US Pat. No. 9,517,843

GENERATOR FOR FLIGHT VEHICLE

Rolls-Royce North America...

1. A flight vehicle comprising:
a fuselage including a nose and a tail spaced apart from the nose,
a gas turbine engine located in a space formed in the fuselage and located between the nose and the tail, and
a generator unit coupled to the gas turbine engine via a generator shaft, at least a portion of the generator shaft located
exterior to the generator unit and the gas turbine engine, to cause a portion of the generator unit to rotate therewith, the
generator unit being located in spaced-apart relation to the gas turbine engine between the nose and the gas turbine engine,

wherein the flight vehicle further includes an air inlet duct coupled to the gas turbine engine and formed to include an air
passageway therein and an air inlet defining an aperture opening into the air passageway to communicate air surrounding the
flight vehicle through the aperture, through the air passageway, and into the gas turbine engine for combustion,

wherein a portion of the generator unit extends into and lies in the air passageway.

US Pat. No. 9,518,510

BLOWER FOR USE WITH AIR PARTICLE SEPARATOR

ROLLS-ROYCE NORTH AMERICA...

1. An apparatus, comprising:
an air particle separator having an air inlet capable of receiving an air carrying a particulate matter, the air particle
separator also having a first air outlet and a second air outlet, each structured to flow portions of the air received by
the air inlet, the first air outlet structured to flow a higher concentration of particulate matter than the second air outlet,
the second air outlet in fluid communication with an air intake of an internal combustion engine; and

an air moving device in fluid communication with the first air outlet having a rotatable air moving member capable of producing
a pressure change to facilitate movement of the particulate matter carried by the air through the first air outlet, the air
moving device structured to receive a rotating force at a force receiving portion located radially outward of a portion of
the air moving member, wherein the rotatable air moving member is positioned in an air duct fluidly coupled with the first
air outlet, the air moving device further includes an annular stator arranged externally around the air duct and aligned axially
with the air moving device; and

the stator is configured to cause the rotating force to be applied to the force receiving portion of the air moving device
through electromagnetic interactions to cause rotation of the rotatable air moving member.

US Pat. No. 9,556,750

COMPARTMENTALIZATION OF COOLING AIR FLOW IN A STRUCTURE COMPRISING A CMC COMPONENT

Rolls-Royce North America...

1. A gas turbine engine structure comprising:
a static metal component;
a CMC component spaced apart from the static metal component and separated therefrom by a cavity having sections with respective
passages for receiving cooling air into the cavity through the static metal component and removing cooling air from the cavity
through the CMC component; and

at least one rope seal located between the static metal component and the CMC component,
the rope seal dividing the cavity into the sections to thereby compartmentalize the cavity and cooling air flow is ensured.

US Pat. No. 9,516,788

HEAT GENERATING ELECTRIC DEVICE HAVING CARBON NANOTUBE HOUSING AND THERMAL STORAGE MEDIUM

Rolls-Royce North America...

1. An apparatus comprising:
an electric device including a heat generating electric component, a thermal storage device, and a housing, the housing being
structured to at least partially envelop the heat generating electric component and being of a composite construction that
includes a plurality of carbon nanotubes useful to dissipate thermal energy produced from the heat generating electric component,
and the thermal storage device including a phase change material operative to at least partly change from a first phase to
a second phase to cool the electric device, wherein the electric device includes a stator and a rotor that includes a plurality
of magnets, and wherein the phase change material is located internal to the rotor and not radially outward of the plurality
of magnets of the rotor.

US Pat. No. 9,482,451

ADAPTIVE TRANS-CRITICAL CO2 COOLING SYSTEMS FOR AEROSPACE APPLICATIONS

Rolls-Royce Corporation, ...

18. An aircraft comprising:
a turbine engine; and
a cooling system for the aircraft comprising:
a heat exchanger through which a refrigerant flows, in which heat is rejected to a fluid;
a set of valves arranged to:
direct the refrigerant through a first circuit having a fluid expansion device; direct the refrigerant through a second circuit
having a fluid expansion machine; or

direct the refrigerant through both circuits, based on ambient conditions; and
a third circuit having a heater coupled to the first and second circuits, the third circuit including a second expansion device
and the heater configured to receive waste heat from the aircraft, wherein the refrigerant is directed through the heater
and the second expansion machine such that the second expansion machine is configured to extract energy from heated refrigerant
that exits from the heater.

US Pat. No. 9,193,311

AIRCRAFT AND SYSTEM FOR SUPPLYING ELECTRICAL POWER TO AN AIRCRAFT ELECTRICAL LOAD

Rolls-Royce North America...

1. An aircraft, comprising:
a wing;
a fuselage coupled to the wing;
an engine coupled to at least one of the fuselage and the wing;
an electrical load of a high energy device associated with the aircraft during flight operations;
a generator coupled to the engine and configured to generate electrical power at a voltage above 270 volts for the electrical
load of the high energy device during flight operations;

first and second conductors electrically disposed between the electrical load and the generator;
first and second conduits arranged along each other and configured to house the respective first and second conductors; and
a dielectric gas disposed in at least one of the first and second conduits;
wherein the at least one of the first and second conduits is configured to envelop the conductor in the dielectric gas.

US Pat. No. 9,145,213

GAS TURBINE ENGINE CONFIGURATION INTERFACE

Rolls-Royce North America...

1. An apparatus for configuring a gas turbine engine, comprising:
a gas turbine engine operation module structured to receive a requested engine configuration and control variable mechanical
features of the gas turbine engine to achieve the requested engine configuration, the gas turbine engine operation module
structured to provide a plurality of candidate operation modes for selection corresponding to a plurality of engine configurations,
the plurality of candidate operation modes including:

a nominal operation mode in which specific fuel consumption (SFC) is non-optimal in favor of rapid thrust response;
an optimum SFC mode in which a cooling feature is in a minimum cooling setting;
a bolster thrust operation mode in which the gas turbine engine is configured to produce a spoiled thrust through a variable
mechanical feature that can be quickly reconfigured to produce a non-spoiled thrust;

a flow holding operation mode in which fan speed is held and flow through an internal duct remains substantially constant
while the gas turbine engine produces varying thrust levels through a configuration of variable mechanical features; and

a power off-take mode in which engine is placed in a state that accommodates a heavy increase in demanded power without destabilizing
the engine.

US Pat. No. 9,371,811

METHODS AND SYSTEMS FOR OPERATING A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A method of operating a turbine engine, comprising:
operating an electrical machine to rotate a spool of the gas turbine engine;
determining a torque of the electrical machine during rotation of the spool using a starter motor current measurement;
controlling the rotation of the spool based on the torque;
simultaneously measuring a speed of rotation of the spool and the starter motor current measurement to determine real time
speed versus torque, wherein the speed and torque are maintained as a record of speed versus torque;

augmenting an engine health monitoring data set with the measured torque and speed; and
utilizing the engine health monitoring data set to do at least one of:
reduce turbine engine start up time;
reduce turbine engine start up temperature;
identify turbine engine failure conditions;
adjust for turbine engine aerodynamic changes; or
adjust turbine engine light-off timing.

US Pat. No. 9,260,976

ENGINE HEALTH MONITORING AND POWER ALLOCATION CONTROL FOR A TURBINE ENGINE USING ELECTRIC GENERATORS

Rolls-Royce North America...

12. A method for controlling the allocation of power extracted from a plurality of shafts of a turbine engine, the shafts
having one or more electrical machines coupled to each of the shafts, the method comprising:
receiving, via feedback control of the turbine engine, data relating to one or more turbine engine operating conditions including
fuel flow, temperature, pressure, and/or speed;

generating an assessment of the health of the turbine engine based on changes in the one or more operating conditions over
time;

using a model-based algorithm to predict a change in the health of the turbine engine based on the operating conditions and
a healthy engine profile, wherein the change in health of the turbine engine is indicative of a wear characteristic of the
turbine engine;

determining an optimal power extraction allocation between the electrical machines based on the predicted change in engine
health, the operating conditions, and one or more optimization parameters including fuel efficiency, engine performance, and/or
engine reliability; and

controlling the electrical machines to implement the optimal power extraction by configuring the one or more electrical machines
coupled to each of the plurality of shafts to generate a part of a total amount of electrical energy.

US Pat. No. 9,228,497

GAS TURBINE ENGINE WITH SECONDARY AIR FLOW CIRCUIT

Rolls-Royce Corporation, ...

1. A gas turbine engine, comprising:
a compressor having an impeller;
a diffuser having a plurality of diffuser vanes;
wherein the impeller is a centrifugal impeller, and wherein the diffuser is a radial diffuser;
wherein the diffuser forms a flowpath downstream of the impeller;
wherein the diffuser vanes extend across the flowpath; and
wherein at least one of the diffuser vanes has a first opening extending through the diffuser vanes and across the flowpath;
a combustor in fluid communication with the compressor;
a turbine in fluid communication with the combustor; and
a secondary flow circuit operative to deliver secondary air flow to the impeller for controlling a temperature of a portion
of the impeller, wherein the secondary air flow is delivered to the impeller from across the flowpath through the first opening;
and

one or more walls defining a cavity separate from the flowpath, the one or more walls further having an opening therein configured
to supply the secondary air flow from the cavity to at least one diffuser vane, wherein the impeller includes a plurality
of blades and a back face opposite the plurality of blades, further comprising a static structure spaced apart from the back
face and configured to direct the secondary air flow from a radially outer tip portion of the impeller radially inward along
the back face of the impeller.

US Pat. No. 9,080,448

GAS TURBINE ENGINE VANES

Rolls-Royce North America...

1. A vane segment for a gas turbine engine, comprising:
an outer shroud;
an inner shroud;
a composite airfoil having a span slidably disposed between said outer shroud and said inner shroud, the entirety of the span
being configured to move in a flowpath direction, the airfoil being configured at its ends to slide against said outer shroud
and inner shroud from a first position to a second position, said airfoil having a passage extending between said outer shroud
and said inner shroud; and

a spoke extending through said passage between said outer shroud and said inner shroud, wherein said airfoil is moveable in
the flowpath direction from the first position to the second position, and in the second position the spoke is operative to
limit the movement of said airfoil.

US Pat. No. 9,567,857

TURBINE SPLIT RING RETENTION AND ANTI-ROTATION METHOD

Rolls-Royce North America...

1. A disk arrangement for a gas turbine engine comprising:
a disk with a bayonet feature;
a cover plate with a bayonet feature;
a split retainer ring;
an anti-rotation peg having a first surface for engaging the split retainer ring,
a second surface for engaging the cover plate, and a third surface for engaging the disk; and
a radial retention peg,
wherein the anti-rotation peg includes at least one of:
a radially-oriented tab that projects from a surface and engages an opening in the retainer ring; or
a stop that engages the retainer ring and prevents the retainer ring from rotating relative to the disk and the cover plate.

US Pat. No. 9,421,733

MULTI-LAYER CERAMIC COMPOSITE POROUS STRUCTURE

Rolls-Royce North America...

1. An article of manufacture, comprising:
a first ceramic matrix composite (CMC) sheet having a plurality of flow passages therethrough;
a permeable structure layer bonded to the first CMC sheet, wherein the permeable structure layer comprises an open-cell foam
layer;

a second CMC sheet bonded to the permeable structure layer;
another open-cell foam layer bonded to the second CMC sheet; and
a third CMC sheet bonded to the another open-cell foam layer.

US Pat. No. 9,458,726

DOVETAIL RETENTION SYSTEM FOR BLADE TRACKS

Rolls-Royce Corporation, ...

1. A turbine shroud comprising
a metallic inner carrier formed to include a plurality of apertures,
a plurality of ceramic blade track segments each including an arcuate runner arranged radially inward from the metallic inner
carrier and a dovetail post extending radially outward from the arcuate runner through one of the apertures formed in the
metallic inner carrier, and

a plurality of segment retainers arranged radially outward from the metallic inner carrier, each segment retainer formed to
include a retention channel mated with a portion of a corresponding dovetail post, and each segment retainer sized to block
movement of the corresponding dovetail post through the aperture so that the plurality of ceramic blade track segments are
coupled to the metallic inner carrier.

US Pat. No. 9,249,684

COMPLIANT COMPOSITE COMPONENT AND METHOD OF MANUFACTURE

Rolls-Royce Corporation, ...

1. A composite component comprising
a bonded portion made from a ceramic matrix composite material and
an un-bonded portion made from the ceramic matrix composite material and coupled to the bonded portion to move relative to
the bonded portion in response to application of a load to cause the composite component to deform in a controlled manner
without fracture of the composite component,

wherein the un-bonded portion includes a first un-bonded section coupled to the bonded portion to move relative to the bonded
portion in response to application of the load, a second un-bonded section coupled to the bonded portion to move relative
to the bonded portion and the first un-bonded section in response to application of the load, and a third un-bonded section
coupled to the bonded portion to move relative to the bonded portion, the first un-bonded section, and the second un-bonded
section in response to application of the load.

US Pat. No. 9,046,056

INLET PARTICLE SEPARATOR SYSTEM FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. An inlet particle separator system for a gas turbine engine, comprising:
a separator inlet structured to receive a vitiated air flow;
an inertial particle separator in fluid communication with the separator inlet positioned upstream of a fan, wherein the inertial
particle separator includes an outer flowpath wall and an inner flowpath wall that define an inlet pathway to receive the
vitiated air flow and discharge a cleaned air flow and a scavenge flow, a first portion of the inlet pathway extending radially
outward to an apex and a second portion of the inlet pathway extending radially inward from the apex such that scavenge flow
is directed radially outward towards the outer flowpath wall;

a flow splitter having an intermediate outer flowpath wall and an intermediate inner flowpath wall positioned downstream of
the inertial particle separator;

a scavenge flowpath defined by the intermediate outer flowpath wall of the flow splitter and the outer flowpath wall of the
inlet pathway positioned to receive the scavenge flow from the inertial particle separator;

a cleaned air flowpath defined by the intermediate inner flowpath wall of the flow splitter and the inner flowpath wall of
the inlet pathway positioned to receive the cleaned air flow from the inertial particle separator;

a variable output ejector in fluid communication with said flowpaths, wherein said variable output ejector is structured to
provide a variable draw on the scavenge flowpath using a portion of the cleaned air flow as a motive fluid for operating said
variable output ejector; and

a cleaned air engine inlet in fluid communication with the cleaned air flowpath, wherein the cleaned air engine inlet is structured
to receive the balance of the cleaned air flow and to direct the balance into the gas turbine engine as at least one of a
core flow and a fan bypass flow of the gas turbine engine.

US Pat. No. 9,435,267

INTEGRATED HEALTH MANAGEMENT APPROACH TO PROPULSION CONTROL SYSTEM PROTECTION LIMITING

Rolls-Royce North America...

1. An engine system comprising:
a controller system that controls demand on a first component of an engine, wherein the controller system includes at least
one controller and is configured to:

access a first set of prognostic data about the first component, wherein the first set of prognostic data includes a remaining
lifespan approximation of the first component operating at a present operating condition;

identify a temporal length of an engine procedure operating at the present operating condition;
decrease a maximum hard current limit of the first component to increase the remaining lifespan approximation of the first
component beyond the temporal length; and

implement the current limit constant associated with the first component so that the first component does not fault during
engine operation.

US Pat. No. 9,410,478

INTERCOOLED GAS TURBINE WITH CLOSED COMBINED POWER CYCLE

Rolls-Royce North America...

13. A turbine engine comprising:
a fan that provides an air flow to the turbine engine as compressor intake air and as compressor bypass air;
a first stage compressor positioned to receive the compressor intake air and output a first stage compressed air;
a boiler positioned to cool the first stage compressed air using a fluid;
a second stage compressor positioned to receive the cooled first stage compressed air;
a pump configured to pump the fluid as a liquid into the boiler, such that the fluid in the boiler extracts energy from the
first stage compressed air and causes the cooling of the first stage compressed air;

a closed cycle turbine configured to receive the fluid from the boiler and extract auxiliary energy therefrom; and
a heat exchanger positioned within an exhaust stream emitted from a turbine assembly, and positioned to receive the fluid
from the boiler and heat the fluid with the exhaust stream prior to entering the closed cycle turbine.

US Pat. No. 9,506,356

COMPOSITE RETENTION FEATURE

Rolls-Royce North America...

14. A retention feature for use in a gas turbine engine comprising
a ceramic post including a body adapted to be coupled to a turbine engine component and a head coupled to the body, the head
formed to include a space,

a first insert arranged in the space, and
a first braze layer extending from the ceramic post to the first insert to bond the first insert to the ceramic post,
wherein the first insert is U-shaped such that a gap exists between two opposed sides thereof.

US Pat. No. 9,534,537

PHASE CHANGE MATERIAL COOLING SYSTEM FOR A VEHICLE

Rolls-Royce North America...

1. An apparatus comprising:
a vehicle having a gas turbine engine for providing a motive force of the vehicle coupled to an outer surface of the vehicle
in a first location;

a first pod coupled to the outer surface of the vehicle in a second location, the first pod having a body extending from a
leading edge to a trailing edge, the body defining a cavity therein, wherein the first location of the gas turbine engine
is different from the second location of the first pod;

a first work providing device positioned in the cavity of the first pod, the first work providing device configured to generate
electrical power in response to a change in a pressure of a first compressed fluid generated by a compressor stage of the
gas turbine engine;

a first thermal conditioning system powered by the first work providing device, the first thermal conditioning system including
a thermal energy transfer system having a first working fluid and a phase transition material capable of transferring thermal
energy with one or more components of the vehicle, wherein thermal energy is transferred between the first thermal energy
transfer system and the phase transition material via the first working fluid;

a second thermal conditioning system powered by the first work providing device, the second thermal conditioning system including
the phase transition material and a second thermal energy transfer system having a second working fluid configured to transfer
thermal energy from the phase transition material;

an electrically driven device positioned in an interior of the vehicle, the electrically driven device having a heat-producing
component which generates thermal energy, the heat-producing component being in thermal communication with the phase transition
material; and

a third thermal energy transfer system having a third working fluid, the third thermal energy transfer system providing thermal
communication between the phase transition material and the heat-producing component, wherein the third working fluid withdraws
thermal energy from the heat-producing component and delivers the thermal energy to the phase transition material.

US Pat. No. 9,435,266

SEALS FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A sealing assembly comprising
a support having a support-seal surface,
an engine component having a component-seal surface, the engine component mounted so that the component-seal surface is arranged
in spaced-apart confronting relation with the support-seal surface to define a gap between the support and the engine component
that grows and shrinks based on the temperature of the support and the engine component, and

a seal adapted to block gasses from passing through the gap between the support and the engine component, the seal including
a mount ring coupled to the support and spaced apart from the engine component, a ceramic tadpole gasket having a compressible
head held in contact with the engine component by the mount ring and a flat body extending from the compressible head into
a retention channel formed by the mount ring, and a retainer ring arranged to extend into the retention channel of the mount
ring to trap the flat body of the tadpole gasket in the retention channel formed by the mount ring so that the tadpole gasket
is coupled to the mount ring,

wherein the mount ring includes a spring portion having a U-shaped cross-section and the spring portion is configured to push
the compressible head of the ceramic tadpole gasket into contact with the engine component, the U-shaped cross-section of
the spring portion included in the mount ring having legs that extend generally parallel to the component seal surface of
the engine component.

US Pat. No. 9,482,156

VEHICLE RECUPERATOR

Rolls-Royce Corporation, ...

1. A system comprising:
an engine having:
a compressor and a turbine coupled together with a flow path there between by which a fluid flow stream is flowable from the
compressor to the turbine, the compressor having a compressor discharge through which the fluid flow stream exits the compressor,
the turbine being configured to generate power;

a combustor located between the compressor and the turbine, the combustor being configured to provide heat to the fluid flow
stream before entering the turbine;

at least one recuperator disposed downstream of the compressor, the recuperator being configured to transfer heat with the
fluid flow stream en route to the combustor;

an electric generator operatively coupled to the turbine to receive at least a portion of the power generated by the turbine,
the electric generator further being in communication with the at least one recuperator; and

a controller configured to operate the engine in a low power mode, a high power mode, and at least one intermediate speed/power
mode, the controller configured to selectively distribute energy from the electric generator between the at least one recuperator
and a load, the controller configured to direct energy to the at least one recuperator while reducing a fuel flow to the engine
when in the intermediate speed/power mode to maintain a speed/power setting of the engine;

wherein the recuperator is configured in an off condition in the low power mode.

US Pat. No. 9,297,304

GAS TURBINE ENGINE SYSTEM WITH BLEED AIR POWERED AUXILIARY ENGINE

Rolls-Royce North America...

1. A gas turbine engine system, comprising:
a compressor forming a source of pressurized air;
a first combustor in fluid communication with said compressor and configured to receive working fluid consisting of pressurized
air directly from the compressor prior to the working fluid passing through a heat addition component or an expansion component;

a second combustor in fluid communication with said compressor in parallel with said first combustor and configured to receive
working fluid consisting of pressurized air directly from the compressor prior to the working fluid passing through a heat
addition component or an expansion component;

a first turbine in fluid communication with and downstream of said first combustor;
a second turbine in fluid communication with said second combustor; and
a third turbine in fluid communication with said second combustor, wherein said third turbine is on a different spool than
said second turbine;

wherein the second combustor, the second turbine, and the third turbine are configured as an auxiliary engine, wherein the
auxiliary engine does not include a compressor.

US Pat. No. 9,109,539

TURBINE BASED COMBINED CYCLE ENGINE

Rolls-Royce North America...

