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. 9,534,537

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. 9,458,855

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. 9,573,853

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. 9,752,592

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. 9,777,627

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. 9,759,079

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. 9,771,870

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. 9,845,688

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. 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. 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. 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. 9,874,110

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. 9,845,700

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. 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. 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. 10,125,622

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. 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. 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. 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. 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. 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. 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. 9,938,198

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. 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. 9,948,216

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. 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. 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. 9,908,635

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. 9,863,366

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. 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. 9,945,256

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. 9,938,845

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. 9,938,846

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. 9,909,430

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. 9,797,270

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. 10,100,654

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. 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. 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. 9,896,191

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. 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. 9,879,601

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. 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. 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. 10,001,084

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. 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. 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. 9,979,339

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. 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. 9,982,676

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. 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. 9,982,541

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. 10,141,874

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. 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. 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. 10,190,434

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. 10,174,619

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. 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. 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. 10,208,668

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. 10,233,764

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. 10,215,056

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. 10,184,352

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. 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. 10,215,035

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. 10,205,415

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. 10,215,028

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. 10,190,418

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. 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. 10,563,579

AIR-INLET DUCT HAVING A PARTICLE SEPARATOR AND AN AGGLOMERATOR FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. An air-inlet duct for a gas turbine engine, the air-inlet duct comprisinga particle separator formed to include an inlet passageway for receiving a stream of air, a compressor passageway that extends downstream from the inlet passageway, and a scavenge passageway that extends downstream from the inlet passageway and that is positioned radially outward of the compressor passageway, the particle separator configured to receive atmospheric air laden with fine particles and large particles and to direct the large particles into the scavenge passageway while allowing the atmospheric air to move into the compressor passageway thereby reducing the number of large particles that enter the compressor passageway, and
an agglomerator configured to emit an electro-magnetic field into the inlet passageway to charge the fine particles in the inlet passageway to cause the fine particles to be attracted to one another and cluster together to form large particles directed into the scavenge passageway by the particle separator to reduce the number of fine particles directed into the compressor passageway,
wherein the particle separator includes an outer wall spaced apart from an engine rotation axis, an inner wall located between the outer wall and the engine rotation axis, the inner wall and the outer wall defining the inlet passageway therebetween, and a splitter located between the outer wall and the inner wall and including an outer splitter surface that cooperates with the outer wall to define the scavenge passageway therebetween and an inner splitter surface cooperating with the inner wall to define the compressor passageway therebetween,
wherein the inner wall, the outer wall, and the splitter include electromagnetically conductive material and the agglomerator includes an electrical charge source configured to electrically charge at least one of the inner wall, the outer wall, and the splitter to emit the electro-magnetic field into the inlet passageway.

US Pat. No. 10,267,177

TURBINE ASSEMBLY HAVING A ROTOR SYSTEM LOCK

Rolls-Royce North America...

1. A turbine assembly for use in a gas turbine engine, the assembly comprisinga case including a support housing and a plurality of vanes coupled to the support housing,
a rotor mounted in the case to rotate relative to the case, the rotor including a wheel and a plurality of blades coupled to the wheel, and
a fuse that extends from at least one vane of the plurality of vanes included in the case to at least one blade of the plurality of blades included in the rotor to block rotation of the rotor relative to the case and that is configured to disintegrate in response to the application of a predetermined current so that the rotor is allowed to rotate relative to the case.

US Pat. No. 10,247,040

TURBINE SHROUD WITH MOUNTED FULL HOOP BLADE TRACK

Rolls-Royce North America...

1. A gas turbine engine comprisinga turbine case arranged around a central axis of the gas turbine engine and formed to include a plurality of outer keyways extending in a radial direction through the turbine case,
a turbine shroud axially aligned with the turbine case and including (i) an annular carrier arranged around the central axis of the gas turbine engine and formed to include a plurality of outer pin receivers and a plurality of inner keyways, wherein the annular carrier includes a plurality of bosses that extend radially outward away from an outer radial carrier surface of the annular carrier and each boss is formed to include one of the outer pin receivers, and (ii) a one-piece annular runner formed to include a plurality of inner pin receivers extending in a radial direction from an outer radial runner surface toward an inner radial runner surface of the one-piece annular runner,
a plurality of outer insert pins, each outer insert pin arranged to extend through one of the outer keyways formed in the turbine case into a corresponding one of the plurality of outer pin receivers formed in the annular carrier to locate the turbine case and the annular carrier relative to the central axis while allowing radial growth of the turbine case and the annular carrier at different rates during use of the gas turbine engine, and
a plurality of inner insert pins, each inner insert pin arranged to extend through one of the inner keyways formed in the annular carrier into a corresponding one of the plurality of inner pin receivers formed in the one-piece annular runner to locate the annular carrier and the one-piece annular runner relative to the central axis while allowing radial growth of the annular carrier and the one-piece annular runner at different rates during use of the gas turbine engine.

US Pat. No. 10,280,872

SYSTEM AND METHOD FOR A FLUIDIC BARRIER FROM THE UPSTREAM SPLITTER

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 core compressor portion; the core turbine portion connected to a shaft;
the bypass stream being compressed by the fan portion;
the core stream and the bypass stream separated by an upstream splitter and a downstream splitter with the fan portion disposed axially between the upstream splitter and the downstream splitter wherein a fluid passage between the core stream and the bypass stream is defined between the upstream splitter and the downstream splitter and a plurality of blades of the fan portion; and
a first plurality of high pressure fluid jets originating from a trailing edge of the upstream splitter restricting migration of the core stream into the bypass stream through the fluid passage; and
a plurality of orifices positioned circumferentially on the trailing edge of the upstream splitter, wherein the first plurality of high pressure fluid jets are injected through the plurality of orifices and directed toward the plurality of blades of the fan portion with at least an axial component relative to a centerline of the shaft.

US Pat. No. 10,287,906

TURBINE SHROUD WITH FULL HOOP CERAMIC MATRIX COMPOSITE BLADE TRACK AND SEAL SYSTEM

Rolls-Royce North America...

1. A gas turbine engine comprisinga blade track arranged around a central axis of the gas turbine engine, the blade track having a leading edge, a trailing edge axially spaced apart from the leading edge, a radial outer surface that extends between the leading and trailing edges, and a radial inner surface spaced apart from the radial outer surface,
a support assembly arranged around the blade track to support the blade track relative to the central axis, the support assembly being formed to define a retention cavity configured to receive relatively high-pressure compressor air that is directed to the leading edge of the blade track to resist gasses from flowing past the leading edge over the radial outer surface of the blade track, a vent cavity configured to receive relatively low-pressure compressor air that is directed to the trailing edge of the blade track to resist the gasses from flowing past the trailing edge over the radial outer surface of the blade track, and a bleed cavity fluidly interconnecting the retention cavity with the vent cavity and configured to receive intermediate-pressure compressor air, and
a seal system including a first seal configured to block fluid communication between the retention cavity and the bleed cavity and a second seal configured to block fluid communication between the bleed cavity and the vent cavity so that a pressure difference across either of the first and second seals is less than a pressure difference between the retention cavity and the vent cavity.

US Pat. No. 10,240,476

FULL HOOP BLADE TRACK WITH INTERSTAGE COOLING AIR

Rolls-Royce North America...

