| Patent application number | Description | Published |
|---|
| 20090056306 | GAS TURBINE ENGINE FRONT ARCHITECTURE - A turbine engine is disclosed that includes a fan case surrounding a fan. A core is supported relative to the fan case by support structure, such as flow exit guide vanes, which are arranged downstream from the fan. The core includes a core housing having an inlet case arranged to receive airflow from the fan. A compressor case is arranged axially adjacent to the inlet case and surrounds a compressor stage. In one example, the example turbine engine includes a gear train arranged between the fan and a spool. The gear train is axially aligned and supported by the inlet case. An intermediate case is arranged axially adjacent to the compressor case. The support structure is arranged axially forward of the intermediate case. In one example, the support structure is axially aligned with the compressor case. | 03-05-2009 |
| 20090056343 | ENGINE MOUNTING CONFIGURATION FOR A TURBOFAN GAS TURBINE ENGINE - An engine mounting configuration reacts engine thrust at an aft mount. The engine mounting configuration reduces backbone bending of the engine, intermediate case distortion and frees-up space within the core nacelle. | 03-05-2009 |
| 20090081035 | GAS TURBINE ENGINE COMPRESSOR CASE MOUNTING ARRANGEMENT - A compressor case support arrangement for a gas turbine engine includes a fan section having a central axis and a compressor case for housing a compressor. An inlet case guides air to the compressor. The compressor case is positioned axially further from the fan section than the inlet case. A support member extends between the fan section and the compressor case. The support member restricts movement of the compressor case relative to the inlet case. The support member is positioned axially further from the fan section than the plumbing access area. | 03-26-2009 |
| 20090081039 | GAS TURBINE ENGINE FRONT ARCHITECTURE MODULARITY - A gas turbine engine is disclosed that includes a spool having a compressor section. An inlet case is arranged axially upstream from the compressor section. A gearbox is secured to the inlet case. The gearbox couples the spool and a fan. A sealing assembly is arranged between the inlet case and the gearbox to provide a sealed bearing compartment. The inlet case, gearbox and seal assembly provide a module. A fastening element removably secures the module to the spool. The gas turbine engine can be serviced by disconnecting the fastening element from the engine core. The gearbox and inlet case can be removed as a module from the engine core without disassembling the gearbox. | 03-26-2009 |
| 20090092487 | Systems and Methods Involving Multiple Torque Paths for Gas Turbine Engines - Systems and methods involving multiple torque paths of gas turbine engines are provided. In this regard, a representative spool assembly for a gas turbine engine, which incorporates a compressor, a turbine and a gear assembly, includes: a shaft operative to be driven by the turbine; a first spool segment operative to couple the shaft to the compressor; and a second spool segment operative to couple the shaft to the gear assembly. The first spool segment and the second spool segment are not coupled to each other. | 04-09-2009 |
| 20090183512 | MOUNTING SYSTEM FOR A GAS TURBINE ENGINE - A mounting system for a gas turbine engine includes a thrust ring and a linkage assembly. The linkage assembly is at least partially received by the thrust ring. The linkage assembly reacts at least a side load and a thrust load communicated from the thrust ring. | 07-23-2009 |
| 20090188232 | THERMAL MANAGEMENT SYSTEM INTEGRATED PYLON - A thermal management system includes at least one heat exchanger in communication with a bypass flow of a gas turbine engine. The placement of the heat exchanger(s) minimizes weight and aerodynamic losses and contributes to overall performance increase over traditional ducted heat exchanger placement schemes. | 07-30-2009 |
| 20090188234 | SHARED FLOW THERMAL MANAGEMENT SYSTEM - A thermal management system includes at least two of a multiple of heat exchangers arranged in an at least partial-series relationship. | 07-30-2009 |
| 20090188334 | Accessory Gearboxes and Related Gas Turbine Engine Systems - Accessory gearboxes and related gas turbine engine systems are provided. In this regard, a representative accessory gearbox for a gas turbine engine is operative to be driven by rotational energy extracted from the gas turbine engine and imparted to the gearbox by multiple tower shafts. | 07-30-2009 |
| 20090191045 | LOW PRESSURE TURBINE WITH COUNTER-ROTATING DRIVES FOR SINGLE SPOOL - A low pressure turbine for a gas turbine engine includes inner and outer counter-rotating rotor sets, with both said rotor sets driving a common shaft. | 07-30-2009 |
| 20090236469 | MOUNTING SYSTEM FOR A GAS TURBINE ENGINE - A mounting system for a gas turbine engine includes a mounting linkage assembly and a tangential link positioned generally transverse to the mounting linkage assembly. The mounting linkage assembly reacts at least a thrust load. The tangential link reacts at least a vertical load, a side load, and a torque load of the gas turbine engine. | 09-24-2009 |