8. An apparatus comprising:
an aircraft engine operable to provide power below and above sonic aircraft speeds, the aircraft engine including a gas turbine
engine core and a ramburner, the gas turbine engine core including a core combustor; and

a gas turbine engine inlet valve operable to provide a first airflow and a second airflow, the first airflow routed to the
gas turbine engine core, the second airflow routed to the ramburner, the first airflow formed by positioning a leading edge
of the gas turbine engine inlet valve axially forward of an apex of an engine centerbody, wherein the apex is positioned forward
of a compressor, wherein an annular flow space is formed between the gas turbine engine inlet valve and the centerbody;

wherein the ramburner is structured to receive products of combustion from the core combustor and to receive the second airflow
provided by the gas turbine engine inlet valve.

US Pat. No. 9,458,855

COMPRESSOR TIP CLEARANCE CONTROL AND GAS TURBINE ENGINE

Rolls-Royce North America...

1. A compressor, comprising:
a rotating compressor blade having a blade tip;
an outer compressor case and an inner compressor case, the inner compressor case having a blade track disposed opposite the
blade tip; and

a tip clearance control system including
a fluid impingement structure having a plurality of impingement openings configured to impinge a fluid onto the inner compressor
case;

a valve in communication with the fluid impingement structure to control flow of the fluid received from a diffuser downstream
from the compressor such that the fluid from the diffuser travels radially through an opening of the diffuser to a cooler,
wherein

the cooler is in fluid communication with the fluid impingement structure to cool the fluid, wherein the tip clearance control
system is configured to control a clearance between the blade tip and the blade track by impinging the cooled fluid that has
been cooled by the cooler onto the inner compressor case and wherein the valve is configured to modulate the cooled fluid
between different flow amounts of the cooled fluid, and wherein the diffuser is in fluid communication with the fluid impingement
structure through the cooler and the valve and a distribution channel; and

a first support structure and a second support structure both configured for radial flexibility, wherein the first support
structure and the second support structure absorb a thermal growth differential between the inner compressor case and the
outer compressor case resulting from impingement of the cooled fluid onto the inner compressor case, wherein the first support
structure includes a first connector attached to the outer compressor case and a second connector attached to the inner compressor
case, wherein the first support structure is angled between the first connector and the second connector, wherein the second
support structure includes a third connector attached to the outer compressor case, and wherein the second support structure
has a bended knee shape and connects the outer compressor case to the inner compressor case.

US Pat. No. 9,551,299

CHECK VALVE FOR PROPULSIVE ENGINE COMBUSTION CHAMBER

Rolls-Royce Corporation, ...

1. A combustion chamber comprising
an inner wall; and
an outer wall surrounding the inner wall;
the inner wall having a plurality of effusion holes in fluid communication with an inside of the combustion chamber, the outer
wall having a plurality of cooling side holes in fluid communication with a cooling source;

the inner wall and the outer wall defining a flow passage therebetween that fluidly connects one or more of the cooling side
holes with one or more of the effusion holes,

the flow passage having a geometric configuration to permit forward flow of gases through the flow passage from the cooling
source to the inside of the combustion chamber and to restrict reverse flow of gases through the flow passage from the inside
of the combustion chamber to the cooling source, and

the permitted flow rate being greater than the restricted flow rate,
wherein the geometric configuration includes a backflow inhibiting portion formed by a forwardly directed pointed conical
portion flanked by a pair of deflector portions and each deflector portion is defined by a forwardly opening curved wall.

US Pat. No. 9,261,019

VARIABLE CYCLE GAS TURBINE ENGINE

Rolls-Royce North America...

1. A variable cycle gas turbine engine, comprising:
a compressor configured to compress a core gas flow;
a combustor in fluid communication with the compressor and configured to combust the core gas flow;
a primary turbine drivingly coupled to the compressor and configured to receive the core gas flow, wherein the primary turbine
is configured to drive the compressor;

an auxiliary turbine drivingly coupled to the compressor;
a valve configured to selectively direct a portion of the core gas flow to the auxiliary turbine; and
a controller structured to selectively operate the valve such that the portion of the core gas flow is directed to the auxiliary
turbine,

wherein the auxiliary turbine is configured to extract power from the portion of the core gas flow and supply the power to
the compressor when the valve is open through action of the controller;

wherein the controller is configured to selectively operate the valve according to at least one of a look-up table, a rate
schedule, sensed or calculated engine parameters, engine inlet conditions, aircraft speed and power lever angle.

US Pat. No. 9,091,229

AIRCRAFT POWERPLANT

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine having a core flow passage, a fan bypass passage, and a third stream bypass passage that is split from
the fan bypass passage to divert a portion of the working fluid traversing the fan bypass away from the fan bypass, the third
stream bypass passage having a flow divider that defines a first branch and a second branch, the first and second branches
configured to convey a portion of working fluid traversing the third stream bypass passage; and

a heat exchanger located in thermal communication with the working fluid conveyed through the first branch of the third stream
bypass passage such that the working fluid exchanges heat with the heat exchanger.

US Pat. No. 9,077,221

GAS TURBINE ENGINE

Rolls-Royce North America...

1. A gas turbine engine, comprising:
a cold section stage being at least one of including a fan blade stage and a compressor blade stage, wherein the fan blade
stage includes a wheel and a plurality of blades and a shroud circumferentially surrounding the plurality of blades;

a combustor in fluid communication with the at least one of the fan blade stage and the compressor blade stage;
a turbine in fluid communication with the combustor; and
an electrical machine located between the fan blade stage and the compressor blade stage, the electrical machine being configured
to generate electrical power, wherein the electrical machine includes a stator and a rotor; wherein the stator is positioned
adjacent to or within the shroud; and wherein the rotor extends from the wheel of the fan blade stage.

US Pat. No. 9,561,763

CONTROL SYSTEM FOR A DUAL REDUNDANT MOTOR/GENERATOR AND ENGINE

Rolls-Royce Corporation, ...

1. An electrical system comprising:
a power plant operable to provide motive power;
a first motor/generator and a second motor/generator operatively coupled to a common gear shaft, and to the power plant via
a gearbox and drive shaft; and

a system controller configured to selectively operate at least one of the first motor/generator and the second motor/generator
in one of a motor operating mode and a generator operating mode based on at least one parameter, and to alter the operating
mode when the at least one parameter exceeds a predetermined threshold;

wherein at least one of the first motor/generator and the second motor/generator provides power to the power plant in the
motor operating mode, and extracts power from the power plant in the generator operating mode;

wherein the at least one parameter includes a temperature of at least one of the first motor/generator and the second motor/generator.

US Pat. No. 9,481,473

DISTRIBUTED CONTROL SYSTEM WITH SMART ACTUATORS AND SENSORS

Rolls-Royce North America...

1. A system for controlling operation of an aircraft engine, comprising:
a master controller including a Full Authority Digital Engine Control (FADEC);
a control node device coupled to a solenoid of the aircraft engine and to the master controller, the solenoid operable based
on a frequency response curve and between a bounded lower curve and a bounded upper curve, wherein the control node device:

tracks performance of the solenoid during operation of the aircraft;
compares the tracked performance with the bounded lower and bounded upper curves;
determines whether the solenoid of the aircraft engine is performing within specification based on the comparison of the tracked
performance against the bounded lower and bounded upper curves; and

if not performing within specification, indicates that the solenoid is out of specification; and
wherein the FADEC controls thrust in response to the indication from the control node device.

US Pat. No. 9,157,328

COOLED GAS TURBINE ENGINE COMPONENT

Rolls-Royce North America...

16. An apparatus comprising:
a coolable gas turbine engine component having a hot side and a film cooling opening operable to discharge a cooling fluid
to create a film cooling for the hot side; and

a depression formed in part by an edge break of a cooling hole and having an upstream portion that descends from the hot side
to a valley, a downstream portion that ascends from the valley toward the hot side, the film cooling opening having a portion
that discharges through the upstream portion and a portion that discharges through the downstream portion of the depression,
wherein the edge break is defined as a surface formed in the coolable gas turbine engine component between the hot side and
a passage that provides the film cooling opening, wherein the edge break is part of a monolithic construction that includes
a portion of the coolable gas turbine engine component defining the cooling hole;

wherein the valley extends in a line from one lateral side of the film cooling opening to another side of the film cooling
opening.

US Pat. No. 9,121,351

GAS TURBINE ENGINE ACCESSORY SYSTEM

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine having a shaft operable to transmit mechanical power;
a first flow path disposed internal to the gas turbine engine and defined by an inner first flow path wall and an outer first
flow path wall;

a second flow path disposed radially inward of the first flow path in the gas turbine engine and defined by an inner second
flow path wall and an outer second flow path wall;

a gearing operably coupled to the shaft and disposed between the first flow path and the second flow path wherein the gearing
is operable to integrate with an accessory system; and

a multi-pad accessory mount disposed between the inner first flow path wall of the first flow path and the outer second flow
path wall of the second flow path, the multi-pad accessory mount coupled to the gearing;

wherein the multi-pad accessory mount is located in a fan frame of the gas turbine engine.

US Pat. No. 9,593,591

ENGINE HEALTH MONITORING AND POWER ALLOCATION CONTROL FOR A TURBINE ENGINE USING ELECTRIC GENERATORS

Rolls-Royce Corporation, ...

1. A method for controlling an allocation of power extracted from a plurality of shafts of a turbine engine, the shafts having
one or more electrical machines coupled thereto, the method comprising, with a controller, during operation of the turbine
engine:
monitoring a plurality of operating parameters of the turbine engine over time;
assessing a health of the turbine engine over time based on the plurality of operating parameters as they are monitored, wherein
the health of the turbine engine is indicative of a wear characteristic of the turbine engine; and

varying the allocation of power extraction between the plurality of electrical machines over time in response to the assessing
of the health of the turbine engine by assigning a part of a total power extraction to the one or more electrical machines
coupled to each of the plurality of shafts.

US Pat. No. 9,162,162

LIQUID FLOW WITH GAS MIXING

Rolls-Royce North America...

1. A fluid de-oxygenation system comprising:
a pressurized gas lumen containing pressurized de-oxygenating gas; and
a liquid transport lumen passing through said pressurize gas lumen, said liquid transport lumen including a plurality of openings
formed along its length within said pressurized gas lumen, a shearing feature downstream of said plurality of openings, and
a continued flow section downstream of said shearing feature;

wherein said de-oxygenating gas enters said liquid transport lumen through said plurality of openings to form a plurality
of bubbles within a transport liquid, said plurality of bubbles removing oxygen from said transport liquid;

wherein said shearing feature is configured to collapse said plurality of bubbles into larger bubbles prior to flow into said
continued flow section.

US Pat. No. 9,310,079

COMBUSTION LINER WITH OPEN CELL FOAM AND ACOUSTIC DAMPING LAYERS

Rolls-Royce North America...

1. A combustion liner, comprising:
an outer combustion liner wall;
an inner combustion liner wall;
an open cell foam disposed between the outer combustion liner wall and the inner combustion liner wall, wherein at least one
of the outer combustion liner wall and the inner combustion liner wall includes a plurality of openings extending therethrough;
and

a plurality of cells exposed to the plurality of openings;
wherein the open cell foam is formed of a composite material, wherein the composite material of the open cell foam is a ceramic
matrix composite.

US Pat. No. 9,586,690

HYBRID TURBO ELECTRIC AERO-PROPULSION SYSTEM CONTROL

Rolls-Royce Corporation, ...

8. A hybrid turbo electric aero-propulsion system comprising:
a turbine engine;
a generator to receive rotational energy from the turbine engine;
an energy storage subsystem to receive electrical input from the generator, the energy storage subsystem comprising one or
more of: an energy storage device, a power converter, and an inverter;

a motor to receive electrical input from the energy storage subsystem;
a fan that may be driven by either the engine or the motor; and
a control to optimize the efficiency of the turbo-electric aero-propulsion system by computing optimal set point values for
each of the turbine engine, the generator, the energy storage subsystem, the motor, and the fan to electronically control
the turbine engine, the generator, the energy storage subsystem, the motor and the fan.

US Pat. No. 9,410,482

GAS TURBINE ENGINE HEAT EXCHANGER

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine having a particle separator for cleaning an airflow and a compressor capable of increasing a pressure
of a working fluid in the gas turbine engine, the particle separator having a clean flow path and a dirty flow path;

a first passageway for the conveyance of a first air flow extracted from a first source of the compressor;
a second passageway for the conveyance of a second air flow extracted from a second source of the compressor, the second source
downstream of the first source; and

a heat exchanger separately maintaining the first air flow and second air flow and constructed such that the first air flow
cools the second air flow;

a passage that conveys a merged flow of a flow of dirty air from the dirty flow path of the particle separator and the first
air flow extracted from the compressor, the passage structured to receive the merged flow from the first passageway downstream
of the heat exchanger and the dirty flow path downstream of the particle separator; and

a blower positioned at a downstream portion of the passage and structured to receive and convey onward the first air flow
from the first source of the compressor and the flow of dirty air from the particle separator;

wherein the first air flow is extracted from the first source of the compressor at a first extraction location and the first
air flow is merged with the flow of dirty air at a merger location, the merger location being downstream of the first extraction
location;

wherein the gas turbine engine has an axial in-flow of an inlet air stream divided by the particle separator into the clean
flow path and the dirty flow path.

US Pat. No. 9,360,219

SUPERCRITICAL OR MIXED PHASE MULTI-PORT FUEL INJECTOR

Rolls-Royce North America...

1. A system, comprising:
a turbine engine having a fueling system, the fueling system comprising:
a fuel injector;
a plurality of fueling passages within the fuel injector;
a body;
a valve in the body, the valve fluidly coupled to a fuel supply of a fuel on an upstream side and fluidly coupled to the fueling
passages on a downstream side, wherein the valve is configured to maintain the fuel on the upstream side at a pressure at
or greater than a critical pressure of the fuel; and

an expansion chamber within the body interposed between the valve and the fueling passages,
wherein the expansion chamber is configured to receive the fuel that passes through the valve,
wherein each of the fueling passages has a corresponding inlet within the body, the inlets of the fueling passages are distributed
over a surface of a disc, and each of the inlets is configured to receive a portion of the fuel from the expansion chamber,
and

wherein each of the fueling passages is configured to flow the portion of the fuel received from the expansion chamber monotonically
downward.

US Pat. No. 9,303,523

SENSOR COMMUNICATION SYSTEM AND MACHINE HAVING THE SAME

Rolls-Royce North America...

1. A turbine engine comprising:
a first fluid passageway having an inlet and an outlet;
at least one combustion chamber positioned along said first fluid passageway between said inlet and said outlet, wherein a
primary fluid stream passes through said first fluid passageway and said at least one combustion chamber for generating power;

a second fluid passageway at least partially distinct from said first fluid passageway, wherein a secondary fluid stream passes
through said second fluid passageway to support the generation of power, and the second fluid passageway is further defined
as being one of a lubricant passageway and a fuel passageway;

a sensor assembly having a sensor operable to sense at least one condition and a transmitter associated with said sensor and
operable to emit a signal corresponding to the at least one condition wirelessly, wherein at least part of said transmitter
is positioned in said second fluid passageway to transmit the signal through said second fluid passageway; and

a receiver operable to receive the signal and positioned in the second fluid passageway, wherein the sensor assembly and receiver
are separate components separated by a portion of the second fluid passageway that includes at least one turn;

wherein said sensor assembly is further defined as being operable to emit the signal at a substantially optimized frequency
relative to a cross-section of said second fluid passageway such that said second fluid passageway functions as a waveguide.

US Pat. No. 9,156,549

AIRCRAFT VERTICAL LIFT DEVICE

ROLLS-ROYCE NORTH AMERICA...

1. An apparatus comprising: an aircraft having an engine operable to provide power; a lift fan powered by the engine and operable
to provide a lift for the aircraft along a vertical axis by producing a lift fan flow stream, the lift fan having a bladed
device structured to rotate about a rotational axis, the rotational axis fixed relative to the aircraft such that the rotational
axis is immobile relative to an aircraft axis during the entire time of operation of the lift fan when the bladed device is
rotated to produce the lift fan flow stream; a vane box fixed relative to the aircraft such that the vane box is substantially
aligned with a reference point on the aircraft and is incapable of moving relative to the aircraft during the entire time
of operation of the lift fan when the bladed device is rotated to produce the lift fan flow stream, the vane box positioned
downstream of the lift fan and housing a vane assembly structured to direct the lift fan flow stream in a downward direction,
the vane assembly having a plurality of vanes operable to direct the fan flow stream to a vertical vector for hovering flight,
a maximum aft vector for forward flight and a maximum forward vector for rearward flight; and wherein the vane box has a maximum
flow area when the plurality of vanes in the vane assembly are at a position other than the vertical vector, wherein the rotational
axis of the lift fan is oriented at an angle relative to the vertical axis.

US Pat. No. 9,151,166

COMPOSITE GAS TURBINE ENGINE COMPONENT

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine airflow member having an airfoil portion and a shroud portion and constructed from a composite material
construction including a matrix and a plurality of fiber plies, at least one of the fiber plies extending from the airfoil
portion to the shroud portion;

wherein the gas turbine engine airflow member includes a sealing knife coupled to the end of the gas turbine engine airflow
member,

wherein the sealing knife includes a plurality of extension members oriented along and extending transverse to an elongate
portion of the airfoil portion,

wherein the extension members include adjacent upright members that are stitched together;
wherein the transverse extension members include a plurality of fiber plies and the shroud portion includes a plurality of
fiber plies, and

a plurality of stitches affixing either a subset of all of the individual plies that form the transverse extension members
of the sealing knife to said at least one of the individual fiber plies extending to the shroud portion wherein a second subset
of all the individual plies that form the transverse extension members is not affixed to the shroud portion,

or a subset of all of the individual plies that form the shroud portion to at least one of the individual fiber plies forming
the transverse extension members wherein a second subset of all the individual plies that form the shroud portion is not affixed
to the transverse extension members.

US Pat. No. 9,573,853

MELT INFILTRATION APPARATUS AND METHOD FOR MOLTEN METAL CONTROL

Rolls-Royce North America...

1. An infiltration apparatus comprising:
an infiltrant having a first melting point,
an infiltrant source adapted to receive the infiltrant;
a solid barrier having a second melting point, the solid barrier being selected from the group of materials consisting of
Si/Zr, Si, SiC-coated silicon wafer, Zr/Si, ZrB2 and Ti;

a component comprising a ceramic matrix composite; and
a wick in fluid communication with the infiltrant source and the component, the wick being configured to draw the infiltrant
from the infiltrant source into the component;

wherein the second melting point is higher than the first melting point; and
wherein the solid barrier is disposed between the infiltrant source and the component and coupled to the wick to block fluid
communication through the wick until the infiltrant melts the barrier to allow the wick to draw the infiltrant from the infiltrant
source into the component.

US Pat. No. 9,784,115

BLADE TRACK ASSEMBLY, COMPONENTS, AND METHODS

Rolls-Royce North America...

10. An apparatus comprising:
a blade track assembly including a blade track having an upstream end and a downstream end,
a forward hanger and an aft hanger, each of the forward hanger and aft hanger having circumferentially extending channels
that receive the respective upstream and downstream ends of the blade track, at least one of the forward hanger and aft hanger
includes a hanger anti-movement member structured to discourage movement of the at least one of the forward hanger and aft
hanger relative to the blade track, and

a spring clip engaged with the blade track to discourage movement of the blade track relative to the forward hanger and the
aft hanger,

wherein the blade track includes a blade track anti-movement cutout that receives the hanger anti-movement member to discourage
movement of the blade track relative to the at least one of the forward hanger and the aft hanger and the spring clip is located
in the blade track anti-movement cutout and arranged around the hanger anti-movement member.

US Pat. No. 9,239,006

GAS TURBINE ENGINE AND SYSTEM FOR MODULATING SECONDARY AIR FLOW

Rolls-Royce North America...

15. The vane stage of claim 12, wherein at least one of the inner band and the outer band include openings in fluid communication with the main flow path
of the gas turbine engine and in fluid communication with the plenum.

US Pat. No. 9,309,778

VARIABLE VANE FOR GAS TURBINE ENGINE

Rolls-Royce North America...

1. A turbomachine, comprising:
a variably positioned vane;
a spindle integral with a vane outer button, wherein the vane is coupled to the vane outer button;
an annular sleeve defining the spindle, wherein the annular sleeve contacts the vane outer button at a radially inward extent
and contacts a turbine casing at a radially outward extent, wherein the annular sleeve comprises a wall portion positioned
perpendicular to the spindle, the wall portion including an aperture and the spindle extending from the vane outer button
to and through the aperture, and wherein the spindle includes threads;

a nut engaged with the threads, the nut applying force to the wall portion toward the radially inward extent;
wherein a radially outward end of the spindle extends though the turbine casing,
and wherein a cantilever rotation actuator is coupled to the radially outward end of the spindle;
a first rolling element engaging the annular sleeve substantially near the radially outward extent, wherein the first rolling
element is coupled to the turbine casing; and

a second rolling element engaging the annular sleeve substantially near the radially inward extent, wherein the second rolling
element is coupled to an outer endwall ring.

US Pat. No. 9,487,303

VEHICLE AND SYSTEM FOR SUPPLYING ELECTRICAL POWER TO A VEHICLE ELECTRICAL LOAD

Rolls-Royce North America...

1. A machine of a vehicle, comprising:
a fixed member;
a structure coupled to the fixed member;
an engine coupled to at least one of the structure and the fixed member;
an electrical load of the vehicle having a high energy device including a high power directed energy device that is associated
with the machine and utilizes above 270 volts during operations of the vehicle;

a generator coupled to the engine and configured to generate electrical power at a voltage above 270 volts for the electrical
load of the high energy device during operations of the vehicle;

first and second conductors electrically disposed between the electrical load and the generator;
first and second conduits arranged along each other and configured to house the respective first and second conductors; and
a dielectric gas disposed in at least one of the first and second conduits;
wherein the at least one of the first and second conduits is configured to envelop the conductor in the dielectric gas.