1. A turbine shroud for use in a gas turbine engine having a central axis, the turbine shroud comprisingan annular carrier formed to define a radially inwardly-opening carrier channel that extends around the central axis and the annular carrier includes an outer pin receiver that extends through the annular carrier and opens into the carrier channel to allow pressurized cooling air to pass through the annular carrier into the carrier channel and a high-pressure cooling air passageway that extends radially through the annular carrier,
a one-piece annular runner aligned axially with the carrier channel of the annular carrier, the one-piece annular runner includes an inner radial runner surface located radially between the annular carrier and the central axis and an outer radial runner surface located radially between the inner radial runner surface and the annular carrier, and the outer radial runner surface cooperates with the annular carrier to form an annular buffer chamber between the annular carrier and the one-piece annular runner, and
a cooling system including an annular impingement plate positioned in the annular buffer chamber to separate the annular buffer chamber into an outer chamber and an inner chamber located radially between the outer chamber and the one-piece annular runner, the outer pin receiver opens into the outer chamber to direct the pressurized cooling air into the outer chamber, and the annular impingement plate includes a plurality of diffusion holes spaced circumferentially around the annular impingement plate and each diffusion hole extends radially through the impingement plate to direct the pressurized cooling air in the outer chamber through the annular impingement plate into the inner chamber and toward the outer radial runner surface of the one-piece annular runner,
wherein the one-piece annular runner includes a forward section, an aft section spaced apart axially from the forward section, and a midsection extending between the forward section and the aft section, the high-pressure cooling air passageway is configured to direct high-pressure air toward the forward section of the one-piece annular runner, the high-pressure air has a greater pressure than the pressurized cooling air, and the turbine shroud further includes a first seal positioned radially between the one-piece annular runner and the annular carrier and positioned axially between the high-pressure cooling air passageway and the annular buffer chamber.

US Pat. No. 10,294,954

COMPOSITE BLISK

ROLLS-ROYCE NORTH AMERICA...

1. A composite turbomachine comprising:a hub comprised of fiber and resin; the hub having a radially inner surface, a radially outer surface and a first edge;
a plurality of blade assemblies, each one of the plurality of blade assemblies comprising:
a blade;
a base having an outer portion, an inner portion, and a radially oriented leg connecting the outer portion and the inner portion, the base defining a slot opening between the outer portion and the inner portion and terminating at the radially oriented leg, wherein said slot receives the hub and the blade is mounted on the outer portion of the base;
a tang axially extending from the outer portion; and
wherein the plurality of blade assemblies are arranged circumferentially around the hub, each interlocking with an adjacent blade assembly and retained in position by the hub and a band overwrapping the respective tang of each of the plurality of blade assemblies.

US Pat. No. 10,273,818

GAS TURBINE ENGINE WITH COMPLIANT LAYER FOR TURBINE VANE ASSEMBLIES

Rolls-Royce North America...

1. A turbine vane assembly for use in a gas turbine engine, the turbine vane assembly comprisinga metallic outer endwall arranged around at least a portion of a central axis of the turbine vane assembly, the outer endwall having a first mating surface facing radially-inward toward the central axis,
a flow path component comprising ceramic material and having a second mating surface facing radially-outward away from the central axis and arranged to face the first mating surface, the second mating surface being spaced apart from the first mating surface to define an outer gap therebetween,
an outer compliant member located in the outer gap between the first and second mating surfaces, the outer compliant member being configured to compress to reduce a size of the outer gap in response to pressure loads acting on the outer endwall and the flow path component and to distribute the pressure loads between the first mating surface of the outer endwall and the second mating surface of the flow path component during use of the turbine vane assembly, and
a plurality of outer load pads located in the outer gap between the first and second mating surfaces, the plurality of outer load pads being rigid and each of the plurality of outer load pads having a fixed first load-pad thickness in a radial direction to limit relative movement between the first mating and second mating surfaces to maintain a minimum distance between the first and second mating surfaces, the minimum distance is greater than or equal to the fixed first load-pad thickness.

US Pat. No. 10,240,460

INSULATING COATING TO PERMIT HIGHER OPERATING TEMPERATURES

Rolls-Royce North America...

11. A method comprising:providing an airfoil-shaped metallic gas turbine engine component and a ceramic matrix composite component, the airfoil-shaped metallic gas turbine engine component positioned in an open interior of at least a portion of the ceramic matrix composite component;
identifying a thermal stress area of the ceramic matrix composite component upon which an adhesive coat of insulation can be applied to insulate the ceramic matrix composite component and discourage adverse thermal stresses; and
applying the coat of insulation to the thermal stress area by depositing the coat of insulation upon an inner surface of the ceramic matrix composite component, the coat of insulation spaced apart entirely from the airfoil-shaped metallic gas turbine engine component.

US Pat. No. 10,227,924

PARTICLE SEPARATOR

Rolls-Royce North America...

1. An air-inlet duct for a gas-turbine engine, the air-inlet duct comprisingan outer wall spaced apart from an engine rotation axis,
an inner wall located between the outer wall and the engine rotation axis, the inner wall and the outer wall defining an air-inlet passageway therebetween,
a splitter located between the outer wall and the inner wall and including an outer splitter surface cooperating with the outer wall to define a scavenge channel therebetween and an inner splitter surface cooperating with the inner wall to define an engine channel therebetween, and
a flow regulator configured to regulate a portion of an inlet flow including particles to cause a size and duration of a separated flow region formed along the outer wall and upstream of a scavenge inlet to the scavenge channel to be minimized so that the particles are collected in the scavenge channel and an amount of particles entering the engine channel are minimized,
wherein the flow regulator includes a series of flow control devices, each flow control device coupled to the outer wall to extend radially inward from the outer wall, and terminating at a free end exterior to the engine rotation axis and the inner wall, and each flow control device has a pair of parallel sides and is arranged to lie in spaced-apart circumferential relation to one another and located upstream of the separated flow region.

US Pat. No. 10,280,768

TURBINE BLISK INCLUDING CERAMIC MATRIX COMPOSITE BLADES AND METHODS OF MANUFACTURE

Rolls-Royce North America...

1. A turbine disk assembly adapted for use in a gas turbine engine, the assembly comprisinga disk comprising metallic materials and forming a radially-outermost surface, the radially-outermost surface being a continuous circumferential surface that extends around a central axis that defines an axial direction parallel thereto and a radial direction perpendicular thereto,
a blade comprising ceramic matrix composite materials, the blade including a root with a radially-innermost surface arranged radially outward of the radially-outermost surface formed by the disk and an airfoil extending radially outward from the root, and
a ring coupled to the disk and extending around the radially-outermost surface formed by the disk, the ring formed to include an opening that receives at least a portion of the root included in the blade to couple the blade to the ring while at least a portion of the airfoil included in the blade is arranged outside of the opening radially outward of the ring to interact with gasses radially outward of the ring, wherein the opening and the at least a portion of the root received by the opening are sized to allow for relative motion between the blade and at least one of the ring and the disk so that differing rates of thermal expansion between the blade and the disk can be accounted for.

US Pat. No. 10,226,797

METHOD OF CLEANING HIGH POWER CLUTCH

Rolls-Royce North America...