| 20090290976 | Gearbox assembly - An assembly for a gas turbine engine includes an intermediate case and a gearbox. The intermediate case defines an annular transition duct. The gearbox comprises a plurality of lobes and is integrally formed with the annular transition duct. | 11-26-2009 |
| 20090314881 | ENGINE MOUNT SYSTEM FOR A TURBOFAN GAS TURBINE ENGINE - A mount system for a gas turbine engine includes an aft mount which reacts at least a portion of a thrust load at an engine case generally parallel to an engine axis. | 12-24-2009 |
| 20090317229 | INTEGRATED ACTUATOR MODULE FOR GAS TURBINE ENGINE - An actuator module for a gas turbine engine includes a multiple of actuators mounted within a common actuator housing. | 12-24-2009 |
| 20100218478 | Turbine engine compressor - A counter-rotating blade stage in lieu of a stator stage may compensate for relatively low rotational speed of a gas turbine engine spool. A first spool may have at least one compressor blade stage and at least one turbine blade stage. A combustor is located between the at least one compressor blade stage and the at least one turbine blade stage along a core flowpath. The at least one counter-rotating compressor blade stage is interspersed with the first spool at least one compressor blade stage. A transmission couples the at least one additional compressor blade stage to the first spool for counter-rotation about the engine axis. | 09-02-2010 |
| 20100242496 | GAS TURBINE ENGINE WITH STACKED ACCESSORY COMPONENTS - An engine accessory system for a gas turbine engine includes a first accessory component defined along an accessory axis and a second accessory component mounted to the first accessory component along the accessory axis. | 09-30-2010 |
| 20100247306 | GAS TURBINE ENGINE WITH 2.5 BLEED DUCT CORE CASE SECTION - A core case section for a gas turbine engine a multitude of discreet radial extending 2.5 bleed ducts defined in part by a structural wall. | 09-30-2010 |
| 20110127368 | SINGLE PLANE MOUNT SYSTEM FOR GAS TURBINE ENGINE - A mounting system for a gas turbine engine assembly includes a plurality of mounting links attached the gas turbine engine along a single plane transverse to the engine centerline for separating the loads from the core engine. | 06-02-2011 |
| 20120087780 | INTEGRATED ACTUATOR MODULE FOR GAS TURBINE ENGINE - An actuator module for a gas turbine engine includes a multiple of actuators mounted within a common actuator housing. | 04-12-2012 |
| 20120090329 | SYSTEMS AND METHODS INVOLVING MULTIPLE TORQUE PATHS FOR GAS TURBINE ENGINES - Systems and methods involving multiple torque paths of gas turbine engines are provided. In this regard, a representative spool assembly for a gas turbine engine, which incorporates a compressor, a turbine and a gear assembly, includes: a shaft operative to be driven by the turbine; a first spool segment operative to couple the shaft to the compressor; and a second spool segment operative to couple the shaft to the gear assembly. The first spool segment and the second spool segment are not coupled to each other. | 04-19-2012 |
| 20120099963 | ENGINE MOUNT SYSTEM FOR A TURBOFAN GAS TURBINE ENGINE - A gas turbine engine according to an exemplary aspect of the present invention includes a gear train defined along an engine centerline axis, and a spool along the engine centerline axis which drives the gear train, the spool includes a low stage count low pressure turbine. | 04-26-2012 |
| 20120121390 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low pressure compressor with four (4) stages. | 05-17-2012 |
| 20120159966 | GEARBOX ASSEMBLY - An assembly for a gas turbine engine includes an intermediate case and a gearbox. The intermediate case defines an annular transition duct. The gearbox comprises a plurality of lobes and is integrally formed with the annular transition duct. | 06-28-2012 |
| 20120167592 | MOUNTING SYSTEM FOR A GAS TURBINE ENGINE - A mounting system for a gas turbine engine includes a mounting linkage assembly and a tangential link positioned generally transverse to the mounting linkage assembly. The mounting linkage assembly reacts at least a thrust load. The tangential link reacts at least a vertical load, a side load, and a torque load of the gas turbine engine. | 07-05-2012 |
| 20120171018 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. An overall pressure ratio, which is provided by a combination of a pressure ratio across said low pressure compressor section and a pressure ratio across said high pressure compressor section, is greater than about 35. The pressure ratio across the low pressure compressor section is between about 3 and about 8 whereas the pressure ratio across the high pressure compressor section is between about 7 and about 15. | 07-05-2012 |
| 20120198815 | GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE - A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low stage count low pressure turbine. | 08-09-2012 |
| 20120198816 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low pressure compressor with four to eight (4-8) stages. | 08-09-2012 |
| 20120198817 | GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE - A gas turbine engine includes a spool along an engine centerline axis which drives a gear train. | 08-09-2012 |