US Pat. No. 9,752,592

TURBINE SHROUD

Rolls-Royce Corporation, ...

8. The turbine shroud of claim 7, wherein the round carrier includes a connection flange adapted to be coupled to a turbine case, a connector arranged to
extend inwardly in the radial direction from the connection flange and having a frustoconical shape, and a support band arranged
to extend inwardly in the radial direction from the connector and connection flange, and the plurality of keys extend inwardly
in the radial direction from the support band.

US Pat. No. 9,284,887

GAS TURBINE ENGINE AND FRAME

Rolls-Royce North America...

1. A gas turbine engine frame, comprising:
a metallic inner hub;
a metallic outer construction;
a composite flowpath structure that is a separate component from the metallic inner hub and has a primary flowpath structure
disposed between said metallic inner hub and said metallic outer construction, wherein the composite flowpath structure includes
carbon bismaleimide composites, ceramic matrix composites, metal matrix composites, organic matrix composites and/or carbon-carbon
composites; and

a primary composite inner flowpath wall spaced radially apart from a primary composite outer flowpath wall defining the primary
flowpath structure for a working fluid of the gas turbine engine;

wherein said composite flowpath structure includes:
a plurality of inner composite struts wherein at least a portion of each inner composite strut extends between said primary
composite inner flowpath wall and said primary composite outer flowpath wall;

wherein said composite flowpath structure is formed as a single piece structure.

US Pat. No. 9,783,909

ANISOTROPIC ETCHING OF METALLIC SUBSTRATES

Rolls-Royce North America...

1. A method comprising:
forming a photoresist layer on a surface of a metallic substrate;
developing the photoresist layer to define a pattern exposing a portion of the surface of the metallic substrate;
forming an electrically conductive layer on a surface of the photoresist layer and the exposed portions of the surface of
the metallic substrate, wherein the electrically conductive layer contacts the exposed portions of the surface of the metallic
substrate; and

submerging the substrate, the photoresist layer, and the electrically conductive layer in an electrolyte solution; and
applying a voltage between a cathode and an anode submerged in the electrolyte solution to anisotropically etch the metallic
substrate where the electrically conductive layer contacts the exposed portions of the surface of the metallic substrate to
form at least one feature in the metallic substrate.

US Pat. No. 9,638,057

AUGMENTED COOLING SYSTEM

Rolls-Royce North America...

1. A cooling system comprising:
a component having an inner wall and an outer wall spaced apart from one another;
a plurality of pedestals extending between the inner and outer walls;
a plurality of inner trip strips projecting from the inner wall towards the outer wall at a predetermined height;
a plurality of outer trip strips projecting from the outer wall towards the inner wall at a predetermined height,
wherein one of either the plurality of inner trip strips or the plurality of outer trip strips extends between adjacent pedestals;
at least one inlet through aperture formed in the inner wall of the component operable for transporting a cooling fluid into
a space between the inner and outer walls of the component; and

a plurality of outlet through apertures formed in the outer wall of the component operable for transporting the cooling fluid
out of the space between the inner and the outer walls of the component;

wherein one or both of: the at least one inlet through aperture is located in an inner well bounded on all sides by a number
of the plurality of the inner trip strips, and, at least one of the plurality of outlet through apertures is located in an
outer well bounded on all sides by a number of the plurality of outer trip strips.

US Pat. No. 9,512,805

CONTINUOUS DETONATION COMBUSTION ENGINE AND SYSTEM

Rolls-Royce North America...

1. A combustion system configured for continuous detonation combustion, comprising:
an annular combustion chamber having a centerline axis, the annular combustion chamber configured to contain therein a rotating
continuous detonation wave; and

a first fluid diode positioned flow-wise upstream of the combustion chamber, including a first rotating diode structure and
a second rotating fluid diode structure; and wherein the first fluid diode is configured to form, with the first diode structure
and the second diode structure, during operation of the combustion system:

a first rotating region rotating at a speed that is the same as a speed of the rotating continuous detonation wave, wherein
the first rotating region has a first flow area through the first fluid diode; and wherein the first rotating region is positioned
adjacent to the rotating continuous detonation wave at a higher pressure zone of the rotating continuous detonation wave;
and

a second rotating region rotating about the combustion chamber at the same speed as the rotating continuous detonation wave,
wherein the second rotating region has a second flow area through the first fluid diode of a greater magnitude than a magnitude
of the first flow area; and wherein the second rotating region is positioned adjacent to a lower pressure zone circumferentially
of the rotating continuous detonation wave downstream of the higher pressure zone,

wherein the first rotating diode structure and the second rotating diode structure rotate around the centerline axis.

US Pat. No. 9,612,017

COMBUSTOR WITH TILED LINER

Rolls-Royce North America...

1. A combustor for use in a gas turbine engine, the combustor comprising
an outer case,
a combustion liner arranged radially inward of the outer case and arranged to define an annular combustion chamber, the combustion
liner including at least one monolithic annular liner tile, and

a mount assembly coupled to the outer case and to the combustion liner to locate the at least one monolithic annular liner
tile relative to the outer case, wherein the mount assembly extends from the outer case to the at least one monolithic annular
liner tile to locate the combustion liner relative to the outer case and wherein the mount assembly includes at least three
mount pins circumferentially spaced apart from one another and a tile hanger that extends from the at least one monolithic
annular liner tile to receive each of the at least three mount pins.

US Pat. No. 9,777,627

ENGINE AND COMBUSTION SYSTEM

Rolls-Royce North America...

1. An engine, comprising:
a combustion system, including:
a plurality of combustion channels, each combustion channel extending between a first open end and a second open end thereof;
a first end structure disposed adjacent to the first open ends of the combustion channels, wherein the first end structure
includes an inlet port configured to permit the first open ends to receive a fuel and an oxidant into the combustion channels;

an ignition source operative to ignite the fuel and the oxidant to form combustion products;
a second end structure disposed adjacent to the second open ends of the combustion channels, wherein the second end structure
includes an exhaust port configured to discharge the combustion products from the combustion channels; wherein the second
end structure is adapted to receive the ignition source; wherein the second end structure includes a cavity formed therein
and disposed between the exhaust port and the ignition source, the second end structure being further formed to include a
pair of turbulators that flank the cavity and the cavity is arranged to receive residual combustion products only from the
second open ends of the combustion channels, the second end structure further including an open face that opens into the cavity,
and the open face has a width that simultaneously exposes at least three combustion channels to the cavity at an interface
with the plurality of combustion channels, wherein the cavity is structured to enhance a fuel/oxidant mixture at the second
open ends; and

a conduit extending between a first opening and a second opening thereof, the first opening of the conduit being in fluid
communication with at least one second open end of the combustion channels via the cavity in the second end structure, the
second opening being in fluid communication with at least one first open end of the combustion channels via the first end
structure, and the conduit being operative to transmit the residual combustion products from the at least one second open
end to the at least one first open end of the combustion channels, and

wherein the combustion channels and the first and second end structures are configured such that operation of the combustion
system includes relative motion between the combustion channels and the first and second end structures; and wherein the combustion
channels, the first and second end structures, the exhaust port, the cavity, and the ignition source are configured relative
to one another, such that the relative motion includes relative movement between one of the second open ends of a corresponding
one of the combustion channels and the second end structure, whereby the second open end is exposed to the exhaust port, then
the cavity, and then the ignition source.

US Pat. No. 9,085,988

GAS TURBINE ENGINE FLOW PATH MEMBER

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine having a fluid cooled airfoil member disposed in a flow path and having a plurality of walls extending
along a span of the member and enclosing an open interior,

the walls forming a cooling passage therebetween,
an inner wall of the plurality of walls extending into the flow path beyond an end of an outer wall of the plurality of walls,
a base extending from the inner wall, the base being located at the end of the outer wall of the fluid cooled airfoil member
and serving as a base portion of a squealer and serving to enclose the open interior of the fluid cooled airfoil member; and

a plurality of apertures in the airfoil member extending from the enclosed open interior and having an upstream inlet and
a downstream exit and operable to pass a fluid therethrough oriented to cool the inner wall that extends beyond the end of
the outer wall,

wherein the downstream exits of the apertures are non-circular, and
wherein the plurality of apertures extending from the enclosed open interior are bounded by the inner wall and the outer wall,
and the plurality of apertures extend around an entire perimeter of the inner wall, the entire perimeter extending from a
trailing edge around a pressure side to a leading edge, and from the leading edge around a suction side returning to the trailing
edge.

US Pat. No. 9,657,969

MULTI-EVAPORATOR TRANS-CRITICAL COOLING SYSTEMS

Rolls-Royce Corporation, ...

1. A multi-evaporator cooling system, comprising:
a compressor circuit that generates multiple levels of evaporating pressures, the circuit comprising at least one compressor
configured to compress a refrigerant to a first pressure;

a heat exchanger configured to receive the compressed refrigerant from the at least one compressor and cool the compressed
refrigerant;

a first evaporator circuit configured to receive the compressed refrigerant from the heat exchanger, expand the compressed
refrigerant to a second pressure that is lower than the first pressure, and return the refrigerant to the compressor circuit;
and

a second evaporator circuit configured to receive the compressed refrigerant from the heat exchanger, expand the compressed
refrigerant to a third pressure that is lower than the second pressure, and return the refrigerant to the compressor circuit;

wherein the first and second evaporator circuits each comprise a respective expansion device and heat load heat exchanger,
wherein refrigerant from each heat load heat exchanger passes to a respective first and second ejector at approximately the
respective second and third pressures, expands in each of the first and second ejectors, and passes to an inlet of a compressor
of the at least one compressor circuit.

US Pat. No. 9,601,970

GAS TURBINE ENGINE AND ELECTRICAL SYSTEM

Rolls-Royce North America...

8. A gas turbine engine, comprising:
a high pressure spool;
a low pressure spool;
a controller; and
a first electrical machine coupled to the low pressure spool and coupled to a first inverter-converter controller and a third
inverter-converter controller, with the first inverter-converter controller being coupled to, and dedicated to, a first electrical
bus and the third inverter-converter controller being coupled to, and dedicated to, a second electrical bus;

a second electrical machine coupled to the high pressure spool and coupled to a second inverter-converter controller and a
fourth inverter-converter controller, with the second inverter-converter controller being coupled to, and dedicated to, the
first electrical bus and the fourth inverter-converter controller being coupled to, and dedicated to, the second electrical
bus;

an energy storage system configured to supply power to and absorb power from the first electrical bus and the second electrical
bus; and

a converter controller coupled to the energy storage system, the controller, the first electrical bus, and the second electrical
bus, wherein the converter controller is configured to control an amount of electrical power supplied to the first electrical
bus and the second electrical bus from the energy storage system under direction from the controller, and to control an amount
of electrical power received from the first electrical bus and from the second electrical bus and supplied to the energy storage
system under direction from the controller;

wherein the first inverter-converter controller and the third inverter-converter controller are configured to provide electrical
power in parallel to the first electrical bus under direction from the controller; and

wherein the second inverter-converter controller and the fourth inverter-converter controller are configured to provide electrical
power in parallel to the second electrical bus under direction from the controller.

US Pat. No. 9,731,836

PROPULSION, ELECTRICAL, AND THERMAL MANAGEMENT DEVICE FOR A SMALL UNMANNED AERIAL VEHICLE

Rolls-Royce North America...

1. A system for providing propulsion, electrical generation and thermal management to an airframe from a single prime mover,
comprising:
an engine providing power to an engine output shaft;
a speed change transmission, said transmission having an input shaft and at least two output shafts, said input shaft operatively
connected to said engine output shaft to distribute power from said engine to a first transmission output shaft and a second
transmission output shaft, said first transmission output shaft operatively coupled to an electric machine and to a propeller
such that both said electric machine and said propeller are driven by said first transmission output shaft, and said second
transmission output shaft operatively coupled to a refrigerant compressor of a refrigeration system, wherein said refrigerant
compressor is coupled with said propeller via said first transmission output shaft and via said second transmission output
shaft, and

a condenser for said refrigeration system mounted on a structure that surrounds said propeller.

US Pat. No. 9,640,959

PLATFORM WITH ENGINE AND WIRING HARNESS SYSTEM, PLATFORM WITH CONTROLLED SYSTEM AND WIRING HARNESS SYSTEM, AND WIRING HARNESS SYSTEM

Rolls-Royce North America...

1. A platform, comprising:
a structure;
a platform electrical system;
an engine coupled to the structure, wherein the engine includes an engine control system; and
a wiring harness system coupled between and connecting the engine control system to the platform electrical system, wherein
the wiring harness system includes an engine wiring harness;

the engine wiring harness coupled to a platform wiring harness; an interface configured to electrically connect the engine
wiring harness to the platform wiring harness; and at least one electromagnetic interference (EMI) protection component coupled
to, and surrounding a pin within a connector head of the engine wiring harness and configured to provide EMI protection for
the engine control system,

wherein the connector head comprises a plurality of pins and at least one of the plurality of pins is not electrically shielded,
and

wherein each of the at least one EMI protection components surrounds a respective individual pin within the connector head.

US Pat. No. 9,759,079

SPLIT LINE FLOW PATH SEALS

Rolls-Royce Corporation, ...

1. A gas turbine engine assembly, the assembly comprising
a first component comprising ceramic matrix materials, the first component including a panel arranged to separate a high pressure
zone from a low pressure zone and formed to include a first chamfer surface that extends from a high pressure surface of the
first component facing the high pressure zone to a first side surface of the first component,

a second component comprising ceramic matrix materials, the second component including a panel arranged to separate the high
pressure zone from the low pressure zone and formed to include a second chamfer surface that extends from a high pressure
surface of the second component facing the high pressure zone to a second side surface of the first component,

a seal assembly arranged in a channel formed by the first chamfer and the second chamfer when the first side surface of the
first component is arranged in confronting relation to the second side surface of the second component, the seal assembly
including a rod configured to block gasses from passing through the interface of the first side surface included in the first
component with the second side surface included in the second component and a rod locator configured to engage the rod to
hold the rod in place relative to the first component and the second component,

wherein the seal assembly includes a bias member configured to push the rod into contact with the first chamfered surface
of the first component and the second chamfered surface of the second component, and

wherein the rod locator and the bias member are included in a singular component.

US Pat. No. 9,624,786

BRAZE MATERIALS AND METHOD FOR JOINING OF CERAMIC MATRIX COMPOSITES

Rolls-Royce Corporation, ...

1. A melt alloy for use as a braze material for joining ceramic matrix composites, the melt alloy comprising:
a homogeneous mixture of two or more materials in powder form, one of the two or more materials being a braze alloy comprising
silicon and another of the two or more materials being a high melting point material or alloy,

wherein the homogeneous mixture of the two or more materials in powder form is selected from the group consisting of (in wt.
%): Ti-9Si (5%)+71Si—Cr (95%), 75Si—Ti (97%)+C (3%), 75Si—Ti (97%)+B (3%), 75Si—Ti (50%)+71Si—Cr (50%), 75Si—Ti (50%)+Ti-43Cr
(50%), 75Si—Ti (50%)+12.5Si—Co (50%), and Ti-25Cr-23Si (97%)+C (3%).

US Pat. No. 9,771,867

GAS TURBINE ENGINE WITH AIR/FUEL HEAT EXCHANGER

Rolls-Royce Corporation, ...

1. An aircraft propulsion gas turbine engine, comprising:
a first compressor stage configured to produce a pressurized air flow;
a second compressor stage disposed downstream of the first compressor stage;
a primary annular flowpath fluidly coupling the first compressor stage and the second compressor stage, wherein the primary
annular flowpath is disposed within the aircraft propulsion gas turbine engine;

a combustor disposed downstream of the second compressor stage;
a fuel injector configured to inject a fuel into the combustor, wherein the combustor is configured to combust the fuel injected
therein by the fuel injector; and

an air/fuel heat exchanger disposed in the primary annular flowpath, the air/fuel heat exchanger including a first flowpath
and a second flowpath, the first flowpath having a first fuel inlet and a first fuel outlet for distributing fuel in a first
direction through a first side of the air/fuel heat exchanger, and the second flowpath having a second fuel inlet and a second
fuel outlet for distributing fuel in a second direction through a second side of the air/fuel heat exchanger;

wherein the air/fuel heat exchanger is an annular heat exchanger with the first direction being a first circumferential direction
and the second direction being a second circumferential direction;

wherein the air/fuel heat exchanger is in fluid communication with the fuel injector, the first compressor stage and the second
compressor stage; and

wherein the air/fuel heat exchanger is configured to receive the pressurized air flow from the first compressor stage, to
discharge the pressurized air flow to the second compressor stage, to heat the fuel by heat exchange with the pressurized
air flow prior to delivery of the fuel to the fuel injector, and to cool the pressurized air flow by heat exchange with the
fuel.

US Pat. No. 9,745,856

PLATFORM FOR CERAMIC MATRIX COMPOSITE TURBINE BLADES

Rolls-Royce Corporation, ...

1. An apparatus comprising:
a disk having at least a first airfoil slot and a second airfoil slot;
a first ceramic matrix composite airfoil coupled to the disk and mounted within the first slot;
a second ceramic matrix composite airfoil coupled to the disk and mounted within the second slot;
a platform disposed between the first airfoil and the second airfoil;
the platform comprising a first axial extension and a second axial extension, and further comprising a first radial extension
and a second radial extension;

the first axial extension extending adjacent to the first airfoil and extending axially fore and aft of the first airfoil;
the second axial extension extending adjacent to the second airfoil and extending axially fore and aft of the second airfoil;
the first radial extension extending radially into and conforming to the first slot;
the second radial extension extending radially into and conforming to the second slot; and
further comprising a first groove disposed between the first axial extension and the first radial extension, wherein the first
groove houses a first sealing dynamic damper.

US Pat. No. 9,771,870

SEALING FEATURES FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A blade for a gas turbine engine comprising
a body including a root configured to engage a turbine rotor, an airfoil extending from the root, and a receiver positioned
distally from the root, and

a floating blade seal received in the receiver such that centrifugal force applied to the floating blade seal during rotation
of the blade about an axis of rotation causes the floating blade seal to move relative to the body and seat against the receiver
to extend a radial height of the blade,

wherein the receiver defines a channel having a first width at a base of the receiver and a second width at an apex of the
receiver, the base positioned closer to the root than the apex, the first width being greater than the second width, and wherein
the receiver has first and second surfaces that extend between the base and the apex, the first and second surfaces converging
from the base to the apex.

US Pat. No. 9,708,914

GAS TURBINE ENGINE AIRFLOW MEMBER HAVING SPHERICAL END

Rolls-Royce Corporation, ...

1. A gas turbine engine comprising
a variable position airfoil member that includes a hub end having a leading edge and a trailing edge, the variable position
airfoil member disposed within a flow path annulus adjacent a hub flow path surface having a divot that includes a downstream
end in which a radial rate of growth of the downstream end is increasing, the variable position airfoil member pivotable around
an adjustment axis and capable of being moved between a first position and a second position where the first position provides
a flow area greater than a flow area in the second position, the hub end offset from the divot to permit pivoting movement
of the variable position airfoil member without interference from the flow path surface between the first position and the
second position, wherein the offset between the flow path surface of the divot and the hub end of the variable position airfoil
member is constant as the variable position airfoil member moves between the first position and the second position, wherein
the hub end is spherical, the variable position airflow member includes a tip end that is spherical, and the axial extent
of spherical shape at the tip end forward of a pivot axis of the variable position airflow member is shifted backward relative
to the axial extent of the spherical shape at the hub end forward of the pivot axis.

US Pat. No. 9,624,870

ADAPTIVE FAN SYSTEM FOR A VARIABLE CYCLE TURBOFAN ENGINE

Rolls-Royce North America...

1. A gas turbine engine, comprising;
a compressor structured to compress an airflow received at the compressor and to output the compressed airflow as a compressor
discharge airflow;

a combustor in fluid communication with said compressor, said combustor being structured to combust a mixture of a fuel and
at least some of said compressor discharge airflow to generate a hot working airflow;

a turbine in fluid communication with said combustor, said turbine being configured to extract a mechanical power from said
hot working airflow;

a shaft coupled to said turbine, said shaft being configured to receive and transmit said mechanical power from said turbine;
a first rotating load powered by said shaft;
a second rotating load powered by said shaft, wherein the first rotating load and the second rotating load rotate in the same
direction; and

a transmission system including:
a geartrain coupled to said shaft;
a clutching mechanism coupled to said shaft in parallel to said geartrain,
wherein said clutching mechanism includes a clutch and a positive locking mechanism disposed axially aft in relation to the
clutch, wherein the second rotating load is disposed forward of the clutch, wherein the clutch couples the shaft to the second
rotating load through alternating engagement of one of the positive locking mechanism and the geartrain such that when the
positive locking mechanism is engaged the second rotating load is driven by the shaft at a same speed as the first rotating
load and when the geartrain is engaged the second rotating load is driven at a speed less than the speed of the first rotating
load, said transmission system being structured to selectively vary a speed at which power is supplied from said shaft to
said second rotating load relative to a speed of at least one of said shaft and said first rotating load, wherein to increase
extraction of mechanical power the second rotating load rotates at a same speed as the first rotating load.

US Pat. No. 9,587,517

BLADE TRACK ASSEMBLY WITH TURBINE TIP CLEARANCE CONTROL

Rolls-Royce North America...