1. A method of cleaning a clutch used to transmit rotation from a driving shaft to a driven shaft, the method comprising:coupling a cleaning agent source to the clutch,
closing a clutch outlet so that cleaning agent dispensed into the clutch is retained in the clutch for a desired period,
dispensing foamed cleaning agent into the clutch so that the foamed cleaning agent contacts friction plates of the clutch,
rotating clutch components while the foamed cleaning agent is retained in the clutch and while the clutch outlet is closed, and
flushing the foamed cleaning agent along with other debris and grime out of the clutch so that the clutch is clean for further use in the transmission of rotation from a driving shaft to a driven shaft.

US Pat. No. 10,443,409

TURBINE BLADE WITH CERAMIC MATRIX COMPOSITE MATERIAL CONSTRUCTION

Rolls-Royce North America...

1. A turbine blade of ceramic matrix composite material construction adapted for use in a gas turbine engine, the turbine blade comprising:a root adapted to attach the turbine blade to a disk,
an airfoil shaped to interact with hot gasses moving through a gas path of the gas turbine engine and cause rotation of the turbine blade when the turbine blade is used in the gas turbine engine, and
a platform having an attachment side facing the root and a gas path side facing the airfoil, the platform arranged between the root and the airfoil and shaped to extend outwardly from the root and the airfoil in order to block gasses from the gas path migrating toward the root when the turbine blade is used in the gas turbine engine,
wherein the turbine blade includes core fiber-reinforcement plies that form part of the root and the airfoil without forming part of the platform, at least one gas path fiber-reinforcement ply that forms part of the airfoil and the platform, and attachment fiber-reinforcement plies that form part of the root and the platform, and wherein the number of attachment fiber-reinforcement plies is greater than the number of gas path fiber-reinforcement plies so that radial forces induced on the platform when the turbine blade is used in the gas turbine engine are transferred to the root without having to pass between fiber-reinforcement plies of the turbine blade.

US Pat. No. 10,370,997

TURBINE SHROUD HAVING CERAMIC MATRIX COMPOSITE SEAL SEGMENT

ROLLS-ROYCE CORPORATION, ...

1. A segmented turbine shroud for radially encasing a rotatable turbine in a gas turbine engine, the shroud comprising:a carrier comprising a forward generally planar flange and an aft generally planar flange, each of said flanges extending radially inward toward the turbine perpendicular to the axis of rotation of the turbine, said forward flange comprising a forward portion defining a pin-receiving forward carrier bore extending through the forward portion and said aft flange comprising an aft portion defining a pin-receiving aft carrier bore extending through said aft portion, wherein both the forward carrier bore and aft carrier bore have an axis parallel to the axis of rotation of the turbine;
a ceramic matrix composite (CMC) seal segment comprising a portion defining a pin-receiving seal segment bore; and
an elongated pin extending through said forward carrier bore, said seal segment bore, and said aft carrier bore,
wherein each of said carrier flange portions defining said pin-receiving forward carrier bore and said pin-receiving aft carrier bore include a respective member extending axially from said flange to thereby define the carrier bore having a length greater than the axial dimension of said flange, said member having a length sufficient to effect radial flexion between said member and said pin received within said carrier bore during operation of the gas turbine engine.

US Pat. No. 10,294,809

GAS TURBINE ENGINE WITH COMPLIANT LAYER FOR TURBINE SHROUD MOUNTS

Rolls-Royce North America...

6. A turbine shroud for use in a gas turbine engine, the turbine shroud comprisinga carrier arranged around a central axis of the turbine shroud, the carrier having a first mating surface,
a blade track segment having a second mating surface configured to be supported by the first mating surface, and
a load-distribution system configured to distribute loads transmitted between the second mating surface of the blade track segment and the first mating surface of the carrier, the load-distribution system including a compliant member engaged with the first and second mating surfaces and a plurality of rigid load pads,
wherein the compliant member has an uncompressed thickness, each of the plurality of rigid load pads have a second thickness that is less than the uncompressed thickness, and the plurality of rigid load pads are configured to engage the first and second mating surfaces in response to the compliant member being compressed to the second thickness,
wherein the compliant member includes an outer surface and an inner surface spaced apart from the outer surface, the complaint member is formed to define a plurality of receiver apertures extending between the outer surface and the inner surface of the complaint member, and each of the plurality of receiver apertures is shaped to receive a corresponding rigid load pad.

US Pat. No. 10,281,045

APPARATUS AND METHODS FOR SEALING COMPONENTS IN GAS TURBINE ENGINES

Rolls-Royce North America...

1. A turbine shroud apparatus adapted for use in a gas turbine engine to surround a turbine wheel assembly and block combustion products from passing over the tips of blades included in the turbine wheel assembly, the apparatus comprisinga first circumferential member comprising ceramic matrix composite materials, the first circumferential member shaped to extend partway about a central axis, an end portion of the first circumferential member defining a recess and including a first seal surface;
a second circumferential member comprising ceramic matrix composite materials, the second circumferential member shaped to extend partway about the central axis, an end portion of the second circumferential member configured to be at least partially disposed within the recess of the first circumferential member, the end portion of the second circumferential member including a second seal surface; and
a seal member configured to seal a gap between the first circumferential member and the second circumferential member, the seal member having a cylindrical shape that defines an seal axis arranged parallel to the central axis,
the first circumferential member and the second circumferential forming at least a portion of a circumferential assembly of a gas turbine when the end portion of the second circumferential member is at least partially disposed within the recess of the first circumferential member, the first seal surface configured to form a first seal with the seal member and the second seal surface configured to form a second seal with the seal member when the end portion of the second circumferential member is at least partially disposed within the recess of the first circumferential member.

US Pat. No. 10,267,160

METHODS OF CREATING FLUIDIC BARRIERS IN TURBINE ENGINES

ROLLS-ROYCE NORTH AMERICA...

1. A method of preventing pressure leakage from a core stream in a high bypass turbojet engine, comprising:dividing an ambient air stream into a bypass stream and a core stream with a upstream splitter;
compressing the bypass and core streams with a fan, said fan between the upstream splitter and a downstream splitter dividing the bypass and core streams downstream of the fan; wherein the core stream has a higher pressure than the bypass stream;
imparting a first momentum into an air stream proximate the fan in a region between the core and bypass streams and the upstream and downstream splitters to form a fluid barrier, wherein the first momentum of the air stream in the region is higher than a second momentum of the air stream adjacent the fluid barrier.

US Pat. No. 10,263,553

SYNCHRONOUS ELECTRICAL POWER DISTRIBUTION SYSTEM

Rolls-Royce North America...

1. A system comprising:a synchronous generator configured to supply polyphase electrical power to a plurality of loads;
a sensor configured to sense a voltage and a current of an output of the synchronous generator;
a controller configured to determine a desired power angle based on the voltage and the current received from the sensor to damp oscillations in a measured power angle between the voltage and the current; and
an exciter configured to excite the synchronous generator to control at least one of the voltage and the current of the output of the synchronous generator,
the controller configured to control the exciter based on the desired power angle to dynamically adjust the excitation of the synchronous generator to damp the oscillations in the measured power angle between the voltage and the current, wherein the oscillations in the measured power angle are at a first frequency, and the voltage and current are at a second frequency, the second frequency being greater than the first frequency.