| 20120216548 | SYSTEMS AND METHODS INVOLVING MULTIPLE TORQUE PATHS FOR GAS TURBINE ENGINES - Systems and methods involving multiple torque paths of gas turbine engines are provided. In this regard, a representative method for reducing overspeed potential of a turbine of a gas turbine engine includes: providing a first load to the turbine via a first torque path; providing a second load to the turbine via a second torque path; and operating the turbine such that: mechanical failure of a component defining at least a portion of the first torque path does not inhibit the second load from being applied to the turbine via the second torque path; and mechanical failure of a component defining the second torque path does not inhibit the first load from being applied to the turbine via the first torque path. | 08-30-2012 |
| 20120233982 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low pressure compressor with four to eight (4-8) stages. | 09-20-2012 |
| 20120234017 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a low pressure compressor section and a high pressure compressor section. A low pressure turbine drives the low pressure compressor section. A gear arrangement is driven by the low pressure turbine to in turn drive a fan section. A pressure ratio across the low pressure compressor section is between about 4-8, and a pressure ratio across the high pressure compressor section is between about 8-15. In a separate feature, a compressor case includes a front compressor case portion and a rear compressor case portion, with the rear compressor case portion being axially further from an inlet case than the front compressor case portion. A support member extends between the fan section and the front compressor case portion. | 09-20-2012 |
| 20120288366 | GAS TURBINE ENGINE COMPRESSOR CASE MOUNTING ARRANGEMENT - A compressor case support arrangement for a gas turbine engine includes a fan section having a central axis and a compressor case for housing a compressor. An inlet case guides air to the compressor. The compressor case is positioned axially further from the fan section than the inlet case. A support member extends between the fan section and the compressor case. The support member restricts movement of the compressor case relative to the inlet case. The support member is positioned axially further from the fan section than the plumbing access area. | 11-15-2012 |
| 20120297791 | CERAMIC MATRIX COMPOSITE TURBINE EXHAUST CASE FOR A GAS TURBINE ENGINE - A turbine exhaust case for a gas turbine engine includes a multiple of CMC turbine exhaust case struts between a CMC core nacelle aft portion and a CMC tail cone. | 11-29-2012 |
| 20120301275 | INTEGRATED CERAMIC MATRIX COMPOSITE ROTOR MODULE FOR A GAS TURBINE ENGINE - A rotor module for a gas turbine engine includes a multiple of CMC airfoil rows which extend from a common CMC drum. | 11-29-2012 |
| 20120301285 | CERAMIC MATRIX COMPOSITE VANE STRUCTURES FOR A GAS TURBINE ENGINE TURBINE - A vane structure for a gas turbine engine according includes a multiple of CMC airfoil sections integrated between a CMC outer ring and a CMC inner ring. | 11-29-2012 |
| 20120301305 | INTEGRATED CERAMIC MATRIX COMPOSITE ROTOR DISK HUB GEOMETRY FOR A GAS TURBINE ENGINE - A rotor disk for a gas turbine engine includes a CMC hub and a rail integrated with the CMC hub opposite the multiple of CMC airfoils, the rail defines a rail platform section that tapers to a rail inner bore. | 11-29-2012 |
| 20120301312 | CERAMIC MATRIX COMPOSITE AIRFOIL STRUCTURES FOR A GAS TURBINE ENGINE - A Ceramic Matrix Composite (CMC) airfoil segment for a gas turbine engine includes a box-shape fiber geometry which defines a rectilinear pressure side bond line and a rectilinear suction side bond line. | 11-29-2012 |
| 20120315130 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. An overall pressure ratio, which is provided by a combination of a pressure ratio across said low pressure compressor section and a pressure ratio across said high pressure compressor section, is greater than about 35. The pressure ratio across the high pressure compressor section is between about 7 and about 15, and a pressure ratio across the fan section is less than or equal to 1.45. | 12-13-2012 |
| 20130014489 | GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE - A gas turbine engine includes a spool along an engine centerline axis which drives a gear train. | 01-17-2013 |
| 20130014490 | GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE - A gas turbine engine includes a spool along an engine centerline axis which drives a gear train, said spool includes a low stage count low pressure turbine. | 01-17-2013 |
| 20130025257 | THREE SPOOL ENGINE BEARING CONFIGURATION - A disclosed gas turbine engine includes a core section having a low spool, intermediate spool and a high spool that rotate about a common axis. The intermediate spool is supported at an aft position by an inter-shaft bearing arrangement on the low spool. The low spool is supported for rotation in an aft position by an aft roller bearing supported on a turbine exhaust case of the gas turbine engine. The high spool is supported by a high spool aft roller bearing disposed within a high spool bearing compartment. The high spool bearing compartment is positioned within a radial space between the combustor and the axis. | 01-31-2013 |