1. A turbine shroud assembly for a gas turbine engine, the turbine shroud assembly comprising
a carrier configured to change size in response to a change in temperature, the carrier arranged around an engine axis to
form a ring, the carrier formed to include a plurality of guide slots extending perpendicular to the engine axis,

a blade track concentric with the carrier made of a plurality of segments, the blade track movable between a radially-inward
position having a first inner diameter and a radially-outward position having a second inner diameter larger than the first
inner diameter, and

a plurality of guide pins positioned in the guide slots for movement therein relative to the carrier, attached to the segments
of the blade track for movement with the blade track relative to the carrier, and configured to couple the carrier to the
blade track, wherein the guide pins are movable within the guide slots so that the segments of the blade track each move substantially
inwardly and outwardly in a radial direction relative to the engine axis as the blade track moves between the radially-inward
position and the radially-outward position in response to movement of the guide pins within the guide slots.

US Pat. No. 9,587,507

BLADE CLEARANCE CONTROL FOR GAS TURBINE ENGINE

Rolls-Royce North America...

8. A method comprising:
spinning a gas turbine engine bladed component within a flow path surface about a first axis to change a pressure of a working
fluid;

moving a pivoting member about the first axis in a circumferential direction;
interacting a cam and a cam follower as the pivoting member moves circumferentially; and
axially adjusting the flow path surface to adjust a clearance between the flow path surface and the gas turbine engine bladed
component as a result of the interacting.

US Pat. No. 9,719,357

TRENCHED COOLING HOLE ARRANGEMENT FOR A CERAMIC MATRIX COMPOSITE VANE

Rolls-Royce Corporation, ...

1. An apparatus comprising:
a shape operable as a gas turbine engine component;
an internal cavity within the shape including a radius;
a trench on an external surface of the shape including a rear face tangential to an arc centered on the radius of the internal
cavity; and

a cooling hole extending from the internal cavity and exiting to the trench through the rear face of the trench,
wherein a cooling fluid introduced to the internal cavity flows through the cooling hole and into the trench during operation
of the gas turbine engine component, the rear face extends to a point spaced a first distance from a suction sidewall defined
by the external surface of the shape, the cooling hole is spaced a second distance from the suction sidewall greater than
the first distance, and the rear face of the trench is centered on a centerline of the radius of the internal cavity.

US Pat. No. 9,617,868

GAS TURBINE ENGINE VARIABLE GEOMETRY FLOW COMPONENT

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine component having a plurality of fixed airfoil members distributed circumferentially about a reference
axis;

at least one movable airfoil extension structured for use in a flow path of a gas turbine engine and positioned intermediate
a first one of the plurality of fixed airfoil members and a second one of the plurality of fixed airfoil members, the at least
one movable airfoil extension configured to be moved between a first operating position adjacent the first one of the plurality
of fixed airfoil members and a second operating position adjacent the second fixed airfoil member, wherein at transition positions
intermediate the first operating position and the second operating position the at least one movable airfoil extension is
free of interfacing contact with the first and second one of the plurality of airfoil members; and

wherein movement of the at least one movable airfoil extension between the first operating position and the second operating
position alters a flow path characteristic of both the first one of the plurality of fixed airfoil members and the second
one of the plurality of fixed airfoil members;

wherein the flow path characteristic is at least one of inlet incidence, exit swirl, effective throat area, and camber;
wherein the at least one movable airfoil extension is a leading edge addition, which further includes another airfoil extension
located at a trailing edge, and wherein the plurality of fixed airfoil members includes cooling apertures that are exposed
in the first operating position and covered in the second operating position.

US Pat. No. 9,587,561

HEAT EXCHANGER INTEGRATED WITH A GAS TURBINE ENGINE AND ADAPTIVE FLOW CONTROL

Rolls-Royce North America...

1. A gas turbine engine, comprising:
an engine including an engine centerline and a passage being one of a fan bypass duct, a third stream duct and a ram air duct;
at least one contiguous heat exchanger having a cold-side inlet surface receiving a cold-side airflow; and
at least one of:
a ring slidably coupled to one of an inner diameter and an outer diameter of the passage, or
a forward-facing scoop disposed within the passage;
wherein the at least one contiguous heat exchanger is disposed within the passage such that a surface normal relative to the
cold-side inlet surface is offset by at least 30 degrees from the engine centerline.

US Pat. No. 9,845,688

COMPOSITE BLADE WITH AN INTEGRAL BLADE TIP SHROUD AND METHOD OF FORMING THE SAME

Rolls-Royce Corporation, ...

1. A gas turbine engine airflow member, comprising:
a blade core portion;
a shroud tip portion extending from the blade core portion;
an airfoil portion formed exteriorly to and surrounding the blade core portion and terminating at the shroud tip portion forming
a shroud interface at a lateral portion of the shroud tip portion; and

wherein the blade core portion and the shroud tip portion are constructed as a first unitary structure and the airfoil portion
is constructed as a second structure,

wherein the first unitary structure is formed of a single-piece monolithic material and the second structure is formed of
a composite material having a plurality of fiber plies with a two dimensional orientation,

wherein the single-piece monolithic material comprises a first composite material and wherein the second structure comprises
a second composite material, and

wherein the first composite material includes a first plurality of fiber plies having a three dimensional orientation, and
wherein the second composite material includes a second plurality of fiber plies having a two dimensional orientation,

further including a sealing knife portion extending from the shroud tip portion opposite the shroud interface.

US Pat. No. 9,822,817

HIGH SPEED BEARING ASSEMBLY

Rolls-Royce Corporation, ...

1. A bearing assembly comprising
an inner race that extends around a central axis,
an outer race that extends around the central axis radially outward of the inner race,
a plurality of internal rotating components arranged radially between the inner race and the outer race to engage the inner
race and the outer race, and

a cage that extends around the central axis radially between the inner race and the outer race, the cage including a cage
rail formed to include a plurality of apertures that receive the plurality of internal rotating components to locate the plurality
of internal rotating components relative to one another within the bearing assembly and a plurality of lubricant-ejector fins
that extend radially outward from the cage rail that are shaped to push lubricant between the cage rail and the outer race
out of the bearing assembly during rotation of the cage relative to the outer race in a direction of rotation so that hot
lubricant is removed from the bearing assembly to make room for cooler lubricant introduced into the bearing assembly during
rotation of the cage.

US Pat. No. 9,803,486

BI-CAST TURBINE VANE

Rolls-Royce North America...

1. A gas turbine engine vane, comprising:
an airfoil having an outer surface extending between a leading edge and a trailing edge and between a first end and a second
end;

a through slot extending between the first and second ends of the airfoil; and
a spar slidingly engaged with the slot of the airfoil, the spar including a pair of extensions with at least one bi-cast groove
formed on opposing ends thereof,

wherein the extensions of the spar are configured to engage with corresponding apertures formed in a pair of opposing endwalls.

US Pat. No. 9,683,488

GAS TURBINE ENGINE IMPELLER SYSTEM FOR AN INTERMEDIATE PRESSURE (IP) COMPRESSOR

Rolls-Royce North America...

1. A gas turbine engine comprising:
a compressor assembly rotationally coupled to a shaft, the compressor assembly having a centrifugal impeller;
a shroud covering a bladed portion of the centrifugal impeller, the shroud having a first flange extending therefrom;
a diffuser having a second flange extending therefrom that is parallel to and in direct contact with the first flange, the
diffuser attached to the shroud via the first flange and the second flange, the diffuser including a strut that is mounted
through an aft-extending leg to a base of an intercase; and

a sealing assembly attached to the diffuser, wherein the sealing assembly is attachable to a transition duct that is positioned
to receive air from the diffuser, and wherein the sealing assembly is configured to prevent air from passing through the sealing
assembly while allowing relative motion to occur between the transition duct and the diffuser.

US Pat. No. 9,845,700

ACTIVE SEAL SYSTEM

Rolls-Royce North America...

1. An active knife seal system, comprising:
a rotor having a rotating seal component and a first electrical generator element; and
a stationary support;
a stationary seal component coupled to the stationary support to move radially relative to the stationary support and disposed
adjacent to the rotor;

a second electrical generator element coupled to the stationary support in a fixed position relative to the stationary support;
and a piezoelectric portion in electrical communication with the second electrical generator element;
wherein the first electrical generator element and the second electrical generator element are configured to cooperate to
generate electrical power in the second electrical generator element when the rotor is rotated;

wherein the piezoelectric portion is configured to change in radial thickness to cause at least a part of the stationary seal
component to move radially in response to changes in an amount of the electrical power received by the piezoelectric portion;

wherein the amount of electrical power received by the piezoelectric portion increases as the first electrical generator element
moves radially toward the second electrical generator element; and

wherein the active knife seal system is configured such that radial growth of the rotor causes the first electrical generator
element to move toward the second electrical generator element,

wherein the stationary seal component includes an abradable surface arranged to face toward the rotating seal component and
engage the rotating seal component to minimize any gap formed between the stationary seal component and the rotating seal
component and

wherein the piezoelectric portion is a continuous ring and decreases in radial thickness in response to outward radial movement
of the rotating seal component toward the stationary seal component to minimize abrasion of the abradable surface while also
minimizing the gap.

US Pat. No. 9,638,056

GAS TURBINE ENGINE AND ACTIVE BALANCING SYSTEM

Rolls-Royce North America...

1. A gas turbine engine, comprising:
a compressor;
a combustor in fluid communication with the compressor;
a turbine in fluid communication with the combustor; and
an active balancing system configured to balance at least a portion of the gas turbine engine during operation of the gas
turbine engine, including:

a rotor component positioned in the turbine and coupled to a shaft of the gas turbine engine for rotation about an axis, the
rotor component having a disk and an arm extending axially from the disk, a first magnetic element coupled to the arm and
spaced apart from the shaft;

a static engine component having a second magnetic element positioned radially inward of the first magnetic element and spaced
apart from the shaft, wherein the second magnetic element and the first magnetic element are disposed to be positioned opposite
each other at one point during a rotation of the rotor component;

a vibration sensor configured to sense a vibration in the gas turbine engine; and
a controller configured to execute program instructions to activate one or both of the first magnetic element and the second
magnetic element to selectively attract one to the other or to selectively repel one from the other based on an output of
the vibration sensor to reduce the vibration.

US Pat. No. 10,047,624

TURBINE SHROUD SEGMENT WITH FLANGE-FACING PERIMETER SEAL

Rolls-Royce North America...

1. A turbine shroud segment comprisinga carrier segment comprising metallic materials, the carrier segment formed to define an attachment-receiving space,
a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend partway around a central axis and an attachment portion that extends radially outward from the runner into an attachment-receiving space channel formed by the carrier segment, and
a seal member configured to resist the movement of gasses into the attachment-receiving space, the seal member shaped to extend around the attachment portion of the blade track segment and arranged to engage a radially-outwardly facing surface of the runner, wherein the seal member is formed to include a plurality of radially-extending bleed holes adapted to conduct a flow of buffer air through the seal member to resist the movement of gasses into the attachment-receiving space.

US Pat. No. 10,018,059

ENGINE NACELLE

Rolls-Royce North America...

1. A nacelle for a jet engine comprisingan inner surface defining an opening for air to flow to an engine intake,
an outer surface positioned external to the inner surface, and
a leading surface circumscribing the opening, the leading surface connecting the inner surface and the outer surface, the leading surface defining a line of stagnation and formed to include a plurality of vortex generators positioned on leading surface along the line of stagnation,
wherein the vortex generators comprise a plurality of tabs extending from the leading surface, the tabs oriented to disrupt air flow flowing laterally across the leading surface,
wherein the tabs comprise a body and a plurality of leading edges that are generally perpendicular to the line of stagnation, and
wherein each of the tabs has converging sides that each terminate in the leading edges.

US Pat. No. 9,771,157

AIRCRAFT AND AIRBORNE ELECTRICAL POWER AND THERMAL MANAGEMENT SYSTEM

Rolls-Royce Corporation, ...

1. An electrical power and thermal management system, comprising:
a gas turbine engine including: a combustor and a turbine in fluid communication with the combustor;
a generator powered by the turbine and configured to provide electrical power to an electrical load;
a refrigerant compressor powered by the turbine;
a condenser in fluid communication with the refrigerant compressor; and
at least one evaporator in fluid communication with the condenser, wherein the at least one evaporator is configured to extract
heat from at least one heat source;

wherein the gas turbine engine, the generator, the refrigerant compressor, the at least one evaporator, and the condenser
are disposed in an external pod for a vehicle.

US Pat. No. 9,874,110

COOLED GAS TURBINE ENGINE COMPONENT

Rolls-Royce North America...

11. A method comprising:
free form fabricating a gas turbine engine component core having an inner surface and an outer surface representing a cooling
space of a cast gas turbine engine component, the fabricating including;

building a core portion representing an internal flow space of the gas turbine engine component;
forming a first cooling hole core fused with the core portion and having a first end and a second end and a bend intermediate
the first and second ends, the first end of the first cooling hole core coupled with the core portion at a leading edge of
the core portion;

forming a second cooling hole core fused with the core portion and having a first end and a second end, the first end of the
second cooling hole core coupled with the core portion at a trailing edge of the core portion on a pressure side of the core
portion; and

wherein the gas turbine engine component is a cooled turbine airflow member, and the fabricating further includes forming
a first internal passage core that extends from a midspan of the core portion toward the leading edge of the core portion
and forming a second internal passage core that extends from the midspan of the core portion on a suction side of the core
portion toward the trailing edge of the core portion, the first internal passage core is coupled with the second end of the
first cooling hole core, and the second internal passage core is coupled with the second end of the second cooling hole core.

US Pat. No. 9,856,824

AIRCRAFT NOZZLE SYSTEM

Rolls-Royce North America...

1. An apparatus comprising:
an aircraft having a combined cycle powerplant and the combined cycle powerplant includes:
a gas turbine engine operable to provide propulsive power in a first configuration and having a gas turbine engine centerline;
a ramjet operable to provide propulsive power in a second configuration and having a ramjet centerline offset from the gas
turbine engine centerline;

an engine partition structure located between the gas turbine engine and the ramjet and oriented to define part of a flowpath
of the gas turbine engine, the engine partition structure being fixed relative to the gas turbine engine centerline; and

a nozzle having a movable nozzle assembly, the movable nozzle assembly configured to translate axially relative to the gas
turbine engine centerline to vary an exhaust of the flowpath of the gas turbine engine, the movable nozzle assembly having
a convergent upstream portion forming a surface of the flowpath of the gas turbine engine in the first configuration, the
movable nozzle assembly also having a divergent portion downstream of the convergent upstream portion that permits an expansion
of an exhaust flow for the combined cycle powerplant,

wherein the convergent upstream portion of the movable nozzle assembly moves relative to the engine partition structure and
cooperates with the engine partition structure to open and close the exhaust of the flowpath of the gas turbine engine.

US Pat. No. 9,845,831

CLUTCH WITH REDUNDANT ENGAGEMENT SYSTEMS

Rolls-Royce North America...

1. A clutch comprising
a first shaft,
a second shaft,
a primary engagement system configured to selectively transmit rotation from the first shaft to the second shaft, the primary
engagement system including a first frustoconical engagement member coupled for common rotation with the first shaft and a
second frustoconical engagement member coupled for common rotation with the second shaft, and

a secondary engagement system configured to selectively transmit rotation from the first shaft to the second shaft, the secondary
engagement system including first shaft splines coupled to the first shaft for movement therewith and second shaft splines
coupled to the second shaft for movement therewith,

wherein the second frustoconical engagement member of the primary engagement system is coupled to the second shaft to slide
relative to the second shaft from a first position disengaged from the first frustoconical engagement member to a second position
engaged with the first frustoconical engagement member.

US Pat. No. 10,012,100

TURBINE SHROUD WITH TUBULAR RUNNER-LOCATING INSERTS

Rolls-Royce North America...

1. A turbine shroud comprisingan annular metallic carrier arranged around a central axis of the turbine shroud and formed to include a plurality of keyways extending in a radial direction into the annular metallic carrier, and
a blade track including a ceramic annular runner and a plurality of ceramic inserts extending outward in a radial direction away from the ceramic annular runner,
wherein each of the plurality of ceramic inserts are tubular and are arranged to extend into a corresponding one of the plurality of keyways formed in the annular metallic carrier to locate the blade track and the annular metallic carrier relative to the central axis while allowing radial growth of the annular metallic carrier and the blade track at different rates during use of the turbine shroud.

US Pat. No. 9,884,789

MELT INFILTRATION APPARATUS AND METHOD FOR MOLTEN METAL CONTROL

Rolls-Royce North America...

1. A method of infiltrating a material into a component, the method comprising:
providing an infiltrant source having an infiltrant contained therein;
providing a component in fluid communication with the infiltrant source;
heating the infiltrant source, the infiltrant, the component and a barrier disposed between the infiltrant source and the
component, the barrier having a higher melting point than the infiltrant;

dissolving the barrier; and
infusing the infiltrant into the component.

US Pat. No. 9,856,884

VALVE FOR GAS TURBINE ENGINE

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine having a compressor that includes a row of rotatable blades for compressing a working fluid between a
first flow surface and a second flow surface of the compressor;

a flow path operable to deliver an injection fluid radially inward to the compressor through an air flow port of the first
flow surface;

an air injection system for the compressor including a valve band extending across the air flow port and moveable in a circumferential
direction relative to the gas turbine engine to selectively permit the injection fluid to flow into the compressor;

wherein the valve band includes a first end that is fixed relative to the gas turbine engine,
wherein the valve band moves radially to open and close the air flow port, and
which further includes a biasing member to resist closing of the air flow port with the valve band.

US Pat. No. 9,797,270

RECESSABLE DAMPER FOR TURBINE

Rolls-Royce North America...

1. A turbine blade damper system comprising
a platform having a leading end, a trailing end, a first circumferential side, a second circumferential side, a radially-outward
side, and a radially-inward side, the radially-inward side defining a pocket extending circumferentially inwardly from the
first circumferential side, the pocket having a ceiling portion, a first wall portion extending radially inward from the ceiling
portion proximate the leading end, a second wall portion extending radially inward from the ceiling portion proximate the
trailing end, a first floor portion extending from the first wall portion toward the second wall portion, and a second floor
portion extending from the second wall portion toward the first wall portion, the ceiling portion defining a pocket beveled
surface extending circumferentially inwardly from the first circumferential side, and the second circumferential side defining
a platform sealing surface; and

a damper having a leading end, a trailing end, a first circumferential side, a second circumferential side, a body portion,
a radially-outward side, a first leg portion extending radially inwardly from the leading end of the body portion, a second
leg portion extending radially inwardly from the trailing end of the body portion, the first circumferential side defining
a damper sealing surface, the radially-outward side defining a damper beveled surface;

the damper beveled surface slidingly engaged with the pocket beveled surface, and the damper movable with respect to the platform
between a first position wherein at least a portion of the damper is recessed within the pocket and a second position wherein
a lesser portion of the damper is recessed within the pocket and the damper sealing surface is engaged with a platform sealing
surface of an adjacent platform, wherein the first and second leg portions of the damper define respectively a first outer
surface and a second outer surface, wherein the first outer surface and the second outer surface are in sliding engagement
with the first and second wall portions of the pocket, wherein the first and second leg portions of the damper define respectively
a first foot and a second foot, wherein the first foot and the second foot are respectively engaged with the first and second
floor portions when the damper is in the first position, and further comprising a first lip extending radially outwardly from
the first floor portion proximate a free end of the first foot and a second lip extending radially outwardly from the second
floor portion proximate a free end of the second foot.

US Pat. No. 10,072,511

ENGINE NACELLE

Rolls-Royce North America...

1. A nacelle for a jet engine comprisingan inner surface defining an opening for air to flow to an engine intake,
an outer surface positioned external to the inner surface, and
a leading surface circumscribing the opening, the leading surface connecting the inner surface and the outer surface, the leading surface defining a line of stagnation and formed to include a plurality of vortex generators positioned on leading surface along the line of stagnation,
wherein the vortex generators comprise a plurality of tabs extending from the leading surface, the tabs oriented to disrupt air flow flowing laterally across the leading surface,
wherein the tabs comprise a body and a plurality of leading edges that are generally perpendicular to the line of stagnation, and
wherein each of the tabs has converging sides that each terminate in the leading edges.

US Pat. No. 9,908,635

AIRCRAFT SYSTEM

Rolls-Royce North America...

1. An apparatus comprising:
an aircraft configured for supersonic flight;
an aircraft power plant structured to provide thrust to the aircraft to achieve supersonic flight, the aircraft power plant
including a compressor, a combustor configured to receive a compressor discharge airflow from the compressor, and a turbine
coupled to the compressor and configured to receive a stream of combustion products from the combustor and a compressed airflow
from the compressor, the aircraft power plant characterized by a thermodynamic cycle;

a ram air turbine that receives a working fluid and that rotates to produce power when the working fluid traverses therethrough,
the ram air turbine is structured to extract work from the working fluid and provide one of heat or work to the thermodynamic
cycle of the aircraft power plant, and the ram air turbine discharges cooled working fluid in response to the working fluid
traversing through the ram air turbine;

a power device structured to receive power from rotation of the ram air turbine, the power device including an electric generator;
and

an electric heat source powered by the power device, the electric heat source positioned to heat at least one of the compressed
airflow, the compressor discharge airflow, and the stream of combustion products during operation of the aircraft power plant,
and the electric heat source is one of an induction heater and a resistive heater,

wherein the apparatus further includes an air-to-air heat exchanger that is in fluid communication with the compressor to
receive the compressed airflow from the compressor, the air-to-air heat exchanger is in fluid communication with the turbine
to conduct the compressed airflow to the turbine, the air-to-air heat exchanger is further in fluid communication with the
ram air turbine such that the cooled working fluid discharged from the ram air turbine is conducted through the air-to-air
heat exchanger and used to cool the compressed airflow before the compressed airflow is provided to the turbine to cool the
turbine of the aircraft power plant.

US Pat. No. 9,938,198

METHOD FOR INTEGRAL JOINING INFILTRATED CERAMIC MATRIX COMPOSITES

Rolls-Royce Corporation, ...