US Pat. No. 10,260,523

FLUID COOLING SYSTEM INTEGRATED WITH OUTLET GUIDE VANE

Rolls-Royce North America...

1. A fan module for a gas turbine engine, the fan module comprisinga fan adapted to rotate about a central axis to pass air at least in part aftward along the central axis and around an engine core of the gas turbine engine,
a plurality of outlet guide vanes spaced aft of the fan along the central axis and configured to receive the air passed aftward along the central axis by the fan,
a fluid cooling system integral with at least one of the plurality of outlet guide vanes and configured to transfer heat from a fluid to the air from the fan to cool the fluid, the fluid cooling system including a plurality of cooling passages radially spaced from one another along the at least one outlet guide vane and extending parallel to one another along the central axis such that when fluid is conducted by the cooling passages during operation of the fan module, the cooling passages cooperate to control a pressure drop of the fluid as the fluid flows through the cooling passages to facilitate heat transfer from the fluid to the air from the fan during a first operating condition of the fan module, and
a back pressure regulator configured to bypass the cooling passages during a second operating condition of the fan module.

US Pat. No. 10,247,015

COOLED BLISK WITH DUAL WALL BLADES FOR GAS TURBINE ENGINE

Rolls-Royce Corporation, ...

1. A blisk for a gas turbine engine having a longitudinal axis, the blisk comprising a disk, a spar, a platform, and a shank portion integrally formed as a unit, the disk disposed about a longitudinal axis and having an upstream side and a downstream side, the spar extending radially outward from the platform relative to the longitudinal axis, the shank portion extending between the platform and the disk, wherein the spar includes a cooling air plenum defined therein disposed along an airfoil axis radially extended from the longitudinal axis, a plurality of standoffs extending away from an outer surface of the spar, wherein the blisk further comprises a cover panel bonded to an outer surface of the standoffs, wherein the standoffs are spaced from one another such that cooling passages are defined between the cover panel and the spar, the spar comprising one or more inlet ports defined therein in communication with the cooling air plenum and the cooling passages, the cover panel having one or more discharge ports defined therein in communication with the cooling passages, wherein a cooling feed channel is defined in the disk or shank portion that is in communication with the cooling air plenum, wherein the cooling feed channel is configured to receive cooling air upstream of the disk for delivery to the cooling air plenum.

US Pat. No. 10,393,021

PARTICLE SEPARATOR

Rolls-Royce North America...

1. An air-inlet duct for use with a gas turbine engine, the air-inlet duct comprisingan outer wall arranged circumferentially about an engine rotation axis of the air-inlet duct,
an inner wall located radially between the outer wall and the engine rotation axis, the inner wall and the outer wall cooperate to define an air-inlet passageway adapted to receive a mixture of air and particles suspended in the air, and
a splitter located radially between the outer wall and the inner wall and configured to separate the mixture of air and particles into a clean flow substantially free of particles and a dirty flow containing the particles, the dirty flow located adjacent the outer wall and the clean flow located radially between the dirty flow and the inner wall,
wherein the outer wall is formed to include a plurality of apertures arranged to extend radially through the outer wall to block a wall-normal vortex from forming in the air-inlet passage ahead of the splitter,
wherein the air-inlet duct further comprises a housing located radially outward of the outer wall, the housing cooperates with the outer wall to define a cavity, the plurality of apertures opens into the cavity to provide fluid communication between the air-inlet passageway and the cavity, and the plurality of apertures provides the only fluid inlets into and exits out of the cavity.

US Pat. No. 10,389,128

POWER CONTROL SYSTEM

Rolls-Royce Corporation, ...

1. A power control system, comprising:an energy storage device;
an engine-driven electrical machine;
a power converter electrically coupled to the energy storage device and the engine-driven electrical machine, wherein the power converter is configured to supply a total load power to an electrical load device; and
a controller coupled to the power converter, the controller receives characteristic data from at least one of the energy storage device, engine-driven electrical machine, and the electrical load device,wherein the characteristic data includes an anticipated load of the electrical loading device, and responsive to receiving the anticipated load, the controller is adapted to adjust first and second proportions of the total load power that are supplied by the energy storage device and the engine-driven electrical machine, respectively.

US Pat. No. 10,371,611

MATERIAL TESTING SYSTEM AND METHOD OF USE

Rolls-Royce North America...

1. A material testing system for loading test articles, the material testing system comprisinga load rod arranged to move along a load axis to provide a load force to a test article,
a rocker arm coupled to the load rod for axial movement therewith, the rocker arm including a rocker axis that intersects the load axis, a first end spaced apart from the rocker axis, and a second end spaced apart from the first end to locate the rocker axis between the first end and the second end of the rocker arm, and the rocker arm configured to pivot relative to the load rod about the rocker axis, and
a load distribution system configured to engage the rocker arm to transfer the load force from the load rod and rocker arm to the test article, the load distribution system including a first load applicator configured to engage the first end of the rocker arm and a second load applicator configured to engage the second end of the rocker arm, the first load applicator and the second load applicator configured to move axially relative to one another and relative to the load axis in response to engaging portions of the test article having different heights to cause the rocker arm to pivot about the rocker axis such that the rocker arm applies the force load equally between the first load applicator and the second load applicator,
wherein the first load applicator and the second load applicator are blocked from moving radially relative to the load axis.

US Pat. No. 10,301,955

SEAL ASSEMBLY FOR GAS TURBINE ENGINE COMPONENTS

Rolls-Royce North America...

1. A gas turbine engine assembly, the assembly comprisinga first component comprising ceramic matrix materials, the first component including a first panel formed to include a first chamfer surface and a first attachment feature that extends from the first panel to mount the first panel relative to other components within the gas turbine engine assembly,
a second component comprising ceramic matrix materials, the second component including a second panel formed to include a second chamfer surface and a second attachment feature that extends from the second panel to mount the second panel relative to other components within the gas turbine engine assembly, and
a seal assembly arranged in a channel formed by the first chamfer and the second chamfer when the first component is arranged in confronting relation to the second component, the seal assembly including a rod seal configured to block gasses from passing through the channel and a biaser seal configured to block gasses from passing through an interface between the first attachment feature and the second attachment feature, wherein the biaser seal is engaged with the rod seal and is configured to push the rod seal toward engagement with the first panel and the second panel,
wherein the biaser seal includes (i) an inner strip seal arranged in slots formed in the first attachment feature and the second attachment feature that engages the rod seal and (ii) a biaser arranged, at least in part, in the slots formed in the first attachment feature and the second attachment feature, and the biaser is configured to push the inner strip seal toward the rod seal.

US Pat. No. 10,263,552

ANTICIPATORY CONTROL USING OUTPUT SHAFT SPEED

ROLLS-ROYCE NORTH AMERICA...

1. A method of controlling engine speed droop, the method comprising:generating electricity from an electric generator powered by an output shaft of an engine;
anticipating, by a processor, an increase in a mechanical load on the engine; and
increasing a speed setpoint of the output shaft of the engine from a first value to a second value in response to anticipation of the increase in the mechanical load on the engine, wherein a control circuit attempts to cause a speed of the output shaft to match the speed setpoint based on a feedback loop, and wherein the speed of the output shaft increases prior to the increase in the mechanical load due to the increase of the speed setpoint from the first value to the second value.