| 20130025258 | GEARED TURBOFAN BEARING ARRANGEMENT - A geared turbofan gas turbine engine includes a fan section and a core section. The core section includes a compressor section, a combustor section and a turbine section. The fan section includes a gearbox and a fan. A low spool includes a low turbine within the turbine section and a forward connection to a gearbox for driving the fan. The low spool is supported for rotation about the axis at a forward most position by a forward roller bearing and at an aft position by a thrust bearing. | 01-31-2013 |
| 20130047623 | ACCESSORY GEARBOX BUFFER COOLING AIR PUMP SYSTEM - A buffer air pump provides pressurized cooling air for cooling components of the gas turbine engine. The buffer air pump is supported on and/or within an accessory gearbox and draws bypass air in through an inlet manifold. An impeller supported within a scroll housing pressurizes the incoming bypass air and directs the pressurized air through passages to a component requiring cooling. The buffer air pump draws in relatively cool air from the bypass flow, pressurizes the air with the impeller and sends the air through conduits and passages within the gas turbine engine to the component that requires cooling such as a bearing assembly. | 02-28-2013 |
| 20130047624 | DISTRIBUTED LUBRICATION SYSTEM - A gas turbine engine includes a spool, a gearbox having gearing driven by the spool, and a lubrication system. The lubrication system includes a first heat exchanger positioned in a first air flow path, a second heat exchanger positioned in a second air flow path, and a lubrication pump fluidically connected to both the first heat exchanger and the second heat exchanger. A first air fan is driven by the gearbox for inducing air flow through the first air flow path. A second air fan is driven by an electric motor for inducing air flow through the second air flow path. | 02-28-2013 |
| 20130067885 | FAN CASE THRUST REVERSER - A fan case of a gas turbine engine includes a fan blade containment section defined about an engine axis, a thrust reverser cascade section downstream of the blade containment section and a Fan Exit Guide Vane section downstream of the thrust reverser cascade section. | 03-21-2013 |
| 20130074517 | GAS TURBINE ENGINE MOUNT ASSEMBLY - A forward mount assembly for connecting a pylon to an intermediate case of a gas turbine engine, the forward mount assembly includes a forward mount platform, a wiffle tree assembly, and first and second A-arms. The platform is connected to the pylon and is disposed adjacent the intermediate case. The wiffle tree assembly is connected to the forward mount platform through a first ball joint. The first A-arm is connected to a first side of the intermediate case and the second A-arm is connected to a second opposing side of the intermediate case. The first and second A-arms are mounted to the forward mount platform and are mounted to opposing ends of the wiffle tree. The aforementioned arrangement allows the first and second A-arms to react a thrust load at the intermediate case substantially parallel to a centerline axis of the gas turbine engine. | 03-28-2013 |
| 20130086922 | Combined Pump System for Engine TMS AOC Reduction and ECS Loss Elimination - A compression pump for an engine is provided. The pump may include a first impeller operatively coupled to a main shaft of the engine, a first inlet configured to at least partially receive bypass airflow, and a first outlet configured to direct compressed air to a thermal management system. | 04-11-2013 |
| 20130097992 | INTEGRATED THERMAL MANAGEMENT SYSTEM AND ENVIRONMENTAL CONTROL SYSTEM FOR A GAS TURBINE ENGINE - A gas turbine engine includes a first and second pump driven by a spool. An Air-Oil Cooler downstream of the first pump. An air-air precooler is downstream of the second pump, the air-air precooler downstream of the Air-Oil Cooler. | 04-25-2013 |
| 20130098046 | INTEGRATED THERMAL SYSTEM FOR A GAS TURBINE ENGINE - An integrated thermal system for a gas turbine engine includes an Air-Oil Cooler and an Air-Air PreCooler within a housing, the Air-Air PreCooler downstream of the Air-Oil Cooler. | 04-25-2013 |
| 20130098057 | CONTROLLABLE SPEED WINDMILL OPERATION OF A GAS TURBINE ENGINE THROUGH LOW SPOOL POWER EXTRACTION - A gas turbine engine that includes at least one component geared to a spool to control a speed of the low spool during a “windmilling” condition. | 04-25-2013 |
| 20130098059 | WINDMILL OPERATION OF A GAS TURBINE ENGINE - A gas turbine engine according to an exemplary aspect of the present disclosure includes a windmill pump driven by a spool. | 04-25-2013 |
| 20130098060 | GAS TURBINE ENGINE ONBOARD STARTER/GENERATOR SYSTEM TO ABSORB EXCESS POWER - A gas turbine engine includes an Integrated Drive Generator (IDG) geared to a low spool to selectively accelerate the low spool during a transient condition. | 04-25-2013 |
| 20130098067 | CONSTANT SPEED TRANSMISSION FOR GAS TURBINE ENGINE - A gas turbine engine includes a constant speed transmission driven by a spool. | 04-25-2013 |