1. A method of making an integrated ceramic matrix composite component for use in a gas turbine engine, the method comprising the steps ofmanufacturing a first green body subpart formed to include a first slot,
manufacturing a second green body subpart formed to include a second slot,
inserting a green body biscuit into the first slot of the first green body subpart and the second slot of the second green body subpart to create a green assembly with a joint between the first green body subpart and the second green body subpart, and
slurry infiltrating the green assembly with ceramic-containing matrix to integrally join the green assembly and produce an integrated ceramic matrix composite component.

US Pat. No. 9,915,149

SYSTEM AND METHOD FOR A FLUIDIC BARRIER ON THE LOW PRESSURE SIDE OF A FAN BLADE

Rolls-Royce North America...

1. A turbofan engine having a fan portion in fluid communication with a core stream and a bypass stream of air; the core stream
being:
compressed by the fan portion and a core compressor portion, heated and expanded through a core turbine portion;
the core turbine portion driving the fan portion and the compressor portion; the core turbine portion connected to a shaft;
the bypass stream being compressed by the fan portion;
the core and the bypass streams separated by an upstream splitter and a downstream splitter with the fan portion disposed
axially between the upstream and downstream splitters wherein a fluid passage between the core and bypass streams is defined
between the splitters; the fan portion having a plurality of blades; each of the blades of the fan portion having a high pressure
side and a low pressure side; and

a plurality of high pressure fluid jets originating from the low pressure side of the blades restricting the migration of
the core stream into the bypass stream through the fluid passage.

US Pat. No. 9,890,640

GAS TURBINE ENGINE TIP CLEARANCE CONTROL

ROLLS-ROYCE NORTH AMERICA...

1. An apparatus comprising:
a gas turbine engine having a flow path wall disposed around an airfoil shaped component;
a battery;
a first thermoelectric device disposed in thermal communication with a first portion of the flow path wall, the first thermoelectric
device operable to generate heat and thereby affect a thermal transfer between the first thermoelectric device and the first
portion of the flow path wall, wherein the first thermoelectric device is powered by the battery;

a controller adapted to regulate heat generated by the thermoelectric device; and
a second thermoelectric device disposed in thermal communication with a second portion of the flow path wall, the second thermoelectric
device adapted to convert waste heat into electrical power used to charge the battery;

wherein thermal transfer between the first thermoelectric device and the flow path wall affects a change in size of the first
portion of the flow path wall and thereby changes a tip clearance between the flow path wall and the airfoil shaped component.

US Pat. No. 9,863,366

EXHAUST NOZZLE APPARATUS AND METHOD FOR MULTI STREAM AIRCRAFT ENGINE

Rolls-Royce North America...

1. A multi stream aircraft fixed geometry nozzle comprising:
an inner nozzle;
an outer nozzle disposed radially outward of the inner nozzle; and
a supersonic ejector disposed axially aft of the inner nozzle and outer nozzle;
a fan supplying motive fluid to form a bypass stream and a third stream;
the inner nozzle being configured to channel a primary stream of a mixture of a propulsive core stream from an engine core
and the bypass stream from a bypass duct surrounding the engine core, from an aft end of the engine core to the supersonic
ejector; and

the outer nozzle being configured to channel the third stream from an aft end of a third stream duct surrounding the bypass
duct to the supersonic ejector to merge the third stream with the primary stream;

the fixed geometry nozzle including a flow control device comprising a plurality of flaps attached to a radially outer wall
of the third stream duct which deploy radially inward from the radially outer wall, the flow control device configured to
operate the fixed geometry nozzle between an SFC mode and a thrust mode such that, when the inner nozzle accelerates the primary
stream supersonically to the supersonic ejector, at which the primary stream is merged with the third stream, in the SFC mode
a total pressure of the primary stream is substantially the same as a total pressure of the third stream, and in the thrust
mode the total pressure of the primary stream is substantially greater than the total pressure of the third stream.

US Pat. No. 10,100,649

COMPLIANT RAIL HANGER

Rolls-Royce North America...

1. A turbine shroud comprisinga plurality of carrier segments comprising metallic materials and arranged circumferentially adjacent to one another around an axis, each carrier segment including a body and a bracket that extends inwardly in a radial direction from the body toward the axis, and
a plurality of blade track segments comprising ceramic-matrix composite materials and arranged circumferentially adjacent to one another around the axis, each blade track segment including a runner and at least one hanger that extends outwardly in the radial direction from the runner,
wherein at least one of the hangers of each blade track segment engages with the bracket of at least one carrier segment to couple the plurality of blade track segments to the carrier segments, the bracket of each carrier segment is formed to include a plurality of circumferentially spaced apart fingers extending generally axially from the body and arranged to be engaged by the hangers of the blade track segments, and the fingers are configured to flex inward in the radial direction when engaged by the hangers of the blade track segments.

US Pat. No. 9,938,845

GAS TURBINE ENGINE VANE END DEVICES

Rolls-Royce Corporation, ...

6. A gas turbine engine assembly comprisinga turbomachinery component having a wall that defines a flow path of the gas turbine engine assembly,
a rotatable vane positioned adjacent the wall and configured to move relative to the wall, and
a brush seal coupled to the rotatable vane for movement therewith, the brush seal being located between the rotatable vane and the wall, and the brush seal including a base member, a plurality of bristles that extend outwardly away from the base member toward the wall, and a clamp arranged around a portion of the base member and the bristles to couple the bristles with the base member,
wherein the rotatable vane includes a leading edge and a trailing edge spaced apart from the leading edge to define a chord of the rotatable vane, the brush seal extends at least partway along the chord, and a height of the plurality of bristles varies along the chord,
wherein the rotatable vane is moveable between a first position and a second position and at least one portion of the bristles does not contact the wall in at least one of the first position and the second position.

US Pat. No. 9,932,844

SEALS FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A sealing assembly comprising:a support having a support-seal surface,
an engine component having a component-seal surface, the engine component mounted so that the component-seal surface is arranged in spaced-apart confronting relation with the support-seal surface to define a gap between the support and the engine component that grows and shrinks based on the temperature of the support and the engine component, and
a seal adapted to block gasses from passing through the gap between the support and the engine component, the seal including a mount ring coupled to the support and spaced apart from the engine component the mount ring formed to include a plurality of spaced apart pusher arms and a ceramic tadpole gasket having a compressible head and a flat body extending from the compressible head, wherein the compressible head is engaged by the plurality of spaced apart pusher arms and the flat body is formed to include receiver slots that receive the pusher arms therethrough so that the tadpole gasket is coupled to the mount ring, wherein the seal includes a retainer ring that cooperates with the mount ring to trap at least a portion of the flat body between the mount ring and the retainer ring.

US Pat. No. 9,714,608

REDUCED NOISE GAS TURBINE ENGINE SYSTEM AND SUPERSONIC EXHAUST NOZZLE SYSTEM USING ELECTOR TO ENTRAIN AMBIENT AIR

Rolls-Royce North America...

17. An exhaust nozzle system for a gas turbine engine, comprising:
a first nozzle configured to discharge a flow stream of the gas turbine engine, wherein the first nozzle is a fixed-position
nozzle that converges toward a central axis at an aft end of the first nozzle and the flow stream is a combination of an engine
core flow and a first bypass flow stream;

a first ejector in fluid communication with another flow stream discharged by the gas turbine engine and configured to entrain
ambient free stream air into the another flow stream to form a mixed flow; and

a mixer disposed upstream of the first nozzle and configured to mix the engine core flow and the first bypass flow stream.

US Pat. No. 9,945,256

SEGMENTED TURBINE SHROUD WITH SEALS

Rolls-Royce Corporation, ...

1. A segmented turbine shroud that extends around a central axis, the segmented turbine shroud comprisinga first shroud segment including a first carrier segment and a first blade track segment, the first blade track segment made from a ceramic-matrix-composite material and coupled to the first carrier segment,
a second shroud segment arranged circumferentially adjacent to the first shroud segment around the central axis, the second shroud segment including a second carrier segment and a second blade track segment, the second blade track segment made from a ceramic-matrix-composite material and coupled to the first carrier segment, and
a circumferential seal arranged between the first shroud segment and the second shroud segment to block gasses from passing through a circumferential interface of the first shroud segment and the second shroud segment, the circumferential seal including a first seal support coupled to the first shroud segment, a second seal support coupled to the second shroud segment, and a seal element that extends from the first seal support to the second seal support,
wherein the first blade track segment includes an arcuate runner and a first attachment post that extends from the arcuate runner to the first carrier segment.

US Pat. No. 9,909,430

TURBINE DISK ASSEMBLY INCLUDING SEPERABLE PLATFORMS FOR BLADE ATTACHMENT

Rolls-Royce North America...

1. A turbine disk assembly adapted for use in a gas turbine engine, the assembly comprising
a disk having an outer surface, the outer surface including a coupling portion;
an attachment member having a coupling portion and defining at least a portion of an opening, the coupling portion of the
attachment member configured to be coupled to the coupling portion of the disk; and

a blade, a portion of the blade configured to be disposed within the opening when the coupling portion of the attachment member
is coupled to the coupling portion of the disk,

wherein the opening is defined by a single attachment member,
wherein the attachment member includes a platform and a pair of engagement portions extending inward in a radial direction
on both sides of the opening from the platform and the engagement portions are configured block a root portion of the blade
from movement through the opening.

US Pat. No. 9,896,191

FLUID-VECTORING SYSTEM

Rolls-Royce North America...

1. An aircraft comprising
a body and
a fluid-vectoring system coupled to the body and configured to control movement of the body as the body moves along a flight
path during flight of the aircraft, the fluid-vectoring system including a first fluid passageway arranged to extend along
an axis of the body and to define a first fluid cavity therein, an environmental fluid passageway defining an environmental
cavity and arranged to communicate a first flow of environmental fluid in a downstream direction from an environment surrounding
the aircraft through the environmental cavity into the first fluid passageway, and a first fluid-control unit coupled to the
body to move between a retracted configuration in which the first flow of environmental fluid moves downstream from the environment
surrounding the aircraft through the environmental cavity, through the first fluid cavity, and to the environment and an engaged
configuration in which the first fluid-control unit blocks the first flow of environmental fluid from flowing through the
first fluid cavity,

wherein the first fluid-control unit includes a first control door coupled to the body to move between an opened position
in which the first flow of environmental fluid is communicated through the first fluid cavity and a closed position in which
the first control door extends into the first fluid cavity to block communication of the first flow of environmental fluid
through the first fluid cavity,

wherein the body includes a first bypass passageway defining a first bypass cavity, the first bypass passageway is arranged
to communicate a first bypass flow of environmental fluid in the downstream direction from the environment surrounding the
aircraft through the first bypass cavity into the first fluid passageway, and

wherein the first control door is spaced apart in the downstream direction from an inlet of the environmental fluid passageway,
spaced apart in the downstream direction from an inlet of the first bypass passageway, and spaced apart in the upstream direction
from an outlet of the first fluid passageway.

US Pat. No. 10,082,085

SEAL FOR GAS TURBINE ENGINES

Rolls-Royce North America...

1. A gas turbine engine assembly comprisinga support component formed to include a notch,
an engine component, the engine component mounted in spaced-apart relation to the support component so that a gap is formed between the support component and the engine component and so that the notch opens into the gap, and
a seal adapted to close the gap, the seal including a rope gasket arranged to block gasses from passing through the gap and a rope-biasing clip arranged between the support component and the rope gasket to push the rope gasket toward engagement with the engine component, the rope-biasing clip being formed to include a first spring lobe and a second spring lobe that cooperate with notch surfaces defining the notch in the support component to define a cavity,
wherein a first portion of the gap between the support component and the engine component extends from a first side of the notch and is arranged to receive a first portion of the gasses at a first pressure, a second portion of the gap between the support component and the engine component extends from a second side of the notch opposite the first side of the notch and is arranged to receive a second portion of the gasses at a second pressure, and the second pressure is lower than the first pressure such that the notch is configured to be pressurized by the first portion of the gasses from the first portion of the gap that extends from the first side of the notch.

US Pat. No. 9,890,647

COMPOSITE GAS TURBINE ENGINE COMPONENT

ROLLS-ROYCE NORTH AMERICA...

1. A gas turbine engine component, comprising:
a structure formed of a composite material, the structure including:
a flowpath surface operable in a hot gas flowpath of a gas turbine engine;
a cavity spaced apart from the flowpath surface by a thickness of the composite material; and
a cooling opening operative to discharge cooling air into the flowpath,
wherein the cooling opening extends through the structure from the flowpath surface to the cavity, the cooling opening being
defined by a plurality of ultrasonically formed geometric shapes; and

wherein the composite gas turbine engine component is disposed at least partially in the flowpath and/or bounds the flowpath;
wherein the composite material is a ceramic matrix composite (CMC), a metal matrix composite (MMC), and/or a carbon-carbon
composite.

US Pat. No. 10,060,264

GAS TURBINE ENGINE AND COOLED FLOWPATH COMPONENT THEREFOR

Rolls-Royce North America...

1. A turbine flowpath component for a gas turbine engine, comprising:a spar having a suction-side wall extending from a leading edge to a trailing edge and a pressure-side wall extending from the leading edge to the trailing edge, each wall having an outer surface;
a coversheet positioned on the spar to at least partially enclose the spar, the coversheet having an exterior surface and an engagement surface opposite the exterior surface, the exterior surface being an outermost surface of the turbine flowpath component, the engagement surface positioned to face the outer surface of the suction-side wall of the spar and the outer surface of the pressure-side wall of the spar, the engagement surface of the coversheet and the outer surface of the suction-side wall of the spar cooperate to define a suction-side gap therebetween, and the engagement surface of the coversheet and the outer surface of the pressure-side wall of the spar cooperate to define a pressure-side gap therebetween;
a plurality of spacers positioned between the spar and the coversheet in the suction-side gap and in the pressure-side gap; and
a hollow pin extending between a first opening formed in the outer surface of the suction-side wall and a second opening formed in the outer surface of the pressure-side wall, wherein the hollow pin provides fluid communication between the suction-side gap and the pressure-side gap,
wherein the turbine flowpath component is defined by a pressure side and a suction side, the coversheet defines at least a portion of a third opening that is located on the pressure side of the turbine flowpath component, and the third opening provides fluid communication between the pressure-side gap and a core flow surrounding the turbine flowpath component to allow fluid in the pressure-side gap to exit the turbine flowpath component on the pressure side through the third opening.

US Pat. No. 10,030,540

FAN CASE LINER REMOVAL WITH EXTERNAL HEAT MAT

Rolls-Royce North America...

1. A method of replacing a fan case liner in a fan case, the method comprisingapplying heat to a portion of an exterior surface of a fan case to soften an adhesive layer bonding a first fan case liner panel to an inner surface of the fan case,
removing the first fan case liner panel and adhesive residue from the fan case to produce an undisrupted inner surface of the fan case, and
bonding a second fan case liner panel to the undisrupted inner surface of the fan case, wherein
applying heat to the portion of the exterior of the fan case is performed on the portion of the fan case containing the first fan case liner panel.

US Pat. No. 10,030,541

TURBINE SHROUD WITH CLAMPED FLANGE ATTACHMENT

Rolls-Royce North America...

1. A segmented turbine shroud that extends around a central axis, the segmented turbine shroud comprising:a carrier segment formed to include a dovetail slot that opens inwardly in a radial direction toward the central axis,
a blade track segment comprising ceramic-containing materials, the blade track segment being formed to include a runner that extends partway around the central axis and a flange that extends radially outward from the runner, and
a track retention assembly including retainer blocks that receive at least a portion of the flange included in the blade track segment,
wherein the retainer blocks are positioned in the dovetail slot of the carrier segment and cooperate to provide a dovetail shape corresponding to the dovetail slot and being sized to block movement of the track retention assembly out of the dovetail slot,
wherein at least one of the retainer blocks is formed to include a recess that receives at least a portion of the flange of the blade track segment, and
wherein the flange from one side to another side has a constant cross-sectional thickness at each point along a length of the flange that extends radially outward from the runner.

US Pat. No. 9,879,698

NOSE CONE AND SHAFT BALANCING ASSEMBLY

ROLLS-ROYCE NORTH AMERICA...

1. A turbine machine comprising:
a rotatable shaft;
a nose cone having a central axis mounted to said rotatable shaft so that said central axis is axially aligned with said rotatable
shaft, said nose cone comprising a flange extending axially from a leading tip of said nose cone to a trailing edge at a base
of said nose cone and radially around said central axis, said flange having an outer surface defining an airflow path and
one or more apertures; and

a shaft balancing assembly comprising one or more balance weights positioned at least partially in said one or more apertures,
one or more of said balance weights being removable from said apertures while said nose cone is mounted to said shaft,

wherein said flange is formed with a thickness that does not vary by more than fifty percent from said leading tip of said
cone to said trailing edge of said flange, and wherein said shaft balancing assembly further comprises one or more alignment
modules, each module defining one or more bores, said one or more modules adhering to an inner surface of said flange and
being positioned so that each module bore is aligned with a flange aperture to thereby form an aligned pair of a flange aperture
and a module bore, each of said aligned pairs forming a recessed cavity.

US Pat. No. 9,683,443

METHOD FOR MAKING GAS TURBINE ENGINE CERAMIC MATRIX COMPOSITE AIRFOIL

Rolls-Royce North America...

11. A method comprising:
fabricating an airfoil preform as a cooling passage preform having a spanwise extending trailing end portion and a plurality
of flow dividing members projecting from the spanwise extending trailing end portion towards a forward end and defining cooling
passages therebetween;

coupling a cooling delivery core to the forward end of the cooling passage preform to close a forward end of the cooling passages;
covering the cooling passages with a ceramic matrix material; and
trimming the spanwise extending trailing end portion of the cooling passage preform to expose the cooling passages.

US Pat. No. 9,982,676

SPLIT AXIAL-CENTRIFUGAL COMPRESSOR

Rolls-Royce North America...

1. A gas turbine engine comprisinga compressor including an axial compression stage and a centrifugal compression stage arranged aft of the axial compression stage along an engine axis,
a turbine arranged aft of the centrifugal compression stage and coupled to the compressor to drive rotation of the axial compression stage and the centrifugal compression stage about the engine axis, and
a transmission coupled to the turbine and the compressor, the transmission configured to transmit rotational power generated by the turbine about the engine axis to at least one of the axial compression stage and the centrifugal compression stage to drive rotation of at least one of the axial compression stage and the centrifugal compression stage at a first speed offset from a turbine speed,
wherein (i) the axial compression stage has an outlet radius and the centrifugal compression stage has an inlet radius that is about equal to the outlet radius of the axial compression stage to facilitate a smooth transition of air from the axial compression stage to the centrifugal compression stage, (ii) the centrifugal compression stage is coupled to the turbine for common rotation therewith about the engine axis, and (iii) the axial compression stage is coupled to the turbine through the transmission for rotation about the engine axis at the first speed offset from the turbine speed.

US Pat. No. 9,979,339

SYNCHRONOUS ELECTRIC POWER DISTRIBUTION STARTUP SYSTEM

Rolls-Royce North America...

1. A system comprising:a excitation system configured to output a variable excitation signal; and
a synchronous generator configured to generate power for a plurality of rotational synchronous motor loads based on the variable excitation signal;
the excitation system configured to output the variable excitation signal based on a voltage and current being supplied by the generator to the rotational synchronous motor loads;
the excitation system configured, in response to the rotational synchronous motor loads not rotating, to provide pulses of the excitation signal in a first stage and in a second stage;
the excitation system configured to selectively provide repetitive pulses of the variable excitation signal in the first stage at a predetermined frequency to temporarily energize the rotational synchronous motor loads prior to rotation of the generator; and
the excitation system further configured to selectively provide pulses of the variable excitation signal at a variable frequency in the second stage after rotation of the generator commences, the pulses of the variable excitation at the second stage provided to coincide with the generator and the rotational synchronous motor loads being substantially in electrical alignment.

US Pat. No. 9,963,979

COMPOSITE COMPONENTS FOR GAS TURBINE ENGINES

Rolls-Royce North America...

1. A turbine wheel for a gas turbine engine, the turbine wheel comprisinga disk formed to include a dovetail slot that extends through the disk in an axial direction from a forward side to an aft side of the disk and inwardly in a radial direction from an outer diameter of the disk toward a central axis,
a blade comprising ceramic-containing materials, the blade formed to include an airfoil that extends outwardly in the radial direction from the outer diameter of the disk and a root that extends into the dovetail slot, the root including a stem that extends from the airfoil into the dovetail slot and a pair of pin receivers arranged in the dovetail slot that extend circumferentially from the stem in opposing directions, wherein the blade includes a first and a second composite ply and each of the first and the second composite plies include a receiver portion that extends circumferentially away from the stem portion to provide at least part of a pin receiver and an airfoil portion that extends radially outwardly from the root to provide at least part of the airfoil, and
a pair of retention pins each arranged in the pin receivers, the pair of retention pins are arranged on circumferentially opposed sides of the stem and are arranged radially between the pin receivers and the disc so that centrifugal forces applied to the blade when the turbine wheel is rotated about the central axis are transferred through the pair of retention pins to the disk.

US Pat. No. 9,948,216

PRE-ALIGNMENT OF SYNCHRONOUS LOADS PRIOR TO STARTING GRID

Rolls-Royce North America...

1. An apparatus for rotor pre-alignment, the apparatus comprising:a partial power converter configured to provide an alignment current through an n-phase supply line to a synchronous alternating current (AC) motor, wherein the synchronous AC motor is connected to the n-phase supply line, wherein the synchronous AC motor is configured to receive polyphase AC power through the n-phase supply line from a synchronous AC grid, and wherein the partial power converter is powered by a power source isolated from the synchronous AC grid; and
a controller configured to direct the partial power converter to provide the alignment current through the n-phase supply line, wherein the alignment current causes a rotor of the synchronous AC motor to move to and stop at a target angular position, wherein the alignment current is provided to the synchronous AC motor prior to startup of the motor when the polyphase AC power from the synchronous AC grid is substantially zero.