US Pat. No. 10,507,930

AIRPLANE WITH ANGLED-MOUNTED TURBOPROP ENGINE

Rolls-Royce North America...

1. A turboprop propulsion system adapted for use in an airplane having an airframe, the turboprop propulsion system adapted for fixed connection to the airframe, the system comprising:a propeller mounted for rotation about a propeller axis,
a gas turbine engine mounted to establish an engine axis of rotation different from the propeller axis to form an offset angle therebetween, and
a gearbox coupled between the propeller and the gas turbine engine and configured to accommodate the offset angle between the propeller axis and the engine axis of rotation,
wherein the propeller axis and the engine axis of rotation are fixed with respect to each and are located within a single nacelle.

US Pat. No. 10,393,380

COMBUSTOR CASSETTE LINER MOUNTING ASSEMBLY

Rolls-Royce North America...

11. A gas turbine engine disposed about a longitudinal axis, comprising:a combustor to receive compressed air from a compressor, the combustor including a casing, an upstream dome coupled to each of an inner wall and an outer wall spaced from the casing, the outer wall comprising a plurality of combustor cassettes coupled to one another in an annular arrangement to define an upstream end and a downstream end of the outer wall;
a turbine disposed downstream of the combustor to receive combustion products from the combustor through a turbine nozzle, the turbine nozzle defined by an inner nozzle shroud and an outer nozzle shroud, the outer nozzle shroud having a nozzle upstream end; and
a ring mount and a mount stake, the ring mount coupled to both of the downstream end of the outer wall and the nozzle upstream end of the outer nozzle shroud, the mount stake coupled between the ring mount and the casing, wherein the downstream end of the outer wall includes a slot configured to receive an axial lip formed along the nozzle upstream end of the outer nozzle shroud and an axial flange of the ring mount extending upstream from the ring mount.

US Pat. No. 10,393,384

WAVE ROTOR WITH CANCELING RESONATOR

Rolls-Royce North America...

11. A wave rotor comprising a rotor drum mounted for rotation about a central axis of the wave rotor, the rotor drum formed to include a plurality of combustion rotor passages that extend along the central axis,an outlet end plate aligned axially with the rotor drum and formed to include an outlet port aperture extending axially through the outlet end plate along an arc around the central axis and aligned radially with the combustion rotor passages, the outlet end plate includes a leading edge wall and a trailing edge wall spaced apart circumferentially from the leading edge wall to define a portion of the outlet port aperture, and the combustion rotor passages are configured to rotate in a direction from the leading edge wall to the trailing edge wall, and
a first canceling resonator including a body and a neck that cooperate to define a cavity, wherein the neck is narrower than the body and is formed to include a mouth positioned adjacent to the leading edge wall, wherein the outlet end plate further includes a radial outer wall interconnecting the leading edge wall and the trailing edge wall and a radial inner wall radially spaced apart from the radial outer wall and interconnecting the leading edge wall and the trailing edge wall to form the port aperture, and the first canceling resonator extends radially away from the outlet port aperture, and
wherein the body and the neck are formed such that the only entrance into the cavity is through the mouth; and
wherein the first canceling resonator has a tuned frequency about equal to a frequency of the combustion rotor passages passing the port aperture when the rotor drum is rotated.

US Pat. No. 10,370,998

FLEXIBLY MOUNTED CERAMIC MATRIX COMPOSITE SEAL SEGMENTS

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 flange;
a ceramic matrix composite (CMC) seal segment comprising a portion defining a pin-receiving bore;
an elongated pin extending through said pin-receiving bore;
a bushing surrounding said elongated pin within said bore and having at least a portion extending axially beyond a first end of the bore; and
a flexible mounting member, said flexible mounting member being connected to said bushing at said portion extending axially beyond the first end of the bore and said carrier flange to thereby flexibly mount said CMC seal segment to said carrier.

US Pat. No. 10,371,008

TURBINE SHROUD

Rolls-Royce North America...

1. A turbine shroud comprisingan annular metallic carrier formed to include a plurality of apertures that extend in a radial direction through the annular metallic carrier,
a solid blade track formed as a continuous full hoop and including a ceramic annular runner and a plurality of pin receivers that extend inwardly in a radial direction partway into the ceramic annular runner from an outer radial surface toward an inner radial surface of the ceramic annular runner, and
a plurality of insert pins each arranged to extend through one of the plurality of apertures formed in the annular metallic carrier into a corresponding one of the plurality of pin receivers to locate the ceramic annular runner relative to the annular metallic carrier.

US Pat. No. 10,316,682

COMPOSITE KEYSTONED BLADE TRACK

Rolls-Royce North America...

1. A blade track for a gas turbine engine, the blade track comprisinga plurality of blade track segments comprising ceramic-matrix composite materials and shaped to extend part-way around a central axis, each blade track segment including opposing circumferential end faces and a radially outer surface extending between the end faces, and
an annular composite-lock structure positioned to engage the radially outer surfaces of the blade track segments, the composite-lock structure including ceramic-matrix materials and at least one reinforcement fiber of ceramic-containing material suspended in the ceramic-matrix materials of the annular composite-lock structure,
wherein the blade track segments are positioned circumferentially around the central axis to form a ring, the end faces of the blade track segments are engaged with one another, and the composite-lock structure continuously extends circumferentially around the entire ring to provide a radially-inward force toward the central axis against each of the plurality of blade track segments such that each blade track segment acts as a keystone to maintain a form of the ring.

US Pat. No. 10,309,257

TURBINE ASSEMBLY WITH LOAD PADS

Rolls-Royce North America...

2. A turbine shroud comprisinga metallic carrier arranged around a central axis, the metallic carrier formed to include a radially-inwardly opening blade track channel defined by a fore-retainer surface, an aft-retainer surface spaced apart axially from the fore-retainer surface, and an intermediate surface extending between the fore and aft-retainer surfaces,
a ceramic blade track segment including a runner and an attachment body extending radially outward away from the runner and positioned in the blade track channel to couple the ceramic blade track segment with the metallic carrier, the attachment body including a fore-attachment surface that faces the fore-retainer surface, an aft-attachment surface that faces the aft-retainer surface, and an outer surface that faces the intermediate surface, and
a plurality of load pads positioned in the blade track channel radially inward of the outer surface included in the attachment body between the fore-retainer surface and the fore-attachment surface and between the aft-retainer surface and the aft-attachment surface to transmit loads between the metallic carrier and the ceramic blade track segment at predetermined locations on the fore and aft-attachment surfaces while allowing growth of the metallic carrier and the ceramic blade track segment at different rates during use of the turbine shroud,
wherein the fore-retainer surface or the aft-retainer surface of the metallic carrier is formed to include a pad recess that extends into one of the fore-retainer surface or the aft-retainer surface, and the pad recess receives a portion of a load pad of the plurality of load pads.

US Pat. No. 10,273,903

ENGINE NACELLE

Rolls-Royce Corporation, ...

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,
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
at least one plasma actuator positioned to disrupt the air flow over the leading surface to reduce laminar separation of air flowing into the engine intake,
wherein the at least one plasma actuator is positioned at the leading surface and centered on the line of stagnation.