| 20130108413 | SECONDARY FLOW ARRANGEMENT FOR SLOTTED ROTOR | 05-02-2013 |
| 20130108445 | SPOKED ROTOR FOR A GAS TURBINE ENGINE | 05-02-2013 |
| 20130108466 | ASYMETRICALLY SLOTTED ROTOR FOR A GAS TURBINE ENGINE | 05-02-2013 |
| 20130192191 | GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION AND BEARING SUPPORT FEATURES - A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is supported by a bearing positioned at a point where the first shaft connects to a hub carrying turbine rotors associated with the second turbine section. | 08-01-2013 |
| 20130192199 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. The shaft includes a main shaft and a flex shaft having bellows. The flex shaft is secured to the main shaft at a first end and includes a second end opposite the first end. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supports the shaft relative to the inlet case. The second bearing is arranged radially outward from the flex shaft. | 08-01-2013 |
| 20130195621 | GAS TURBINE ENGINE WITH HIGH SPEED LOW PRESSURE TURBINE SECTION AND BEARING SUPPORT FEATURES - A gas turbine engine includes a very high speed low pressure turbine such that a quantity defined by the exit area of the low pressure turbine multiplied by the square of the low pressure turbine rotational speed compared to the same parameters for the high pressure turbine is at a ratio between about 0.5 and about 1.5. The high pressure turbine is supported by a bearing positioned at a point where the first shaft connects to a hub carrying turbine rotors associated with the second turbine section. | 08-01-2013 |
| 20130195645 | GEARED TURBOMACHINE ARCHITECTURE HAVING A LOW PROFILE CORE FLOW PATH CONTOUR - An exemplary geared turbomachine assembly includes a core inlet having a radially inner boundary that is spaced a first radial distance from a rotational axis of a turbomachine, and a compressor section inlet having a radially inner boundary that is spaced a second radial distance from the rotational axis. A ratio of the second radial distance to the first radial distance is of about 0.65 to about 0.9. | 08-01-2013 |
| 20130195646 | GAS TURBINE ENGINE SHAFT BEARING ARRANGEMENT - A gas turbine engine includes a shaft supported by first and second bearings for rotation relative to an inlet case. The first and second bearings are positioned within a common bearing compartment. | 08-01-2013 |
| 20130202415 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. The engine includes a combination of quantities providing beneficial operation. | 08-08-2013 |
| 20130219856 | COUNTER-ROTATING LOW PRESSURE TURBINE WITH GEAR SYSTEM MOUNTED TO MID TURBINE FRAME - A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. The gear system is mounted to a mid-turbine frame. | 08-29-2013 |
| 20130219859 | COUNTER ROTATING LOW PRESSURE COMPRESSOR AND TURBINE EACH HAVING A GEAR SYSTEM - A compressor section includes a counter rotating low pressure compressor that includes outer and inner compressor blades interspersed with one another and are configured to rotate in an opposite direction than one another about an axis of rotation. A transmission couples at least one of the outer and inner compressor blades to a shaft. A turbine section includes a counter rotating low pressure turbine having an outer rotor that includes an outer set of turbine blades. An inner rotor has an inner set of turbine blades interspersed with the outer set of turbine blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system couples at least one of the outer and inner rotors to the shaft. | 08-29-2013 |
| 20130219860 | COUNTER-ROTATING LOW PRESSURE TURBINE WITHOUT TURBINE EXHAUST CASE - A gas turbine engine includes a shaft defining an axis of rotation. An inner rotor directly drives the shaft and includes an inner set of blades. An outer rotor has an outer set of blades interspersed with the inner set of blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system is engaged to the outer rotor and is positioned upstream of the inner set of blades. | 08-29-2013 |
| 20130223983 | COUNTER ROTATING LOW PRESSURE TURBINE WITH SPLITTER GEAR SYSTEM - A gas turbine engine includes a shaft defining an axis of rotation. An outer turbine rotor directly drives the shaft and includes an outer set of blades. An inner turbine rotor has an inner set of blades interspersed with the outer set of blades. The inner turbine rotor is configured to rotate in an opposite direction about the axis of rotation from the outer turbine rotor. A splitter gear system couples the inner turbine rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. | 08-29-2013 |
| 20130223991 | GAS TURBINE ENGINE DRIVING MULTIPLE FANS - A gas turbine engine includes a core engine with a compressor section, a combustor and a turbine. The turbine drives an output shaft, and the output shaft drives at least four gears. Each of the at least four gears extends through a drive shaft to drive an associated fan rotor. | 08-29-2013 |
