchable between an on-state and an off-state, and a controller operable to perform a starting sequence for the gas turbine engine. The starting sequence includes alternating on and off commands to an electromechanical device coupled to the discrete starter valve to achieve a partially open position of the discrete starter valve to control a flow from a starter air supply to the pneumatic starter motor to drive rotation of a starting spool of the gas turbine engine below an engine idle speed.

Abstract: A disclosed gas turbine engine includes a case for a case for a compressor section including a bleed air slot. The bleed air slot includes an inlet having a first area radially inward of an outlet having a second area with the second area being greater than the first area. The bleed air slot further includes a center portion disposed along a radial line extending from an axis of the engine and an elongated portion extending from the opening at an angle relative to a line normal to the radial line.

Abstract: An example method of cooling a compressor of a gas turbine includes, among other things, diverting a flow from a compressor, and directing the flow at the compressor in a direction, the direction having a circumferential component and an axial component.

Abstract: A disclosed gas turbine engine includes a case for a case for a compressor section including a bleed air slot. The bleed air slot includes an inlet having a first area radially inward of an outlet having a second area with the second area being greater than the first area. The bleed air slot further includes a center portion disposed along a radial line extending from an axis of the engine and an elongated portion extending from the opening at an angle relative to a line normal to the radial line.

Abstract: A multiple of circumferential grooves within said arcuate engine casing and an abradable material located radial inboard of the multiple of circumferential grooves.

Abstract: A multiple of circumferential grooves within said arcuate engine casing and an abradable material located radial inboard of the multiple of circumferential grooves.

Abstract: A turbine rotor for a turbine engine includes multiple rotor disks having rotor blades mounted about the circumference of each of the rotor disks. A fluid seal extends about the circumference of each rotor disk in close proximity to a stationary component of the rotor to separate the space between the rotor blades and a stationary component into separate cavities. The fluid seal includes a seal land on the rotor disk and a brush seal extending from the stationary component. A plurality of disk land segments and a plurality of blade land segments form the seal land. The disk seal segments are located on the rotor disk between the rotor blades. The blade land segments are located on the rotor blades. After the rotor blades are assembled, the blade land segments and the disk land segments fit together to form a segmented ring-like seal land around the circumference of the rotor disk.

Abstract: A gas turbine engine includes a turbine rotor and a compressor, which provides discharge air. A nozzle, typically referred to a tangential on-board injector (TOBI), is arranged near the rotor to deliver the discharge air near the turbine rotor for cooling it. The TOBI receives pollution air leaking past seals within the gas turbine engine. The TOBI swirls the discharge air and the pollution air before it reaches the turbine rotor. The TOBI provides multiple passages separated by vanes. At least some of the passages include discharge inlets and outlets for carrying the discharge air from the compressor to the turbine rotor. Typically, several of the passages are unused and blocked. However, the example arrangement provides a pollution inlet and outlet in at least one of the normally unused, blocked passages.