US Pat. No. 10,227,896

FLOW SEGREGATOR FOR INFRARED EXHAUST SUPPRESSOR

ROLLS-ROYCE CORPORATION, ...

1. A system for turbine exhaust treatment comprising:a turbine engine disposed in an engine compartment;
an exhaust region; and
a primary assembly disposed between the engine compartment and the exhaust region, the primary assembly having a center body, a circumferential member radially spaced from the center body, and a plurality of vanes extending radially outward from the center body and through the circumferential member, wherein the primary assembly is axially segregated radially outward from the circumferential member to form an axially forward portion adapted to receive air flow from the engine compartment and an axially aft portion adapted to receive air flow from a secondary air source;
wherein an axially-extending wall separates the engine compartment from the exhaust region, the axially-extending wall terminating in a flow segregator which contacts the primary assembly to segregate the engine compartment from the exhaust region.

US Pat. No. 10,393,383

VARIABLE PORT ASSEMBLIES FOR WAVE ROTORS

Rolls-Royce North America...

1. A wave rotor combustor comprisinga rotor drum mounted for rotation about a central axis of the wave rotor combustor, the rotor drum formed to include a plurality of rotor passages that extend along the central axis, the rotor drum extending between an inlet end adapted to receive a flow into the plurality of rotor passages and an outlet end adapted to emit the flow out of the plurality of rotor passages, the outlet end spaced apart from the inlet end along the central axis, and
an inlet assembly positioned adjacent to the inlet end of the rotor drum, the inlet assembly comprising:
a seal plate formed to include an inlet port extending axially through the seal plate along a first arc around the central axis to allow the flow to enter one or more of the rotor passages aligned with the inlet port,
a restrictor plate extending along a second arc around the central axis, the restrictor plate coupled to the seal plate to pivot about the central axis relative to the seal plate to cover selectively a portion of the inlet port and vary an area of the inlet port for the flow to pass through the seal plate, and
a restrictor-plate mover mounted to translate radially inwardly and outwardly relative to the central axis the restrictor-plate mover formed to define a mover surface that is elongated and curved along a length of the restrictor-plate mover, the mover surface abutting and engaging the restrictor plate during radial translation of the restrictor-plate mover relative to the restrictor plate to cause the restrictor plate to pivot relative to the seal plate around the central axis,
wherein the seal plate is fixed against rotation relative to the central axis.

US Pat. No. 10,370,994

PRESSURE ACTIVATED SEALS FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A gas turbine engine assembly adapted to be arranged between a high pressure zone and a low pressure zone within a gas turbine engine, the assembly comprisinga first component comprising metallic materials formed to include an axially extending seal-support slot,
a second component comprising ceramic-matrix composite materials arranged adjacent to the first component to form a radially extending gap therebetween, and
a pressure-activated seal comprising a sheet of material having a generally uniform thickness shaped to form a support portion that extends into the seal-support slot of the first component and a sealing portion arranged in the gap between the first component and the second component, wherein the sealing portion is shaped to be pushed by high pressure gas on one side of the gap into contact with the second component to resist flow through the gap,
wherein the sealing portion of the pressure-activated seal forms a radially-outwardly opening channel that receives high pressure gas from the high pressure zone and wherein the support portion of the pressure-activated seal forms an axially opening channel facing the gap that receives high pressure gas from the high pressure zone.

US Pat. No. 10,329,950

NOZZLE GUIDE VANE WITH COMPOSITE HEAT SHIELD

Rolls-Royce North America...

1. A nozzle guide vane for a gas turbine engine, the nozzle guide vane comprisinga metallic support structure including an inner endcap formed to include an inner attachment aperture and an outer endcap formed to include an outer attachment aperture, the outer endcap spaced from the inner endcap in a radial direction,
an airfoil including an aerodynamic feature shaped to redirect gasses moving through a gas path between the inner end cap and the outer endcap, an inner attachment feature that extends from the aerodynamic feature into the inner attachment aperture of the inner endcap, and an outer attachment feature that extends from the aerodynamic feature into the outer attachment aperture of the outer endcap, and
a ceramic-matrix composite heat shield system adapted to shield the metallic support structure from hot gasses moving through the gas path, the ceramic-matrix composite heat shield system including an inner heat shield arranged radially between the inner endcap and the gas path and an outer heat shield comprising ceramic-matrix composite materials arranged radially between the outer endcap and the gas path,
wherein the inner heat shield is formed to include an inner locator aperture, the outer heat shield is formed to include an outer locator aperture, and the entire inner locator aperture and the entire outer locator aperture are smaller than the aerodynamic feature when the aerodynamic feature is viewed in the radial direction to overlap the heat shields with the aerodynamic feature, a cooling gap between the outer endcap and the outer heat shield.

US Pat. No. 10,480,336

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 (i) 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, (ii) a ceramic tadpole gasket having a compressible head and a flat body extending from the compressible head, and (iii) a retainer ring configured to retain the ceramic tadpole gasket in place relative to the mount ring with a portion of the retainer ring folded over the mount ring to couple the retainer ring to the mount ring along with the flat body of the ceramic tadpole gasket, 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.

US Pat. No. 10,476,418

SYNCHRONOUS ELECTRIC POWER DISTRIBUTION STARTUP SYSTEM

ROLLS-ROYCE NORTH AMERICA...

1. A system comprising:a synchronous generator rotatable by a prime mover;
a controller configured to provide excitation signals to the synchronous generator to excite the synchronous generator to output electric power;
the controller further configured to provide an excitation signal at a first magnitude, during rotation of the synchronous generator at a speed less than rated speed, to supply a load bus with a first magnitude of current flow, the load bus electrically connected to a plurality of synchronous motor loads that are non-rotational in response to receipt of the first magnitude of current flow;
the controller further configured to determine a rotating rotor position of the synchronous generator during rotation of the synchronous generator at less than the rated speed and a rotor position of at least one of the synchronous motor loads;
the controller further configured to provide a pulse of the excitation signal at a second magnitude, the pulse provided coincident with a predetermined relative position of the rotating rotor position of the synchronous generator with respect to the rotor position of the at least one of the synchronous motor loads to supply the load bus with a second magnitude of current flow such that the plurality of synchronous motor loads rotate with the synchronous generator at the speed less than rated speed; and
the controller further configured to synchronously ramp rotational speed of the synchronous generator and the plurality of synchronous motor loads to the rated speed of the synchronous generator.

US Pat. No. 10,392,946

TURBINE BLADE WITH REINFORCED PLATFORM FOR COMPOSITE MATERIAL CONSTRUCTION

Rolls-Royce North America...