| 20130223992 | COUNTER-ROTATING LOW PRESSURE TURBINE WITH GEAR SYSTEM MOUNTED TO TURBINE EXHAUST CASE - A gas turbine engine includes a shaft defining an axis of rotation. An inner rotor directly drives the shaft and includes an inner set of blades. An outer rotor has an outer set of blades interspersed with the inner set of blades. The outer rotor is configured to rotate in an opposite direction about the axis of rotation from the inner rotor. A gear system couples the outer rotor to the shaft and is configured to rotate the inner set of blades at a lower speed than the outer set of blades. | 08-29-2013 |
| 20130223993 | COUNTER-ROTATING LOW PRESSURE TURBINE WITH GEAR SYSTEM MOUNTED TO TURBINE EXHAUST CASE - A gas turbine engine includes a shaft defining an axis of rotation. An outer rotor directly drives the shaft and includes an outer set of blades. An inner rotor has an inner set of blades interspersed with the outer set of blades. The inner rotor is configured to rotate in an opposite direction about the axis of rotation from the outer rotor. A gear system couples the inner rotor to the shaft and is configured to rotate the inner set of blades at a faster speed than the outer set of blades. | 08-29-2013 |
| 20130232768 | TURBINE ENGINE CASE MOUNT AND DISMOUNT - A method for mounting a gas turbine engine having a compressor section, a combustor section, a turbine section, a pylon and a rear mount bracket, includes positioning the mounting bracket between the gas turbine engine and the pylon. The mounting bracket is connected to the turbine case reacting a least a vertical load, a side load, a thrust load, and a torque load from the gas turbine engine through the mounting bracket. The mounting bracket is attached to the pylon reacting the same loads from the gas turbine engine. | 09-12-2013 |
| 20130233997 | TURBINE ENGINE CASE MOUNT - A mount for a turbine engine has a semi-circular yoke with a first leg and a second leg. The mount also has a stanchion with a cylindrical section attached to the yoke, and a conical section attached to the cylindrical section. A mounting bracket is attached to the conical section. | 09-12-2013 |
| 20130239582 | CONSTANT SPEED PUMP SYSTEM FOR ENGINE ECS LOSS ELIMINATION - A gas turbine engine has an impeller pump for delivering air to an environmental control system and a speed control pump connected to the impeller pump for driving the impeller pump at a constant speed. | 09-19-2013 |
| 20130239583 | PUMP SYSTEM FOR HPC EPS PARASITIC LOSS ELIMINATION - An engine includes a duct containing a flow of cool air and a pump system for providing air to an environmental control system. The pump system has an impeller having an inlet for receiving cool air from the duct and an outlet for discharging air to the environmental control system. | 09-19-2013 |
| 20130239584 | CONSTANT-SPEED PUMP SYSTEM FOR ENGINE THERMAL MANAGEMENT SYSTEM AOC REDUCTION AND ENVIRONMENTAL CONTROL SYSTEM LOSS ELIMINATION - A gas turbine engine has a spool, a towershaft connected to the spool, an impeller pump, and a speed control pump connected to the towershaft and to the impeller pump for driving the impeller pump at a constant speed. | 09-19-2013 |
| 20130239587 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, a gear arrangement configured to drive the fan section, a compressor section, including both a low pressure compressor section and a high pressure compressor section. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio provided by the combination of a pressure ratio across the low pressure compressor section and a pressure ratio across the high pressure compressor section is greater than about 35. The pressure ratio across a first of the low and high pressure compressor sections is between about 3 and about 8. The pressure ratio across a second of the low and high pressure compressor sections is between about 7 and about 15. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. | 09-19-2013 |
| 20130239588 | PUMP SYSTEM FOR TMS AOC REDUCTION - An engine includes a duct containing a flow of cool air and a pump system having an impeller with an inlet for receiving air from the duct and an outlet for discharging air into a discharge manifold. The discharge manifold containing at least one heat exchanger which forms part of a thermal management system. | 09-19-2013 |
| 20130259672 | INTEGRATED INLET VANE AND STRUT - A gas turbine engine case structure includes inner and outer annular case portions radially spaced from one another to provide a flow path and circumferentially arranged airfoils extend radially and interconnect the inner and outer annular case portions. The airfoils include multiple vanes and multiple strut-vanes. Each vane has a vane leading edge. Each strut-vane includes a strut-vane leading edge. The vane leading edges and strut-vane leading edges are aligned in a common plane. The vanes include a first axial length and the strut-vanes include a second axial length that is at least double the first axial length. | 10-03-2013 |
| 20130312419 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that has an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. The shaft supports a compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A geared architecture is coupled to the shaft, and a fan coupled to and rotationally driven by the geared architecture. The geared architecture includes a sun gear supported on the second end. A first bearing supports the shaft relative to the intermediate case and a second bearing supporting the shaft relative to the inlet case. The second bearing is arranged radially outward from the shaft. | 11-28-2013 |