US Pat. No. 9,938,846

TURBINE SHROUD WITH SEALED BLADE TRACK

Rolls-Royce North America...

1. A segmented turbine shroud that extends around a central axis, the segmented turbine shroud comprisinga first shroud segment including a first carrier segment, a first blade track segment, and a first retainer that couples the first carrier segment to the first blade track segment, the first blade track segment made from a ceramic-matrix-composite material,
a second shroud segment arranged circumferentially adjacent to the first shroud segment around the central axis, the second shroud segment including a second carrier segment, a second blade track segment, and a second retainer that couples the second carrier segment to the second blade track segment, the second blade track segment made from a ceramic-matrix-composite material, and
a circumferential seal arranged between the first shroud segment and the second shroud segment to block gasses from passing through a circumferential interface of the first shroud segment and the second shroud segment, the circumferential seal engaging a first seal-locating feature formed in the first carrier segment and a second seal-locating feature formed in the second carrier segment so that the circumferential seal is held in place circumferentially between the first shroud segment and the second shroud segment,
wherein the each carrier segment includes a central attachment body engaged with a corresponding blade track segment and a pair of end caps that extend circumferentially in opposing directions from the central attachment body and each end cap is formed to include a seal-locating feature, and
wherein a closed cavity is formed by the first shroud segment between the first carrier segment and the first blade track segment, the closed cavity extends along the central attachment body of the first carrier segment, the closed cavity is radially bounded by the blade track segment, and the closed cavity is circumferentially bounded by the pair of end caps.

US Pat. No. 9,879,601

GAS TURBINE ENGINE COMPONENT ARRANGEMENT

Rolls-Royce North America...

1. An apparatus comprising
a cooled gas turbine engine component having an outer surface and an internal space for the conveyance of a relatively pressurized
cooling fluid,

a trench formed in the cooled gas turbine engine component having an upstream side disposed below and at an angle relative
to the outer surface of the cooled gas turbine engine component and a downstream side that intersects the upstream side of
the trench at a bottom of the trench, the downstream side being continuous without holes, and

a plurality of cooling holes configured to exit substantially perpendicular to the upstream side of the trench intermediate
the outer surface of the cooled gas turbine engine component and the bottom of the trench, each of the plurality of cooling
holes being j-hook shaped and including an upstream portion in proximity to the internal space, a downstream portion spaced
apart from the upstream portion and having an exit formed in the upstream side of the trench, and an intermediate curved portion
extending between the upstream portion and the downstream portion, and the downstream portion of each of the plurality of
cooling holes is spaced apart from the bottom of the trench.

US Pat. No. 10,125,622

SPLAYED INLET GUIDE VANES

ROLLS-ROYCE NORTH AMERICA...

1. A system for directing the flow of a fluid and controlling the rate of flow of the fluid, said system comprising:a channel for directing the flow of the fluid;
at least a pair of articulating vanes positioned within said channel for controlling the flow rate of the fluid within said channel, each of said vanes comprising a pair of lateral major surfaces forming a leading edge and a trailing edge of said vane, and an axis of articulation intersecting said vane at a point spaced from the aerodynamic center of said vane; and
a linkage between said vanes coupling the articulation of each of said vanes to the other of said vanes, wherein each vane imparts a force on said linkage when the relative angle of attack is greater than zero, wherein the force imparted on said linkage by one of said vanes is at least partially cancelled by the force imparted on the linkage by the other of said vanes during the articulation of said vanes, wherein the axis of articulation of one of said vanes intersects said vane between the aerodynamic center and said leading edge of said vane, and the axis of articulation of the other of said vanes intersects said other vane between the aerodynamic center and said trailing edge of said other vane.

US Pat. No. 10,087,770

SHROUD CARTRIDGE HAVING A CERAMIC MATRIX COMPOSITE SEAL SEGMENT

Rolls-Royce Corporation, ...

1. A segmented turbine shroud for radially encasing a turbine in a gas turbine engine, the shroud comprising:a carrier comprising a portion defining a pin-receiving carrier bore;
a ceramic matrix composite (CMC) seal segment comprising an arcuate flange having a surface facing the turbine and a portion defining a pin-receiving seal segment bore;
an elongated pin extending through said carrier bore and said seal segment bore,
wherein said elongated pin comprises a lateral cross-sectional dimension of at least three-eighths inches; and
wherein said CMC seal segment portion defining the pin-receiving seal segment bore is radially spaced from said arcuate flange by a spacing flange extending radially outward from said arcuate flange to thereby effect receipt within the seal segment bore of said elongated pin, said spacing flange having an axial dimension and a circumferential dimension, wherein said axial dimension is greater than said circumferential dimension of said spacing flange.

US Pat. No. 10,001,084

AIRCRAFT POWERPLANT WITH MOVEABLE NOZZLE MEMBER

Rolls-Royce North America...

1. An apparatus comprising:a gas turbine engine having a core flow passage and a fan bypass passage that together are merged into a merged flow passage;
an annular shaped third stream bypass passage at a proximal end of the apparatus which is unwrapped from a coaxial axis with the merged flow passage and ducted to an underslung configuration to form a third stream nozzle passage at a distal end of the apparatus; and
a nozzle that receives the merged flow passage and the third stream nozzle passage, the nozzle having a dual-use moveable member disposed between the merged flow passage and the third stream nozzle passage and structured to change an area of the merged flow passage and the third stream nozzle passage whereby movement of the dual-use moveable member increases a flow area of the merged flow passage while it decreases a flow area of the third stream nozzle passage.

US Pat. No. 9,739,202

THERMAL ADJUSTMENT MEMBER FOR A FUEL NOZZLE OF A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A fuel nozzle for use in a gas turbine engine, comprising:
a fuel nozzle tip having an outer tip housing, wherein the outer tip housing has a recess formed at an inner surface of the
outer tip housing;

a shim disposed within the recess, wherein a surface of the shim contacts a diaphragm of the fuel nozzle tip; and
a thermal adjustment member disposed within the recess against the shim, wherein the outer tip housing, diaphragm and shim
are formed of a base material having a linear coefficient of thermal expansion (?1) and the thermal adjustment member is formed of a different material having a linear coefficient of thermal expansion (?2) higher than the linear coefficient of thermal expansion (?1) of the base material.

US Pat. No. 10,100,659

HANGER SYSTEM FOR A TURBINE ENGINE COMPONENT

Rolls-Royce North America...

14. A method of assembling a component of a turbine, the method comprising:advancing a seal ring segment forward to engage a hanger of the seal ring segment with a rail of a carrier segment,
positioning a retainer aft the hanger of the seal ring segment, and
securing the retainer to the carrier segment, via a fastener extending through the carrier segment into a bottom surface of a groove in the retainer such that the hanger is secured between the carrier segment and the retainer.

US Pat. No. 10,024,537

COMBUSTOR ASSEMBLY WITH CHUTES

Rolls-Royce North America...

1. A combustor for use in a gas turbine engine, the combustor comprising:a combustion liner that defines a combustion chamber, the combustion liner including an outer liner surface facing away from the combustion chamber, an inner liner surface facing toward the combustion chamber, and a chute-receiving aperture that extends through the outer liner surface and the inner liner surface, and
a chute that extends through the chute-receiving aperture of the combustion liner and defines a passageway sized to convey air from an environment outside the combustion chamber through the combustion liner into the combustion chamber,
wherein the chute includes a chute body that extends through the chute-receiving aperture and defines the passageway, a flared head located outside of the combustion chamber that extends outwardly from the chute body away from the passageway so that the flared head is sized to block movement of the chute through the combustion liner into the combustion chamber, and a locating shoulder located inside of the combustion chamber that extends outwardly from the chute body away from the passageway so that the locating shoulder is sized to block movement of the chute through the combustion liner away from the combustion chamber,
wherein the combustion liner includes a plurality of cooling holes that extend through the combustion liner and that are arranged around the chute-receiving aperture and the locating shoulder is formed to include a plurality of cooling scallops arranged to face away from the chute body to provide a flow path for the air to flow from the environment outside of the combustion chamber through the plurality of cooling holes into the combustion chamber.

US Pat. No. 9,878,798

AIRCRAFT WITH COUNTER-ROTATING TURBOFAN ENGINES

Rolls-Royce North America...

1. An aircraft comprising
a frame,
a first turbofan engine coupled to the frame, the first turbofan engine including a first turbine, a first fan coupled to
the first turbine to be driven by rotation of the first turbine, and a first transmission coupled between the first turbine
and the first fan to transmit rotation from the first turbine to the first fan, the first transmission having a star gearset,
and

a second turbofan engine coupled to the frame, the second turbofan engine including a second turbine, a second fan coupled
to the second turbine to be driven by rotation of the second turbine, and a second transmission coupled between the second
turbine and the second fan to transmit rotation from the second turbine to the second fan, the second transmission having
a planetary gearset,

wherein the first turbine and the second turbine are configured to rotate in a first direction, the first transmission is
configured to transmit rotation from the first turbine to the first fan to cause rotation of the first fan in the first direction,
and the second transmission is configured to transmit rotation from the second turbine to the second fan to cause rotation
of the second fan in a second direction opposite the first direction.

US Pat. No. 9,714,609

GAS TURBINE ENGINE AND ELECTRIC MACHINE

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine comprising:
a rotatable turbomachinery component comprising a plurality of blades;
a surface defining a flow path boundary;
an electrical machine, wherein the electrical machine comprises a stator portion and a rotor portion configured to interact
and produce electrical power when the rotor portion is rotated relative to the stator portion, wherein the stator portion
comprises conductive coils, wherein the rotor portion of the electrical machine is integrated with the rotatable turbomachinery
component, wherein the rotor portion comprises a magnetic field element extending between respective blades of the plurality
of blades and located radially inward from the surface defining a flow path boundary, and wherein the conductive coils are
located radially outwardly of the flow path boundary such that the flow path boundary protects the conductive coils from a
working fluid that flows through the turbomachinery component; and

a magnetic field adjuster, wherein the magnetic field adjuster comprises a moveable member capable of altering a magnetic
field that interacts with the conductive coils, wherein the moveable member is urged from a first position to a second position
via centripetal acceleration, and wherein the moveable member is configured to move in both a radial and an axial direction
along a conical surface of a blade of the plurality of blades to alter the magnetic field as a function of rotational speed
of the shaft.

US Pat. No. 9,803,553

METHOD TO CONTROL ELECTRIC STARTER GENERATOR FOR GAS TURBINE ENGINES

Rolls-Royce North America...

1. A gas turbine engine starting system comprising:
an electric start generator (ESG) free of temperature sensors and configured to provide torque to a gas turbine engine, wherein
the ESG comprises a plurality of subcomponents;

a fuel metering module configured to provide a quantity of fuel to the gas turbine engine; and
an electronic control system (ECS) that:
determines a future temperature of the ESG based on a plurality of historical ESG thermal trending information and an input
ambient temperature;

determines whether at least one of an ongoing start and an uninitiated start will be unsuccessful, wherein the determination
of whether at least one of the ongoing start and the uninitiated start will be unsuccessful is based on the future temperature
of the ESG; and

indicates the determination of whether at least one of the ongoing start and the uninitiated start will be unsuccessful to
an operator.

US Pat. No. 10,141,874

SYNCHRONOUS ELECTRICAL POWER DISTRIBUTION SYSTEM STARTUP AND CONTROL

ROLLS-ROYCE NORTH AMERICA...

1. A system comprising:a prime mover configured to provide mechanical energy to the system by spinning a shaft;
a synchronous AC generator comprising a rotor mechanically coupled to the shaft;
an exciter mechanically coupled to the shaft and configured to output a variable field current to excite the synchronous AC generator;
a plurality of synchronous electric motors electrically direct coupled to the synchronous AC generator and each comprising a rotor rotatable operable to drive one or more mechanical loads; and
a controller configured to establish and maintain a magnetic coupling between the rotor of the synchronous AC generator and all of the rotors of the synchronous electric motors by control of a level of the field current during a ramped increase in rotation of the rotor of the synchronous AC generator from zero rotational speed based on a difference in an angle of deflection between a position of the rotor of the synchronous AC generator and a position of the rotors of the synchronous electric motors.

US Pat. No. 10,132,194

SEAL SEGMENT LOW PRESSURE COOLING PROTECTION SYSTEM

Rolls-Royce North America...

1. A turbine shroud comprisinga plurality of carrier segments comprising metallic material and arranged circumferentially adjacent to one another around an axis,
a plurality of blade track segments comprising ceramic-matrix composite material and arranged circumferentially adjacent to one another around the axis, each blade track segment coupled to one of the carrier segments,
a plurality of thin-walled tubes each defining an internal cooling air plenum, each thin-walled tube extending into one of the carrier segments and configured to direct a flow of cooling air toward a radially-outward facing side of the blade track segment, and
a plurality of track-segment couplers, each track-segment coupler coupled to one of the carrier segments and configured to hold one of the blade track segments on the carrier segment,
wherein each track-segment coupler is configured to receive one of the thin-walled tubes to hold the thin-walled tube in place relative to the carrier segment.

US Pat. No. 10,100,654

IMPINGEMENT TUBES FOR CMC SEAL SEGMENT COOLING

Rolls-Royce North America...

1. A turbine shroud comprisinga plurality of carrier segments comprising metallic material and arranged circumferentially adjacent to one another around an axis,
a plurality of blade track segments comprising ceramic-matrix composite material and arranged circumferentially adjacent to one another around the axis, each blade track segment coupled to one of the carrier segments, and
a plurality of impingement tubes, each impingement tube extending into one of the carrier segments and configured to direct a flow of cooling air toward a radially-outward facing side of the blade track segment,
wherein each blade track segment includes a runner and at least two attachment features extending radially outward from the runner, the at least two attachment features axially spaced apart from one another and circumferentially extending along the runner, and
wherein each impingement tube is positioned between one of the at least two attachment features and the runner of a corresponding blade track segment.

US Pat. No. 9,879,861

GAS TURBINE ENGINE WITH IMPROVED COMBUSTION LINER

Rolls-Royce Corporation, ...

1. A gas turbine engine variable porosity combustor liner comprising:
a laminated alloy structure having combustion chamber facing holes on one side and cooling plenum facing holes on a radially
opposite side, the combustion chamber facing holes being in fluid communication with the cooling plenum facing holes via axially
and circumferentially extending flow passages disposed between metal alloy sheets of the laminated alloy structure;

wherein at least one of the combustion chamber facing holes and the cooling plenum facing holes are oriented at a first angle
with respect to an axis of the combustor liner;

wherein porous zones having respective different cooling flow amounts are formed in the laminated alloy structure;
wherein the combustion chamber facing holes include holes of a first diameter and holes of a second diameter that is different
from the first diameter; and

wherein the holes oriented at the first angle are diffusion shaped holes, the diffusion shaped holes having an entrance at
a first size, and an exit at a second size that is greater than the first size.

US Pat. No. 9,982,541

GAS TURBINE ENGINE FLOW PATH MEMBER

Rolls-Royce North America...

1. A turbine blade adapted for use in a gas turbine engine, the turbine blade comprisingan airflow device having an interior volume configured to receive cooling fluid during use of the turbine blade in a gas turbine engine,
a rubbing tip set back from an edge of the airflow device,
and cooling openings sized to discharge cooling fluid from the interior volume of the airflow device, the cooling openings defined at least in part by the rubbing tip and at least in part by the airflow device, and the cooling openings located around an entire periphery of the rubbing tip including a pressure side of the airflow device and a suction side of the airflow device,
wherein the airflow device includes a base located toward an end of the airflow device, the base is arranged to enclose the interior volume of the airflow device, the rubbing tip extends outwardly away from the base beyond an end of the airflow device, each cooling opening includes an entrance that opens directly into the interior volume, and the rubbing tip defines at least a portion of each of the entrances.

US Pat. No. 10,048,168

SYSTEM AND METHOD FOR OPTIMIZING COMPONENT LIFE IN A POWER SYSTEM

Rolls-Royce North America...

1. An aircraft, comprising:first and second gas turbine engines, each having a first power circuit and a second power circuit, the first power circuit having a first aircraft component coupled thereto, and the second power circuit having a second aircraft component coupled thereto that is redundant with the first aircraft component; and
an engine health monitoring system (EHMS) coupled to the first and second gas turbine engines, wherein the EHMS:
receives sensor feedback from the first aircraft component and the second aircraft component;
generates a remaining useful life (RUL) of the first aircraft component and of the second aircraft component;
generates a rate-of-life consumption of the first aircraft component and of the second aircraft component;
indicates when failure of the first aircraft component and failure of the second aircraft component will occur based on the RUL of each and based on the respective rate-of-life consumption of each;
provides instructions for altering operation of the aircraft that affects the RUL of the first aircraft component or the second aircraft component; and
the EHMS varies an amount of load on the first power circuit and the second power circuit to increase the RUL of the aircraft component with the lowest RUL, by reducing an amount of load on the power circuit having the aircraft component with the lowest RUL and increasing an amount of load on the other power circuit, such that the first aircraft component and the second aircraft component each have a generally converged time of failure.

US Pat. No. 10,094,239

VANE ASSEMBLY FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A vane assembly for a gas turbine engine, the assembly comprisingan inner platform made from a metallic material,
an outer platform made from a metallic material,
a ceramic-containing airfoil that extends from the inner platform to the outer platform and engaged with at least one of the inner platform and the outer platform so that some aerodynamic loads applied to the ceramic-containing airfoil are transferred directly to at least one of the inner platform and the outer platform, and
a reinforcement spar made from a metallic material that extends from the inner platform to the outer platform through a hollow core of the ceramic-containing airfoil and engages an interior surface of the ceramic-containing airfoil so that some aerodynamic loads applied to the ceramic-containing airfoil are transferred to at least one of the inner platform and the outer platform,
wherein the ceramic-containing airfoil includes a first end and a second end and the interior surface of the ceramic-containing airfoil engages the reinforcement spar adjacent to the second end of the ceramic-containing airfoil,
wherein the first end of the ceramic-containing airfoil is received in one of the inner platform and the outer platform to transfer load,
wherein the ceramic-containing airfoil is disengaged from the inner platform adjacent to the second end of the ceramic-containing airfoil.

US Pat. No. 9,963,990

CERAMIC MATRIX COMPOSITE SEAL SEGMENT FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A segmented turbine shroud for radially encasing a turbine in a gas turbine engine, the shroud comprising:a carrier;
a ceramic matrix composite (CMC) seal segment comprising an arcuate flange having a surface facing the turbine and a portion defining a bore for receiving an elongated pin; and
one or more elongated pins,
wherein said CMC seal segment is carried by said carrier by at least one of said elongated pins being received within said bore,
wherein said CMC seal segment portion defining a pin-receiving bore comprises a curved outer surface and is radially spaced from said arcuate flange by an elongated spacing flange extending radially outward from said arcuate flange, said elongated spacing flange having an axial dimension parallel to the axis of said bore and a lateral dimension less than the axial dimension of said flange and less than a lateral dimension of said portion defining the bore, and wherein said CMC seal segment portion defines a bore having a length that is at least 70% of the length of said elongated pin received therein.

US Pat. No. 9,790,893

GAS TURBINE ENGINE FLOW DUCT HAVING INTEGRATED HEAT EXCHANGER

Rolls-Royce North America...

7. A flow duct for a gas turbine engine comprising:
the flow duct disposed along a centerline of the gas turbine engine and defining a flow duct passage;
a plurality of heat transfer components integrated in the flow duct and configured to have a variable geometry arrangement
for adjusting static pressure in the flow duct passage downstream of the heat transfer components; and

an adjustable nozzle positioned within the flow duct, the adjustable nozzle having a first nozzle wall assembly and a second
nozzle wall assembly that define a nozzle span, such that air passing through the flow duct passes through the nozzle span;

wherein:
the first nozzle wall assembly includes a first nozzle wall and the second nozzle wall assembly includes a second nozzle wall;
and

the first and second nozzle wall assemblies are each adjustable to define the nozzle span between the first nozzle wall and
the second nozzle wall;

further comprising an adjustable diffuser positioned downstream of the adjustable nozzle, the adjustable diffuser including
a first diffuser wall assembly and a second diffuser wall assembly, the first diffuser wall assembly having a first diffuser
wall and the second diffuser wall assembly having a second diffuser wall, wherein the first and second diffuser wall assemblies
are each adjustable to adjust a diffuser span between the first diffuser wall and the second diffuser wall;

wherein each of the first and second nozzle wall assemblies are four-body linkages, each having pivotal joints that adjust
the nozzle span between the first nozzle wall and the second nozzle wall; and

wherein each of the first and second diffuser wall assemblies are four-body linkages, each having pivotal joints that adjust
the diffuser span between the first diffuser wall and the second diffuser wall.

US Pat. No. 9,920,634

METHOD OF MANUFACTURING A TURBOMACHINE COMPONENT, AN AIRFOIL AND A GAS TURBINE ENGINE

Rolls-Royce Corporation, ...

13. An airfoil for a turbomachine, comprising:
a body having an airfoil shape, wherein the body includes:
an internal portion of the airfoil having a composite foam cooling passage configured to pass a cooling fluid into the composite
foam cooling passage at a first end face and out of the composite foam cooling passage at a second end face, the first end
face and the second end face positioned on a same side of the body; and

a plurality of composite wrap plies enveloping the internal portion of the airfoil.

US Pat. No. 9,882,513

SYNCHRONIZING MOTORS FOR AN ELECTRIC PROPULSION SYSTEM

Rolls-Royce North America...