1. A turbine blade of ceramic matrix composite material construction adapted for use in a gas turbine engine, the turbine blade comprisinga root adapted to attach the turbine blade to a disk,
an airfoil shaped to interact with hot gasses moving through the gas path of a gas turbine engine and cause rotation of the turbine blade when the turbine blade is used in a gas turbine engine, and
a platform having an attachment side facing the root and a gas path side facing the airfoil, the platform arranged between the root and the airfoil and shaped to extend outwardly from the root and the airfoil in order to block migration of gasses from the gas path toward the root when the turbine blade is used in a gas turbine engine,
wherein the turbine blade includes at least one attachment fiber-reinforcement ply having a first section that forms part of the root and a second section that extends in a first direction away from the root to form part of the attachment side of the platform, at least one gas path fiber-reinforcement ply having a first section that forms part of the airfoil and a second section that extends in the first direction away from the airfoil to form part of the gas path side of the platform, and at least one through thickness reinforcement that extends in a second direction perpendicular to the first direction through the second section of the at least one attachment fiber-reinforcement ply and the second section of the at least one gas path fiber-reinforcement ply to secure the at least one attachment fiber-reinforcement ply to the at least one gas path fiber-reinforcement ply to facilitate resistance to loads induced on the platform when the turbine blade is used in a gas turbine engine.

US Pat. No. 10,378,439

GAS TURBINE ENGINE WITH VARIABLE SPEED TURBINES

Rolls-Royce North America...

1. A gas turbine engine, comprising:a fan configured to generate and transmit a bypass flow in a bypass duct;
a compressor in fluid communication with the fan;
a combustor in fluid communication with the compressor;
a first turbine in fluid communication with the combustor and operative to drive the compressor;
a second turbine in fluid communication with the first turbine and operative to drive the fan, wherein the second turbine includes at least a first stage and a second stage downstream of the first stage, both the first and second stages operative to drive the fan; and
a clutch configured to allow the first stage to rotate at a faster speed than the second stage,
further comprising a valve fluidly disposed between the first stage and the second stage, wherein the valve is configured to vent an output of the first stage, thereby reducing an amount of gas that is delivered to the second stage and increasing a speed of the first stage,
wherein the valve is configured as a sleeve valve,
wherein the clutch is configured to prevent the second stage from rotating faster than the first stage.

US Pat. No. 10,364,744

DEEP HEAT RECOVERY GAS TURBINE ENGINE

Rolls-Royce Corporation, ...

5. A system comprising:an air compressor configured to compress a flow of intake air;
a combustor configured to receive the compressed flow of intake air and provide exhaust output gas to produce thrust for an aircraft;
a fluid compressor configured to compress a working fluid, the working fluid being carbon dioxide;
an expander coupled to the fluid compressor and configured to receive and expand the compressed working fluid to generate mechanical energy and output a decompressed working fluid, wherein a temperature of the working fluid is reduced as the compressed working fluid expands;
a compressed air heat exchanger positioned ahead of an inlet of the combustor and configured to recuperatively transfer heat from the decompressed working fluid to the compressed flow of intake air;
an exhaust output gas heat exchanger positioned after an outlet of the combustor and configured to recuperatively transfer heat from the exhaust output gas of the combustor to the compressed working fluid; and
an ambient air heat exchanger positioned to receive an ambient air flow and transfer heat from the decompressed working fluid into the ambient air flow to lower a temperature of the decompressed working fluid to within a temperature range of a transcritical mode of the working fluid.

US Pat. No. 10,487,670

GAS TURBINE ENGINE COMPONENT INCLUDING A COMPLIANT LAYER

Rolls-Royce Corporation, ...

1. A method of producing a gas turbine engine component comprisingproviding a metallic wheel formed to define a dovetail-shaped groove, the metallic wheel arranged around an axis,
forming a blade from a ceramic matrix composite material, the blade including a blade portion and an engagement portion, the engagement portion sized to be received in the dovetail-shaped groove,
bonding at least one compliant layer to the engagement portion to cause a width of the engagement portion and the at least one compliant layer to be greater than a width of the dovetail-shaped groove, wherein the at least one compliant layer is applied using one of brazing, electroless deposition, spray coating, chemical vapor deposition, or plasma spraying, and
machining the at least one compliant layer to provide a mating surface of the at least one compliant layer for engagement with the metallic wheel, wherein the machining step removes at least a portion of the at least one compliant layer to cause the width of the engagement portion and the at least one compliant layer to be less than the width of the dovetail-shaped groove,
wherein the at least one compliant layer includes a first edge extending along a surface of the engagement portion and a second edge spaced apart axially relative to the axis from the first edge and extending along the surface of the engagement portion parallel to the first edge, and machining the at least one compliant layer includes chamfering the first and second edges to provide a first chamfer surface and a second chamfer surface,
wherein the at least one compliant layer further includes a bonding surface bonded with the engagement portion of the blade, the first chamfer surface is planar and extends between and contacts directly the bonding surface and the mating surface, and the second chamfer surface is planar and extends between and contacts directly the bonding surface and the mating surface,
wherein the at least one compliant layer is solid, the engagement portion of the blade is dovetail shaped, and the at least one compliant layer further includes a radial outer chamfer surface and a radial inner chamfer surface that is spaced apart radially from the radial outer chamfer surface relative to the axis, the radial outer chamfer surface is planar and extends between and contacts directly the bonding surface and the mating surface, the radial inner chamfer surface is planar and extends between and contacts directly the bonding surface and the mating surface.

US Pat. No. 10,480,408

ENERGY WEAPON SYSTEM HAVING A GAS TURBINE GENERATOR WITH IDLE ASSIST

Rolls-Royce North America...

1. A weapon system platform comprisinga high-energy beam unit configured to discharge high-energy beams,
a gas turbine engine configured to provide power for the high-energy beam unit, the gas turbine engine including a compressor, a combustor, and a turbine, the combustor adapted to combine air received from the compressor with fuel and to burn the fuel to supply high pressure gasses toward the turbine to rotate an output shaft of the gas turbine engine, the gas turbine engine adapted to operate between an idling output level and a main power generation output level higher than the idling output level,
a generator coupled to the output shaft of the gas turbine engine and adapted to generate electricity when driven by the gas turbine engine,
an energy storage unit coupled to the generator and configured to store the electricity generated by the generator,
a load applicator coupled to the gas turbine engine and configured to apply at least one of an electrical load on the energy storage unit and a mechanical load on the gas turbine engine, and
a system controller configured to selectively operate the load applicator to increase an output level of the gas turbine engine above the idling output level and below the main power generation output level to cause an increase in an exhaust temperature of the gas turbine engine when the exhaust temperature is below a threshold level.

US Pat. No. 10,480,519

HYBRID COMPRESSOR

Rolls-Royce North America...

1. A compressor for a gas turbine engine, the compressor comprisinga first compression stage mounted for rotation about a central axis, the first compression stage including a plurality of first-stage blades, and
a second compression stage mounted for rotation along the central axis aft of the first compression stage to receive air compressed by the first compression stage, the second compression stage including a plurality of second-stage blades, the second compression stage shaped to conduct air in a substantially radial direction away from the central axis and to discharge air in a substantially axial direction parallel to the central axis,
wherein an average exit radius of the plurality of first stage-blades is greater than an inlet tip radius of the plurality of first-stage blades.

US Pat. No. 10,378,477

NOZZLE FOR JET ENGINES

Rolls-Royce North America...