| 20140096508 | SYSTEMS AND METHODS INVOLVING MULTIPLE TORQUE PATHS FOR GAS TURBINE ENGINES - A turbofan engine includes a fan, a compressor section, a combustor in fluid communication with the compressor section, a turbine section in fluid communication with the combustor, a shaft configured to be driven by the turbine section and coupled to the compressor section through a first torque load path, and a speed reduction mechanism configured to be driven by the shaft through a second torque load path separate from the first load path for rotating the fan. | 04-10-2014 |
| 20140157752 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A method of designing a gas turbine engine includes providing a fan section including a fan; driving the fan section via a gear arrangement; providing a compressor section, including both a first compressor and a second compressor; and driving the compressor section and the gear arrangement via a turbine section. The pressure ratio across the first compressor is greater than or equal to about 7. | 06-12-2014 |
| 20140157753 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A method of designing a gas turbine engine includes providing a fan section including a fan; driving the fan section via a gear arrangement; providing a compressor section, including both a first compressor and a second compressor; and driving the compressor section and the gear arrangement via a turbine section. An overall pressure ratio, being provided by the combination of a pressure ratio across the first compressor and a pressure ratio across the second compressor, is greater than or equal to about 35. The pressure ratio across the first compressor is greater than or equal to about 7. | 06-12-2014 |
| 20140157754 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor, with the overall pressure ratio being greater than or equal to about 35. A pressure ratio across the fan section is less than or equal to about 1.50. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. | 06-12-2014 |
| 20140157755 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. A bypass ratio which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section is greater than or equal to about 8. | 06-12-2014 |
| 20140157756 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, the geared arrangement defining a gear reduction ratio greater than or equal to about 2.6. A compressor section includes both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor. | 06-12-2014 |
| 20140157757 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. The turbine section includes a fan drive turbine configured to drive the fan section, a pressure ratio across the fan drive turbine being greater than or equal to about 5. An overall pressure ratio is provided by the combination of the first compressor and the second compressor. | 06-12-2014 |
| 20140165534 | GAS TURBINE ENGINE COMPRESSOR ARRANGEMENT - A gas turbine engine has a fan section, including a fan, a gear arrangement configured to drive the fan section, and a compressor section, including both a first compressor and a second compressor. A turbine section is configured to drive the compressor section and the gear arrangement. An overall pressure ratio is provided by the combination of the first compressor and the second compressor, with the overall pressure ratio being greater than or equal to about 35. The fan is configured to deliver a portion of air into the compressor section, and a portion of air into a bypass duct. A bypass ratio, which is defined as a volume of air passing to the bypass duct compared to a volume of air passing into the compressor section, is greater than or equal to about 8. | 06-19-2014 |
| 20140165588 | TURBO COMPRESSOR FOR BLEED AIR - A disclosed bleed air system utilizes high pressure air from a high pressure compressor to drive the turbo compressor to increase a pressure of bleed air drawn from the low pressure compressor. Air drawn from the low pressure compressor is at a lower temperature and pressure than that encountered from the high pressure compressor. The turbo compressor increases the pressure of airflow and provides that airflow into the main bleed air passage to be communicated to systems utilizing the bleed air. | 06-19-2014 |
| 20140169972 | FAN WITH INTEGRAL SHROUD - An integrally bladed rotor for use in a gas turbine engine includes a central hub; a plurality of airfoils extending from the central hub, each airfoil with a tip, a leading edge and a trailing edge; and a shroud with a metallic portion connecting to the tip of each airfoil to rotate with the airfoils. | 06-19-2014 |
| 20140174056 | GAS TURBINE ENGINE WITH LOW STAGE COUNT LOW PRESSURE TURBINE - A gas turbine engine comprises a gear train defined along an axis. A spool along the axis drives the gear train and includes a low stage count low pressure turbine. A fan i s rotatable at a fan speed about the axis and driven by the low pressure turbine through the gear train. The fan speed is less than a speed of the low pressure turbine. A core is surrounded by a core nacelle defined about the axis. A fan nacelle i s mounted at least partially around the core nacelle to define a fan bypass airflow path for a fan bypass airflow. A bypass ratio defined by the fan bypass passage airflow divided by airflow through the core is greater than about ten (10). | 06-26-2014 |