1. An electric propulsion system comprising:
an AC drive circuit that includes:
a plurality of propulsor motors;
an AC power bus; and
an AC generator that delivers AC electrical power to the AC power bus for simultaneously driving the plurality of propulsor
motors;

a synchronization circuit configured to individually synchronize, with the AC generator, a single propulsor motor from the
plurality of propulsor motors, one at a time; and

a control unit configured to:
determine whether the single propulsor motor is not synchronized with the AC generator; and
maintain synchronicity between the single propulsor motor and the AC generator by engaging the synchronization circuit with
the single propulsor motor in response to determining that the single propulsor motor is not synchronized with the AC generator.

US Pat. No. 9,853,335

THERMAL MANAGEMENT OF ENERGY STORAGE

Rolls-Royce North America...

1. An energy storage thermal management system comprising:
an energy storage compartment including a liquid coolant bath portion and a vapor portion;
a plurality of energy storage cells positioned within said energy storage compartment and submerged within said liquid coolant
bath;

a compressor in communication with said vapor portion, said compressor removing vapor from said vapor portion; and
a condenser in communication with said compressor, said condenser returning liquid coolant to said energy storage compartment;
at least one distribution manifold positioned between an adjoining pair of said plurality of energy storage cells;
at least one expansion valve positioned on said distribution manifold between said adjoining pair of energy storage cells,
said at least one expansion valve allowing said liquid coolant bath to flow through said at least one distribution manifold
towards regions of localized heat.

US Pat. No. 9,803,485

TURBINE SEGMENTED COVER PLATE RETENTION METHOD

Rolls-Royce North America...

1. A gas turbine engine disc and cover plate assembly comprising:
a turbine disc having a first groove and a retaining arm including a radially extending portion;
a turbine blade having a second groove;
at least one segmented cover plate having an upper end and a lower end, the at least one segmented cover plate dimensioned
such that when the upper end is positioned within the second groove the lower end clears the first groove;

at least one retainer member positioned within the first groove of the turbine disc below the lower end of the segmented cover
plate such that the at least one retainer member is positioned against the bottom of the first groove and the radially extending
portion, the retainer member having a contact surface comprising a notched portion for retaining the lower end of the segmented
cover plate in place relative to the turbine disc such that the segmented cover plate is secured between the first groove
and the second groove; and

at least one smash plate positioned within the first groove of the turbine disc;
wherein axial retention of the turbine blade relative to the turbine disc is aided by the smash plate; and
wherein the notched portion is configured to force the at least one segmented cover plate into the second groove and inwards
against the turbine blade.

US Pat. No. 9,937,803

ELECTRIC DIRECT DRIVE FOR AIRCRAFT PROPULSION AND LIFT

Rolls-Royce North America...

1. An apparatus of an aircraft propulsion system, comprising:a prime mover coupled to a generator;
an electric motor directly electrically coupled to said generator;
a propeller coupled to and driven by said electric motor; and
a bi-directional power converter coupled to said generator and further coupled to an energy storage device;
wherein said energy storage device is selectively coupled to said electric motor, and
wherein said energy storage device comprises an engine generator.

US Pat. No. 9,771,945

GAS TURBINE ENGINE HAVING CONFIGURABLE BYPASS PASSAGE

Rolls-Royce North America...

1. An apparatus comprising
a gas turbine engine having a core flow path and a bypass duct structured to bypass a working flow around a combustor of the
gas turbine engine;

a moveable component associated with the bypass duct and structured to change a flow area by moving between a first position
and a second position, the moveable component actuated to the second position in the duct such that a geometry of the duct
forms a resonance space tuned to attenuate a noise, and

an actuator coupled to the moveable component and energized by a controller and a sensor structured to detect a noise information,
the sensor in communication with the controller, wherein the controller is structured to operate upon the noise information
when energizing the controller.

US Pat. No. 9,915,202

GAS TURBINE ENGINE HEAT EXCHANGER SYSTEM

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine including an inlet particle separator that separates inlet air into scavenge air and clean air, the clean
air passing into the gas turbine engine;

a heat transfer system;
a scavenge air path that conveys all the scavenge air from the inlet particle separator to the heat transfer system;
a heat exchange fluid path that conveys a heat exchange fluid to the heat transfer system and away from the heat transfer
system, wherein the heat transfer system transfers heat from the heat exchange fluid path to the scavenge air path to cool
the heat exchange fluid;

a flow control mechanism positioned between the heat transfer system and the inlet particle separator such that all of the
scavenge air passes through the flow control mechanism; and

a blower positioned downstream from the heat transfer system such that the scavenge air that passes through the heat transfer
system also passes through the blower;

wherein the flow control mechanism controls a flow of the scavenge air to modulate a temperature of the heat exchange fluid;
and

wherein the flow control mechanism is a pump.

US Pat. No. 10,208,668

TURBINE ENGINE ADVANCED COOLING SYSTEM

ROLLS-ROYCE CORPORATION, ...

1. A gas turbine engine defining a longitudinal axis, the gas turbine engine comprising:a compressor;
a turbine;
a combustor arranged axially between the compressor and the turbine and configured to drive the turbine with a stream of gas, the combustor comprising at least one liner defining a combustion chamber;
a shield positioned axially between the compressor and the turbine, wherein the shield extends from a first plane positioned axially forward of the combustor to a second plane extending through an aft end of the combustion chamber, the first plane and the second plane each being perpendicular to the longitudinal axis;
a combustor case;
a drive shaft mechanically coupling the compressor to the turbine;
a conduit configured to supply a cooling fluid to an annulus defined by the combustor case and the shield, the annulus configured to supply the cooling fluid to a first end of a gap between an outer surface of the drive shaft and the shield, the drive shaft and the shield configured to guide the cooling fluid from the first end of the gap to a second end of the gap, wherein the cooling fluid flows from the second end of the gap into an outlet of the compressor during operation of the gas turbine engine;
a flow restrictor positioned in the gap and located between the first end of the gap and the second end of the gap, wherein the flow restrictor is a seal, wherein the seal is configured to limit fluid flow through the gap; and
a radial pre-swirler configured to direct the cooling fluid from the annulus toward the drive shaft and into the gap, the radial pre-swirler configured to swirl the cooling fluid in a direction of rotation of the drive shaft, the radial pre-swirler positioned at an aft end of the shield and axially aft of the flow restrictor.

US Pat. No. 10,174,619

GAS TURBINE ENGINE COMPOSITE VANE ASSEMBLY AND METHOD FOR MAKING SAME

Rolls-Royce North America...

1. A method for forming a gas turbine engine airfoil assembly, comprising:providing at least two gas turbine engine airfoil composite preform components,
interlocking the airfoil composite preform components with a first locking component, and
interlocking the first locking component and at least one of the airfoil composite preform components with a second locking component, and
after interlocking with each of the first and second locking components, rigidizing the assembly of aircraft component preform components, the first locking component, and the second locking component,
wherein the airfoil composite preform components comprise an airfoil and an endwall, and the first locking component interlocks the airfoil and the endwall and the second locking component interlocks the first locking component and the endwall, and the second interlocking comprises inserting the second locking component into a through hole in the first locking component and a through hole in the endwall.

US Pat. No. 9,958,159

COMBUSTOR ASSEMBLY FOR A GAS TURBINE ENGINE

Rolls-Royce Corporation, ...

1. A combustor assembly for a gas turbine engine, comprising: a hanger having a first flange portion extending radially inward therefrom, said first flange portion forming a step including a first wall portion at a first wall thickness and a second wall portion at a second wall thickness that is less than said first wall thickness;a combustor liner fixed to said hanger, said combustor liner having an inner surface extending along an axis of a combustion chamber;
a heat shield having a-second flange portions extending radially outward therefrom, said second flange portion in contact with said first wall portion and said second wall portion of said first flange portion, said heat shield at least partially overlapping and confronting said inner surface of said combustor liner along said axis to form an axially extending gap, said heat shield releasably engaged with said hanger, said heat shield forming a plurality of radially outer surfaces between said second flange portions, said radially outer surfaces oppose said first wall portion to form circumferentially extending gaps between said second flange portions, said heat shield forming recesses that oppose said second wall portion to form radially extending gaps, such that air flow passes from the circumferentially extending gaps into said radially extending gaps and into said axially extending gap; and
fasteners passing through said first flange portion and said second flange portions and extending in a direction parallel with said axis.

US Pat. No. 9,790,864

PROGNOSTIC HEALTH MANAGEMENT APPROACHES FOR PROPULSION CONTROL SYSTEM

Rolls-Royce North America...

15. An aircraft, comprising: an engine; and a controller system having a system controller with a processor, the controller
system being coupled to an engine and configured to control demand on a first and second component of the engine, the controller
system is configured to: operate the engine in a control mode associated with one of the first component and the second component;
identify, by way of the system controller, a set point reference of the first component; identify, by way of the system controller,
a set point reference of the second component; identify, by way of the system controller, a data set indicative of a level
of deterioration of the engine by at least referencing at least one reference schedule that associates a state of at least
one of the first component and the second component with a given thrust; and based on the level of deterioration, determine
and implement a trim scheme, which includes increasing the set point reference of the first component if the control mode
is a rotor speed control mode, and decreasing the set point reference of the control mode is one of an engine pressure ratio
control mode and a turbine fan power ratio control mode; and notify a user to change the set point reference of the second
component wherein the set point reference of the first component is a fan speed set point reference and the set point reference
of the second component is one of an engine pressure ratio set point reference and a turbo fan power ratio set point reference.

US Pat. No. 10,060,277

TURBINE WHEEL WITH CLAMPED BLADE ATTACHMENT

Rolls-Royce North America...

1. A turbine wheel for a gas turbine engine, the turbine wheel comprising a disk formed to include a fir tree slot that extends through the disk in an axial direction from a forward side to an aft side of the disk and inwardly in a radial direction from an outer diameter of the disk toward a central axis,a blade comprising ceramic-containing materials, the blade integrally formed to include an airfoil that extends outwardly in the radial direction from the outer diameter of the disk and a root that extends into the fir tree slot, the root including a stem that extends from the airfoil into the fir tree slot and a retention head arranged in the fir tree slot that extends from the stem, and
a blade retention assembly including at least two clamp blocks that cooperate to form a fir tree shape corresponding to the fir tree slot, wherein the at least two clamp blocks engage the retention head of the root and are received in the fir tree slot of the disk so that centrifugal forces applied to the blade when the turbine wheel is rotated about the central axis are transferred from the retention head through the blade retention assembly to the disk, wherein the clamp blocks are shaped to extend radially inward to a point substantially similar to a radially-innermost surface of the blade, and wherein the retention head extends in axially forward and aft direction out from the stem.

US Pat. No. 10,060,300

SECTIONED GAS TURBINE ENGINE DRIVEN BY SCO2 CYCLE

Rolls-Royce North America...

1. A sectioned heat exchanger system for a gas turbine engine, comprising:an inlet manifold configured to receive a working fluid;
a plurality of circuits including at least first and second circuits configured to transfer heat with respect to the working fluid, each of the first and second circuits having a circuit inlet valve, a circuit heat exchange channel, and a circuit outlet valve;
an outlet manifold configured to pass the working fluid to an outlet;
a first sensor configured to measure a first parameter of the first circuit;
a second sensor configured to measure a second parameter of at least one of the outlet and the second circuit; and
a controller of the gas turbine engine and configured to selectively isolate at least one of the plurality of circuits by closing at least one of the first and second circuits in response to detecting a leak based on a parameter difference between the first and second parameters.

US Pat. No. 10,038,397

MULTIPLE ENGINE CONDITION MATCHING VIA ELECTRICAL POWER EXTRACTION CONTROL

Rolls-Royce Corporation, ...

1. A multi-engine power system comprising:a load requiring a total amount of electrical power;
a first engine configured to provide a first portion of the total amount of electrical power to be provided to the load;
a second engine configured to provide a second portion of the total amount of electrical power to be provided to the load; and
a controller configured to:
determine the total amount of electrical power to be provided to the load;
estimate a respective service time associated with each of the first and second engines; and
control each of the first and second engines to provide the total amount of electrical power to the load and to coordinate the respective service times associated with the first and second engines.

US Pat. No. 9,840,967

RAM AIR THERMAL MANAGEMENT SYSTEM

Rolls-Royce North America...

1. A thermal management system for an aircraft having an engine and a heat generating component, at least one of which generates
a heat load, the engine having an engine fan configured to draw in an engine inlet air stream, at least a portion of which
is to be used as an engine air stream downstream of the engine fan, the thermal management system comprising:
a cooling circuit configured to circulate a fluid through the heat load such that at least a portion of the heat load is transferrable
to the fluid;

a heat exchanger in fluid communication with the cooling circuit, the heat exchanger being configured to enable heat transfer
between the fluid and a cooling air stream, the heat exchanger being located upstream of the engine fan;

a plenum starting at an air outlet side of the heat exchanger; and
a pumping device located within the plenum and configured to draw the cooling air stream through the heat exchanger and into
a portion of the engine air stream downstream of the engine fan.

US Pat. No. 10,190,434

TURBINE SHROUD WITH LOCATING INSERTS

Rolls-Royce North America...

1. A turbine blade track comprisingan annular ceramic runner formed to include a plurality of cutouts extending inward in a radial direction from an outer radial surface of the annular ceramic runner toward an inner radial surface of the annular ceramic runner, and
a plurality of inserts coupled to the annular ceramic runner, each insert including a stem arranged in the cutout and a cap arranged outside the cutout that extends from the stem in a circumferential direction and in an axial direction along the outer radial surface of the annular ceramic runner.

US Pat. No. 10,180,071

COMPOSITE BLADES FOR GAS TURBINE ENGINES

Rolls-Royce North America...

1. A turbine wheel for a gas turbine engine, the turbine wheel comprisinga disk formed to include a dovetail slot that extends through the disk in an axial direction from a forward side to an aft side of the disk and inwardly in a radial direction from an outer diameter of the disk toward a central axis, and
a blade comprising ceramic-matrix materials, the blade formed to include an airfoil that extends outwardly in the radial direction from the outer diameter of the disk and a root that extends into the dovetail slot,
wherein the root includes a stem that extends from the airfoil into the dovetail slot, a root core, and a root casing that extends from the stem around the root core to couple the root core to the stem, the root casing comprising ceramic-matrix materials and positioned to engage an inner surface of the dovetail slot formed by the disk to retain the blade in place relative to the disk during rotation of the disk,
wherein portions of the airfoil and root casing are formed by at least one continuous ply of ceramic-containing material having a first end forming a portion of the airfoil and extending radially inward from the airfoil to wrap around the root core and extending along itself back radially outward to a second end also forming a portion of the airfoil.

US Pat. No. 10,151,219

GAS TURBINE ENGINE AND FRAME

ROLLS-ROYCE NORTH AMERICA...

1. A gas turbine engine frame, comprising:a metallic inner hub;
a metallic flange;
a metallic outer construction; and
a composite flowpath structure comprising a primary flowpath structure disposed between the metallic inner hub and the metallic outer construction, wherein the composite flowpath structure includes at least one of carbon bismaleimide composites, ceramic matrix composites, metal matrix composites, organic matrix composites or carbon-carbon composites, and wherein the metallic flange is configured to secure the composite flowpath structure to the metallic inner hub;
wherein the primary flowpath structure comprises a primary composite outer flowpath wall and a primary composite inner flowpath wall spaced radially apart from the primary composite outer flowpath wall, and the primary composite outer flowpath wall and the primary composite inner flowpath wall together define the primary flowpath structure for a working fluid of the gas turbine engine;
wherein the primary flowpath structure is configured to direct the working fluid to a compressor of a gas turbine engine; and
wherein the composite flowpath structure further comprises a plurality of inner composite struts, wherein at least a portion of each inner composite strut extends between the primary composite inner flowpath wall and the primary composite outer flowpath wall.

US Pat. No. 9,976,514

PROPULSIVE FORCE VECTORING

Rolls-Royce North America...

1. An apparatus for vectoring a propulsive force imparted to an object, the apparatus comprising:a fluid acceleration unit adapted to eject a fluid from a nozzle to thereby provide a first component of said propulsive force, the nozzle defining an exit area substantially perpendicular to a path of the ejected fluid;
a first array of rotatable members disposed in the path of said ejected fluid, each of said rotatable members being fully rotatable about a respective rotation axis extending through the respective member, wherein a flow of said ejected fluid around said rotatable members when spinning provides a second component of said propulsive force, the rotatable members of the first array positioned downstream of the nozzle and distributed substantially equally across the exit area; and
one or more motors adapted to spin one or more of said rotatable members in said first array in a first rotational direction to vector said propulsive force in a first direction and further adapted to spin one or more of said rotatable members in said first array in a second rotational direction to vector said propulsive force in a second direction, wherein all the propulsive force is provided by manipulation of the fluid;
further comprising a second array of rotatable members disposed in the path of said ejected fluid and downstream of said first array, wherein the flow of said ejected fluid over the rotatable members of said second array when said members of the second array are spinning provides a third directional component to said propulsive force.

US Pat. No. 9,932,851

ACTIVE SYNCHRONIZING RING

Rolls-Royce NOrth America...

1. A system for controlling individual vane angles in a gas turbine engine, comprising:a synchronization ring having a plurality of micro-actuators coupled to said synchronization ring;
a plurality of strain sensors on a face of the synchronization ring to measure distortion of the synchronization ring and communicate the measured distortion by way of a control signal of a controller in communication with the plurality of micro-actuators such that each of said micro-actuators of the plurality of actuators create a bending moment in said synchronization ring about an axis substantially parallel with a central axis of the gas turbine engine to counter the measured distortion of the synchronization ring, each of said micro-actuators having an input for receiving the control signal; and
an actuator pivotally attached to said synchronization ring for rotating said ring about the central axis.

US Pat. No. 9,752,607

RETENTION PIN AND METHOD OF FORMING

Rolls-Royce North America...

1. A retention pin assembly comprising:
a stud including a head, a shaft, and a deformable end disposed at a distal end thereof, wherein the shaft includes a stepped
surface having a greater diameter than the deformable end;

a housing wall defining a base;
a support member extending transversely from the base, the support member having an opening extending therethrough for receiving
the deformable end, the opening of the support member further including a pocket on one side of the support member;

a nub disposed on the support member between the base and the opening, the nub extending transversely to the base along the
support member towards the opening and projecting outwardly from the support member on a side opposite the pocket;

a leaf seal pivotally mounted on the shaft; and
a bias member coaxially arranged on the shaft and disposed between the head and the leaf seal, wherein the bias member exerts
a force on the leaf seal;

wherein the deformable end is deformed to occupy the pocket, and wherein the deformable end and the stepped surface provide
a mechanical retainer to secure the stud to the support member; and

wherein the nub is configured to position the leaf seal apart from a distal end of the support member.

US Pat. No. 9,845,768

THREE STREAM, VARIABLE AREA, VECTORABLE NOZZLE

Rolls-Royce North America...

16. An exhaust nozzle for an engine comprising:
a fan that generates first, second, and third streams of air, wherein the first and second streams of air diverge downstream
of at least one compressor of the engine;

the first stream of air that exits a core of the engine;
the second stream of air that traverses a length of the nozzle and is directed to a plenum of the nozzle, wherein the first
and second streams of air combine rejoin downstream of a combustor of the engine and prior to passing through a turbine of
the engine; and

the third stream of air is separated from the second stream of air, the third stream of air traverses the length of the engine
and is injectable to the plenum of the nozzle, the third stream of air includes fan bypass air and is conditioned by a heat
exchanger, the exhaust nozzle including a primary convergent flap and a secondary divergent flap that is attached to the primary
convergent flap, the primary convergent flap pivotal inwardly and outwardly to define a throat area of the exhaust nozzle,
and the secondary divergent flap pivotably attached to the primary convergent flap, the secondary divergent flap pivotable
to control a flow rate of the third stream of air.

US Pat. No. 10,215,056

TURBINE SHROUD WITH MOVABLE ATTACHMENT FEATURES

Rolls-Royce Corporation, ...

1. A segmented turbine shroud that extends around a central axis, the segmented turbine shroud comprisinga carrier segment that extends partway around the central axis and that forms a radially inwardly-opening cavity,
a blade track segment comprising ceramic-containing materials, the blade track segment formed to include a runner that extends partway around the central axis and a positioner attachment post that extends radially outward from the runner into the radially inwardly-opening cavity of the carrier segment, the positioner attachment post formed to include a track-positioning surface that extends both radially and axially, and
a track attachment system adapted to couple the blade track segment to the carrier segment, and the track attachment system including a positioner coupled to the carrier segment to move axially from a disengaged position out of contact with the positioner attachment post to an engaged position contacting the positioner attachment post to engage the track-positioning surface of the positioner attachment post with a position-setting surface that extends both radially and axially at an angle corresponding to that of the track-positioning surface.

US Pat. No. 10,205,415

MULTIPLE GENERATOR SYNCHRONOUS ELECTRICAL POWER DISTRIBUTION SYSTEM

Rolls-Royce North America...

1. A power system comprising:a first controller configured to control a first generator;
a second controller configured to control a second generator, the second generator electrically coupled with the first generator; and
a plurality of rotational loads electrically coupled with the first generator and the second generator;
the first controller configured to excite the first generator to generate alternating current (AC) electric power at a time of commencement of rotation of the first generator;
the second controller configured to excite the second generator at the time of commencement of rotation of the first generator such that the second generator is energized to operate as a motor in response to receipt of the AC power generated by the first generator; and
the second generator and the plurality of rotational loads configured to commence rotation with the first generator at the time of commencement of rotation of the first generator due to receipt of the AC electric power.

US Pat. No. 10,184,352

TURBINE SHROUD SEGMENT WITH INTEGRATED COOLING AIR DISTRIBUTION SYSTEM

Rolls-Royce North America...

1. A turbine shroud segment adapted for use in a gas turbine engine, the turbine shroud segment comprisinga carrier segment comprising metallic materials, the carrier segment formed to include an attachment-receiving space,
a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend at least partway around a central axis and an attachment box portion that extends radially outward from the runner into the attachment-receiving space formed by the carrier segment, and
an attachment assembly including an attachment post that extends from the carrier segment into the attachment box portion of the blade track segment, the attachment post formed to include a post passageway that provides part of a cooling system configured to conduct cooling air into the attachment box portion of the blade track segment to cool the blade track segment when the turbine shroud segment is used in a gas turbine engine, wherein the attachment box portion included in the blade track segment defines a box interior bounded by a radially-outwardly facing surface of the runner included in the blade track segment so that cooling air conducted into the box interior by the post passageway cools the radially-outwardly facing surface of the runner, and wherein the box interior defined by the attachment box portion of the blade track segment is open for fluid communication with the attachment-receiving space formed by the carrier segment and the carrier segment is formed to include a plurality of vent holes configured to conduct used cooling air out of the attachment-receiving space.

US Pat. No. 9,982,607

SHAFT FAILURE DETECTION USING PASSIVE CONTROL METHODS

Rolls-Royce North America...

1. A method for use in a turbine control system, the method comprising:a) controlling fuel supply to a gas turbine engine at least in part using a fuel supply limit determined as a first function of a rotational speed of a shaft of the gas turbine engine.
b) obtaining a first value representative of a rotational speed of the shaft;
c) differentiating the first value within a processing unit;
d) employing the processing unit to determine an adjusted fuel supply limit as an adjusted function of the first value, the adjusted function based on the first function and the differentiated first value; and
e) controlling the fuel supply to the gas turbine engine at least in part using the adjusted fuel supply limit.

US Pat. No. 9,952,044

MULTI-AXIS CALIBRATION BLOCK

Rolls-Royce North America...

1. A system comprising:a measurement table;
a calibration block for calibrating a touch probe, the calibration block comprising:
a calibration block body forming a bored hole providing a concave measurement surface;
a magnetic base comprising one or more permanent magnets and a manual release mechanism configured to facilitate releasing the magnetic base from the measurement table; and
a three dimensional object protruding from the calibration block body and providing a convex measurement surface, wherein the convex measurement surface provides opposing measurement contact points in at least two dimensions;
a five-axis mechanical holding arm configured to manipulate a touch probe to measure the calibration block and a component; and
a computing device configured to:
send control signals to the five-axis mechanical holding arm to locate the touch probe mounted in the five-axis mechanical holding arm relative to the calibration block;
send control signals to the five-axis mechanical holding arm to measure the three dimensional object with a distal tip of the touch probe by contacting multiple points of the three dimensional object;
send control signals to the five-axis mechanical holding arm to measure the bored hole with sides of the distal tip of the touch probe by contacting multiple points of the concave measurement surface;
generate calibration factors for the touch probe by comparing the sizes of the three dimensional object and the bored hole as measured by manipulating the touch probe with the five-axis mechanical holding arm with predefined actual sizes of the three dimensional object and the bored hole;
store the calibration factors for the touch probe in a non-transitory computer readable medium;
after storing the calibration factors for the touch probe, send control signals to the five-axis mechanical holding arm to measure features of the component with the distal tip of the touch probe, wherein the calibration block and the component are both secured to the measurement table during the measurement of the calibration block and the measurement of the component;
store values of the measured features of the component in the non-transitory computer readable medium, the values being based on the calibration factors;
after storing values of the measured features of the component in the non-transitory computer readable medium, send control signals to the five-axis mechanical holding arm to again measure at least one of the three dimensional object and the bored hole with the touch probe;
compare the sizes of the at least one of the three dimensional object and the bored hole as measured by manipulating the touch probe with the five-axis mechanical holding arm with the predefined actual sizes of the at least one of the three dimensional object and the bored hole to generate updated calibration factors for the touch probe; and
in the event that the updated calibration factors are substantially different than the calibration factors, update the stored values of the measured features of the component in the non-transitory computer readable medium based on the updated calibration factors.

US Pat. No. 9,885,369

VARIABLE VANE FOR GAS TURBINE ENGINE

ROLLS-ROYCE NORTH AMERICA...

1. An apparatus, comprising:
a vane;
a rotation support coupled to an end of the vane;
a spindle coupled to the rotation support, wherein the spindle, the vane, and the rotation support are rotationally aligned;
an annular sleeve engaging the spindle, wherein the annular sleeve contacts the rotation support at a radially inward extent
and contacts a turbine casing at a radially outward extent of the annular sleeve;

a first rolling element directly engaging the annular sleeve substantially near the radially outward extent, wherein the first
rolling element is coupled to the turbine casing and the annular sleeve is disposed between the first rolling element and
the spindle; and

a second rolling element engaging the annular sleeve substantially near the radially inward extent, wherein the second rolling
element is coupled to an outer endwall ring and wherein a center of mass of the annular sleeve is positioned between the first
and second rolling elements.

US Pat. No. 10,087,852

FUEL FLOW SPLITTER AND GAS TURBINE FUEL SYSTEM HEALTH MONITORING

Rolls-Royce North America...

1. A gas turbine health monitoring fuel system comprising:a fuel splitter;
a first fuel pump downstream from the fuel splitter;
a first fuel nozzle fluidly coupled to the first fuel pump by a first fuel stream;
a second fuel pump downstream from the fuel splitter;
a second fuel nozzle fluidly coupled to the second fuel pump by a second fuel stream; a common drive coupled to the first and second fuel pumps and configured to drive the first and second fuel pumps at a same speed;
a controller system configured to:
identify a first fuel pressure in the first fuel stream;
identify a second fuel pressure in the second fuel stream; and
determine a difference between the first and second fuel pressures to diagnose if at least one of the fuel nozzles is faulty; and
a primary fuel source configured to supply a primary fuel stream to the fuel splitter,
wherein the controller system is further configured to:
identify a primary fuel pressure in the primary fuel stream;
compare the first fuel pressure to the primary fuel pressure;
compare the second fuel pressure to the primary fuel pressure;
identify a difference between at least one of the first and second fuel pressures and the primary fuel pressure; and
notify a user that at least one of the fuel nozzles is faulty.

US Pat. No. 9,982,629

ENGINE DRIVEN BY SC02 CYCLE WITH INDEPENDENT SHAFTS FOR COMBUSTION CYCLE ELEMENTS AND PROPULSION ELEMENTS

Rolls-Royce Corporation, ...

1. A gas turbine engine, comprising:a first shaft coupled to a first turbine and a first compressor;
a second shaft coupled to a second turbine and a second compressor;
a third shaft coupled to a third turbine and a fan assembly;
a heat rejection heat exchanger configured to reject heat from a closed loop system with bypass air passed from the fan assembly to provide cooling to the heat rejection heat exchanger;
a combustor positioned to receive compressed air from the second compressor as a core stream;
wherein the closed-loop system includes the first, second, and third turbines and the first compressor and receives energy input from the combustor; and
wherein the closed-loop system is configured to provide power from the combustor to the first, second, and third turbines.

US Pat. No. 9,879,862

GAS TURBINE ENGINE AFTERBURNER

Rolls-Royce North America...

1. An apparatus comprising:
a gas turbine engine having a core flow path, a bypass flow path, and an afterburner flame holder; and
a toroidal afterburner combustor pilot structured to receive working fluid from the bypass flow path, the toroidal afterburner
combustor pilot oriented around an axis of revolution and having:

an annular inlet positioned on a first lateral side to feed a flow of working fluid to a top region of the toroidal afterburner
combustor pilot, the annular inlet having a plurality of swirler vanes oriented to impart a circumferential flow component
to the working fluid;

a combustion chamber displaced laterally from the annular inlet and shaped to receive the working fluid at the top region,
the combustion chamber having a curved far wall that acts to turn the working fluid downward and form a circumferential vortex
of flow; and

an outlet positioned on the first lateral side of the toroidal afterburner combustor pilot and located radially beneath the
annular inlet, the outlet structured to deliver products of combustion in a radially inward direction into the afterburner
flame holder.

US Pat. No. 9,759,090

GAS TURBINE ENGINE COMPONENT HAVING FOAM CORE AND COMPOSITE SKIN WITH COOLING SLOT

Rolls-Royce North America...

1. An apparatus comprising:
a cooled gas turbine engine airfoil member having an outer skin that forms a gas path surface of the cooled gas turbine engine
airfoil member, the outer skin including a composite material having a matrix material;

a foam core also having the matrix material and disposed internal to the composite material of the outer skin, the foam core
occupying substantially an entirety of space internal to the composite material of the outer skin; and

a foam based discharge slot having the matrix material also located at a trailing edge portion of the cooled gas turbine engine
airfoil member wherein a width of the slot along a span of the airfoil member is greater than a height of the airfoil member
in a thickness direction of the airfoil member, the foam core defining the foam based discharge slot and having a cooling
passage through which cooling fluid exits the cooled gas turbine engine airfoil member.

US Pat. No. 10,233,764

FABRIC SEAL AND ASSEMBLY FOR GAS TURBINE ENGINE

Rolls-Royce North America...

1. A fabric seal for sealing between components of a gas turbine engine, the fabric seal comprisinga number of free tows each having a length and extending to a free end, and
a number of cross tows each extending across the length of the free tows and woven together therewith,
wherein the fabric seal defines a seal aperture, and the free ends of the free tows terminate within the seal aperture to provide a fringe and are configured for compliant contact with a component of a gas turbine engine inserted into the seal aperture to provide fluid sealing around the component.

US Pat. No. 10,215,035

TURBINE WHEELS WITH PRELOADED BLADE ATTACHMENT

Rolls-Royce North America...

1. A wheel assembly for a gas turbine engine, the assembly comprisinga disk arranged for rotation about a central axis, the disk formed to include a plurality of slots circumferentially arranged adjacent one another,
a plurality of blades, the plurality of blades including roots sized to be received in the plurality of slots so that the plurality of blades are coupled to the disk for common rotation about the central axis, and
a plurality of blade biasers positioned in the plurality of slots between the disk and the roots of each of the plurality of blades, the blade biasers being engaged with the disk and the roots of the plurality of blades to preload the plurality of blades away from the central axis when the wheel assembly is at rest and reduce the range of centrifugal loads experienced by the disk and the plurality of blades during rotation of the wheel assembly within the gas turbine engine,
wherein the plurality of blade biasers move in an aft direction during rotation of the wheel assembly within the gas turbine engine.

US Pat. No. 10,196,904

TURBINE ENDWALL AND TIP COOLING FOR DUAL WALL AIRFOILS

ROLLS-ROYCE NORTH AMERICA...

1. An airfoil for a gas turbine engine, the airfoil comprising:a spar comprising a passageway inside of the spar for a cooling fluid, a pedestal on an outer surface of the spar, and an inlet configured to direct the cooling fluid from the passageway to the outer surface of the spar; and
a coversheet, wherein an inner surface of the coversheet is positioned on the pedestal of the spar, wherein an edge of the coversheet is positioned along a tip of the spar,
wherein the inner surface of the coversheet, the pedestal, and the outer surface of the spar define a cooling path from the inlet to an outlet at the edge of the coversheet, wherein the coversheet, the pedestal, and the outer surface of the spar define an opening of the outlet, wherein the outlet is configured to direct the cooling fluid onto the tip of the spar, wherein the edge of the coversheet is flush with the tip of the spar and wherein the cooling path is unobstructed from the inlet to the outlet.

US Pat. No. 10,190,418

GAS TURBINE ENGINE AND TURBINE BLADE

Rolls-Royce North America...

1. A turbine blade for a gas turbine engine, comprising:an airfoil body having a pressure side, a suction side, a leading edge portion having a leading edge of the airfoil body, and a trailing edge portion having a trailing edge of the airfoil body, wherein the pressure side, leading edge portion, suction side, and trailing edge portion collectively form a continuous outer surface of the airfoil body, and the airfoil body culminates at a tip surface; and
a squealer tip extending outwardly from the tip surface, said squealer tip having a pressure side rail portion extending along the pressure side wall from the leading edge portion towards the trailing edge portion and a suction side rail portion extending along the suction side wall from the leading edge portion to the trailing edge, said pressure side rail portion and suction side rail portion forming a cavity therebetween on the tip surface of the airfoil body;
wherein the squealer tip includes a passage extending between the pressure side rail portion and the suction side rail portion, said passage configured to fluidly couple the trailing edge portion to the cavity; and
wherein the entire pressure side rail portion and the entire suction side rail portion are respectively offset from the pressure side and the suction side of the airfoil body between the leading edge portion and the trailing edge portion, to define a shelf on the tip surface between each of the pressure and suction sides and the pressure side rail portion and the suction side rail portion, respectively, wherein the suction side rail portion of the squealer tip extends to the trailing edge and is coincident with the continuous outer surface only at the trailing edge of the trailing edge portion, and the pressure side rail portion terminates at an end thereof near the trailing edge portion, and defines the passage as a gap between the end of the pressure side rail portion and the suction side rail portion.

US Pat. No. 9,889,807

VEHICLE AND SYSTEM FOR SUPPLYING ELECTRICAL POWER TO A VEHICLE ELECTRICAL LOAD

Rolls-Royce North America...

1. A system of a vehicle, the system comprising:
an electrical load of the vehicle including a high energy device that utilizes above 270 volts during operations of the vehicle;
a generator coupled to an engine of the vehicle and configured to generate electrical power at a voltage above 270 volts for
the electrical load of the high energy device during operations of the vehicle; and

first and second conduits arranged along each other to house respective first and second conductors that are electrically
disposed between the electrical load and the generator.

US Pat. No. 10,196,919

TURBINE SHROUD SEGMENT WITH LOAD DISTRIBUTION SPRINGS

Rolls-Royce North America...

1. A turbine shroud segment adapted for use in a gas turbine engine having a central axis, the turbine shroud segment comprisinga carrier segment comprising metallic materials, the carrier segment formed to include an attachment-receiving space,
a blade track segment comprising ceramic matrix composite materials, the blade track segment formed to include a runner shaped to extend at least partway around the central axis and an attachment portion that extends radially outward from the runner into the attachment-receiving space formed by the carrier segment,
an attachment assembly including an attachment post that extends from the carrier segment through an attachment hole formed in the attachment portion of the blade track segment, an attachment support coupled to the attachment post to block withdrawal of the attachment post through the attachment hole, and a load distributor configured to distribute clamp force applied by the attachment post and the attachment support along the attachment portion of the blade track segment.

US Pat. No. 10,047,632

RADIALLY STACKED INTERSHAFT BEARING

Rolls-Royce North America...

1. An intershaft bearing for a gas turbine engine, comprising:an outer stubshaft including outer and inner diameter surface portions disposed concentrically with respect to one another;
wherein the outer diameter surface portion is rotatably connected to a body of the gas turbine engine and the inner diameter surface portion is rotatably connected to an inner stubshaft of the gas turbine engine, so as to provide a stacked configuration of the inner stubshaft, the outer stubshaft, and the body of the gas turbine engine along a radial direction that is orthogonal to a longitudinal axis of the gas turbine engine; and
wherein the outer stubshaft includes first and second passages in respective fluid communication with the outer and inner diameter surface portions.

US Pat. No. 10,329,924

TURBINE AIRFOILS WITH MICRO COOLING FEATURES

Rolls-Royce North America...

1. An airfoil for use in a gas turbine engine and having a pressure side and a suction side, the airfoil comprisinga spar formed to define a cooling air plenum adapted to receive a flow of cooling air, and
a skin coupled to an exterior surface of the spar and positioned to at least partially cover the spar along the pressure side and the suction side,
wherein at least one axially extending groove is formed in the exterior surface of the spar on the pressure side that defines at least one cooling passageway between the spar and the skin, at least one inlet port is formed in the spar adjacent a trailing edge of the spar, the at least one inlet port is in fluid communication with the cooling air plenum and the at least one cooling passageway to pass the flow of cooling air into the at least one cooling passageway from the cooling air plenum, at least one outlet port is formed through the skin on the pressure side and axially forward of the at least one inlet port, the at least one outlet port is configured to pass the flow of cooling air from the at least one cooling passageway to an exterior of the airfoil, and at least one turbulator is positioned within the at least one cooling passageway,
at least a second, axially extending groove formed in the exterior surface of a tail section of the spar and defining at least one second cooling passageway between the spar and skin
at least a second inlet port is formed in the spar and in fluid communication with the cooling air plenum and the at least one second cooling passageway to pass a second portion of the flow of cooling air into the at least one second cooling passageway from the cooling air plenum,
a radially extending separator wall is defined between the at least one cooling passageway and the at least one second cooling passageway and configured to separate the flow of cooling air within the at least one cooling passageway from the second portion of the flow of cooling air within the at least one second cooling passageway, and at least one outlet slot is defined between the spar and the skin and configured to pass the second portion of the flow of cooling air from the at least one second cooling passageway to an exterior of the airfoil.

US Pat. No. 10,132,529

THERMAL MANAGEMENT SYSTEM CONTROLLING DYNAMIC AND STEADY STATE THERMAL LOADS

Rolls-Royce Corporation, ...

18. A method of operating a thermal management system, comprising:thermally coupling a dynamic trans-critical cooling circuit with a steady-state trans-critical cooling circuit via a thermal energy storage (TES) system, wherein each of the cooling circuits has a respective compressor, heat rejection exchanger, and expansion device;
receiving a dynamic load in the TES;
cooling the TES to absorb thermal energy by the TES when the dynamic thermal load is ON; and
cooling the TES with the steady-state cooling circuit when the dynamic thermal load is OFF.

US Pat. No. 10,215,028

TURBINE BLADE WITH HEAT SHIELD

Rolls-Royce North America...

1. A turbine-blade assembly adapted for use in a gas turbine engine, the turbine-blade assembly comprising a root including a root platform and a stem adapted to attach the turbine-blade assembly to the gas turbine engine for rotation about a central axis of the gas turbine engine, an airfoil including a spar comprising metallic materials and a heat shield comprising ceramic materials, the spar extending outward from the root platform and formed to include a core body, a tail forming a trailing edge of the airfoil, and an airfoil retainer extending outwardly away from the core body, the heat shield shaped to extend around the core body to form a leading edge, a forward portion of a pressure side of the airfoil, and a forward portion of a suction side of the airfoil, the heat shield located in spaced-apart relation to the core body at all locations to define cooling passages between the spar and the heat shield, and a tip shroud spaced-apart from the root and coupled to the spar to block radial movement of the heat shield relative to the spar, wherein the core body includes a suction side portion and a pressure side portion, the suction side portion having an inner circumference located on a suction side of the airfoil and the pressure side portion having an inner circumference located on a pressure side of the airfoil, the suction side portion formed to include a suction side pocket and the pressure side portion formed to include a pressure side pocket, wherein the heat shield further includes a suction side segment and a pressure side segment spaced apart from the suction side segment, the suction side segment having an outer circumference located on the suction side of the airfoil and the pressure side segment having an outer circumference located on the pressure side of the airfoil, the suction side segment formed to include a first extension extending from the suction side segment towards the core body and the pressure side segment formed to include a second extension extending from the pressure side towards the first extension and the core body, wherein the first and second extensions are positioned in closer relation to the trailing edge of the airfoil than the leading edge of the airfoil, wherein the suction side pocket and the pressure side pocket of the core body are positioned substantially opposite each other, the suction side pocket sized to receive the first extension and the pressure side pocket sized to receive the second extension, wherein the heat shield is arranged so a portion of the first extension of the suction side segment is located within a portion of the suction side pocket of the core body to form a first dovetail joint and a portion of the second extension of the pressure side segment is located within a portion of the pressure side pocket of the core body to form a second dovetail joint.

US Pat. No. 10,066,550

FAN BY-PASS DUCT FOR INTERCOOLED TURBO FAN ENGINES

Rolls-Royce North America...

1. A system for providing intercooling to an engine, comprising:a fan bypass duct communicating with fan bypass air, said fan bypass duct having a fan bypass duct nozzle and disposed about a centerline of the engine;
a secondary bypass duct radially inwardly spaced from said fan bypass duct, said secondary bypass duct having an inlet with a diffuser for decelerating air passing therethrough, said secondary bypass duct in communication with said fan bypass duct for directing a portion of the fan bypass air from the fan bypass duct to said secondary bypass duct, said secondary bypass duct having an outlet in communication with the bypass air at a location downstream of a throat of said fan bypass duct nozzle and comprising a secondary bypass duct nozzle disposed at said outlet of said secondary bypass duct; and
one of a microchannel or minichannel heat exchanger disposed about an entire annulus of the engine and within said secondary bypass duct to receive the portion of the fan bypass air that passes into said secondary bypass duct, and oriented non-perpendicularly with respect to the centerline of the engine, said one of a microchannel or minichannel heat exchanger in communication with a heat transfer fluid.

US Pat. No. 9,816,963

HIGH PRESSURE COMPRESSOR THERMAL MANAGEMENT

Rolls-Royce North America...

8. A method of assembling a gas turbine engine, comprising:
positioning an inner shaft to extend along a rotational axis of the gas turbine engine;
positioning disks to extend radially inward toward the inner shaft;
forming a first hole in a first of the disks and a second hole in a second of the disks that is adjacent to the first of the
disks;

forming a center bore aperture in each of the first and second disks; and
positioning an obstruction in the center bore aperture of the first of the disks, such that a bore flow that flows along the
rotational axis and along the inner shaft is obstructed from flowing along the shaft by the obstruction, and flows radially
outward from the obstruction, through the first hole, radially inward toward the inner shaft, and through the center bore
aperture of the second disk, the obstruction extending through the center bore apertures of the first and second disk.