| 20140186158 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A geared architecture is arranged within the inlet case. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged axially between the inlet case flow path and the intermediate case flow path. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case. | 07-03-2014 |
| 20140205439 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a housing including an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A geared architecture is arranged within the inlet case. The geared architecture includes an epicyclic gear train. A fan is rotationally driven by the geared architecture. A shaft provides a rotational axis. A hub is operatively supported by the shaft. First and second bearings support the shaft relative to the intermediate case and the inlet case, respectively. | 07-24-2014 |
| 20140230403 | GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flow path. A first shaft supports a low pressure compressor section that is arranged axially between the inlet case flow path and the intermediate case flow path. A first bearing supports the first shaft relative to the inlet case. A second bearing supports a second shaft relative to the intermediate case. A low pressure compressor hub is mounted to the first shaft. The low pressure compressor hub extends to the low pressure compressor section between the first bearing and the second bearing. | 08-21-2014 |
| 20140248129 | LPC FLOWPATH SHAPE WITH GAS TURBINE ENGINE SHAFT BEARING CONFIGURATION - A gas turbine engine includes a core housing that includes an inlet case and an intermediate case that respectively provide an inlet case flow path and an intermediate case flowpath. A shaft provides a rotational axis. A hub is operatively supported by the shaft. A rotor is connected to the hub and supports a compressor section. The compressor section is arranged in a core flow path axially between the inlet case flow path and the intermediate case flow path. The core flowpath has an inner diameter and an outer diameter. At least a portion of inner diameter has an increasing slope angle relative to the rotational axis. A bearing is mounted to the hub and supports the shaft relative to one of the intermediate case and the inlet case. | 09-04-2014 |
| 20140250862 | REVERSE FLOW GAS TURBINE ENGINE AIRFLOW BYPASS - A gas turbine engine has a propulsor including a fan and a power turbine, an engine core aerodynamically connected to the propulsor by a transition duct, and a bypass valve in the transition duct that allows for air from the engine core to bypass the power turbine. | 09-11-2014 |
| 20140250863 | REAR MOUNTED REVERSE CORE ENGINE THRUST REVERSER - In one embodiment, a gas turbine engine for mounting to a rear of an aircraft fuselage has a propulsor that rotates on a first axis, and an engine core including a compressor section, a combustor section, and a turbine section, with the turbine section being closer to the propulsor than the compressor section. The engine core is aerodynamically connected to the propulsor and has a second axis. A nacelle is positioned around the propulsor and engine core. The nacelle is attached to the wing of the aircraft. A downstream end of the nacelle has at least one pivoting door with an actuation mechanism to pivot the door between a stowed position and a vertical deployed position in which the door inhibits a flow to provide a thrust reverse of the flow. | 09-11-2014 |
| 20140252160 | REVERSE FLOW GAS TURBINE ENGINE REMOVABLE CORE - An engine mounting arrangement includes a propulsor mounted to an aircraft wing, and an engine core aerodynamically connected to the propulsor and positioned rearward of the propulsor. | 09-11-2014 |
| 20140252167 | REVERSE CORE ENGINE THRUST REVERSER FOR UNDER WING - A gas turbine engine for mounting under a wing of an aircraft has a propulsor that rotates on a first axis, and an engine core including a compressor section, a combustor section, and a turbine section, with the turbine section being closer to the propulsor than the compressor section. The engine core is aerodynamically connected to the propulsor and has a second axis. A nacelle is positioned around the propulsor and engine core. The nacelle is attached to the wing of the aircraft. A downstream end of the nacelle has at least one pivoting door with an actuation mechanism to pivot the door between a stowed position and a horizontal deployed position in which the door inhibits a flow to provide a thrust reverse of the flow. | 09-11-2014 |
| 20140294589 | ASYMMETRICALLY SLOTTED ROTOR FOR A GAS TURBINE ENGINE - A spool for a gas turbine engine includes at least one rotor disk defined along an axis of rotation and at least one rotor ring defined along the axis of rotation, with the rotor ring being in contact with the rotor disk. The rotor disk and rotor ring are contoured to define a smooth rotor stack load path. | 10-02-2014 |