1. A gas turbine engine comprising:an engine core configured to discharge a first stream of pressurized core air that is passed through the engine core,
a fan coupled to the engine core to be driven by the engine core, the fan configured to discharge a second stream of pressurized bypass air and a third stream of pressurized bypass air that are passed around the engine core, and
an exhaust system coupled to the engine core, the exhaust system including a primary-stream nozzle arranged to adjust the first stream of pressurized core air discharged from the engine core and the second stream of pressurized bypass air discharged from the fan and including a third-stream nozzle arranged to adjust the third stream of pressurized bypass air discharged by the fan,
wherein the third stream of pressurized bypass air is adjustable via the third-stream nozzle based on adjustment of the first stream of pressurized core air and the second stream of pressurized bypass air by the primary-stream nozzle,
wherein the primary-stream nozzle includes at least one primary-stream duct wall, a plurality of forward primary-stream adjustment flaps mounted to pivot relative to the at least one primary-stream duct wall to adjust a primary-nozzle-throat defined between the plurality of forward primary-stream adjustment flaps, a plurality of aft primary-stream adjustment flaps mounted to move relative to the plurality of forward primary-stream adjustment flaps to adjust an outlet area of the primary-stream nozzle defined between the plurality of aft primary-stream adjustment flaps, one primary-stream-nozzle actuator coupled to at least one of the forward primary-stream adjustment flaps and configured to pivot the at least one of the forward primary-stream adjustment flaps relative to the primary-stream duct wall, and another primary-stream-nozzle actuator coupled to at least one of the aft primary-stream adjustment flaps and configured to move the at least one of the aft primary-stream adjustment flaps relative to the at least one of the forward primary-stream adjustment flaps.

US Pat. No. 10,329,943

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 variable-ratio unit coupled to the turbine and the compressor, the variable-ratio unit configured to transmit rotational power generated by the turbine about the engine axis to the axial compression stage to drive rotation of the axial compression stage at various speeds offset from a turbine speed,
wherein 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, the centrifugal compression stage is coupled to the turbine for common rotation therewith about the engine axis, and the variable-ratio unit is arranged forward of the axial compression stage about the engine axis.

US Pat. No. 10,494,935

BRAZED BLADE TRACK FOR A GAS TURBINE ENGINE

Rolls-Royce North America...

1. A full hoop blade track for a gas turbine engine, the blade track comprisinga first segment comprising ceramic-matrix composite materials and shaped to extend part-way around a central axis, the first segment having a first and a second circumferential end face,
a second segment comprising ceramic-matrix composite materials and shaped to extend part-way around the central axis, the second segment having a first and a second circumferential end face, and
a joint that couples the first segment to the second segment, the joint including a first biscuit that extends into the first segment along the second circumferential end face of the first segment and into the second segment along the first circumferential end face of the second segment to securely fix the second segment in place relative to the first segment.

US Pat. No. 10,480,328

FORWARD FLOWING SERPENTINE VANE

Rolls-Royce Corporation, ...

1. A vane adapted for use in a gas turbine engine, the vane comprisingan outer platform,
an inner platform, and
an airfoil that extends from the outer platform to the inner platform in a radial direction, the airfoil having a leading edge and a trailing edge and being formed to include a serpentine cooling passage with an inlet arranged through the outer platform at the trailing edge of the airfoil and an outlet arranged through the inner platform adjacent to the leading edge of the airfoil so that cooling air supplied to the inlet moves toward the leading edge of the airfoil during use in the gas turbine engine.

US Pat. No. 10,408,084

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 and an outer platform spaced apart from the inner platform to define a gas path therebetween, and
a ceramic-matrix composite airfoil that extends from the inner platform to the outer platform across the gas path, the ceramic-matrix composite airfoil formed to include a pressure side, a suction side, and a plurality of axially-aligned ribs spaced radially apart from one another between the inner platform and the outer platform that interconnect the pressure side with the suction side,
wherein the ceramic-matrix composite airfoil includes (i) a core comprising ceramic-containing reinforcements and having a first tube formed to include windows and a second tube formed to include windows, and (ii) an overwrap that wraps around the core to form the pressure side and the suction side of the airfoil,
wherein the windows of the second tube are aligned with the windows of the first tube so that the plurality of axially-aligned ribs are formed by material of the first tube and the second tube arranged between the windows of the first tube and the windows of the second tube, and
wherein the second tube is coupled to the first tube by a plurality of stitches in the plurality of axially-aligned ribs to fix the windows of the second tube in alignment with the windows of the first tube.

US Pat. No. 10,401,885

DC TO DC CONVERTER OUTPUT BUS VOLTAGE CONTROL SYSTEM

Rolls-Royce North America...

1. A system comprising:a DC to DC converter coupled with a load;
a power source bus coupled with an input of the DC to DC converter;
a capacitor coupled in parallel across an output of the DC to DC converter; and
a controller configured to dynamically adjust a bus voltage set point of a bus voltage on the output of the converter up or down to prepare for supply of the bus voltage and energy stored in the capacitor to an anticipated load event having a load step change that occurs in less than about 5 milliseconds and is greater than about 85% of a rated output of the DC to DC converter.

US Pat. No. 10,370,995

GAS TURBINE ENGINE VANE END DEVICES

Rolls-Royce North America...

6. An apparatus comprising:an airfoil member for a gas turbine engine, the airfoil member having a first end and a second end spaced apart radially from the first end relative to an axis, the airfoil member formed to include an interior that opens into the first end;
a wall arranged circumferentially around the first end of the airfoil member; and
a member that extends through the first end of the airfoil member into the interior, the member configured to provide fluid obstruction between the airfoil member and the wall,
wherein the airfoil member includes a pressure side and a suction side spaced apart from the pressure side, the member includes a piston that is located in the interior formed in the airfoil member and a head that is coupled to the piston, the head extends toward the pressure side and the suction side of the airfoil member, and the head has a convex-shaped outer surface when viewed circumferentially.

US Pat. No. 10,371,093

AIRCRAFT NOZZLE WITH A VARIABLE NOZZLE AREA OF A SECOND FLOW PATH

Rolls-Royce North America...

1. An apparatus comprising:an aircraft powerplant having a plurality of flow paths;
a first flow path of the plurality defined by a movable first member; and
a second flow path of the plurality radially outward of the first flow path and defined between a movable second member and a moveable third member, the movable third member is articulably coupled to a downstream end of the moveable first member, the movable third member structured to move with movement of the movable first member,
wherein the moveable first member comprises a plurality of movable first flaps and the moveable second member comprises a plurality of movable second flaps, and a first actuator is coupled with a first sync ring to move the plurality of movable first flaps and a second actuator is coupled with a second sync ring to move the plurality of movable second flaps independently relative to the moveable first member, and
wherein a variable nozzle area of the second flow path is configured to form a convergent-divergent geometry.

US Pat. No. 10,577,978

TURBINE SHROUD ASSEMBLY WITH ANTI-ROTATION FEATURES

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,
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, the attachment portion formed to include an anti-rotation notch, and
a mounting system configured to couple the blade track segment to the carrier segment, the mounting system including a brace that forms a bracket engaged with the attachment portion of the blade track segment to block radially inward movement of the blade track segment relative to the carrier segment, at least one shaft that extends from the bracket through the carrier to couple the blade track segment to the carrier segment, and an anti-rotation pin that extends into the anti-rotation notch to block circumferential movement of the blade track segment relative to the carrier segment,
wherein the attachment portion of the blade track segment has a L-shaped cross sectional shape including a substantially radially-extending portion and a substantially axially-extending portion and the bracket of the brace included in the mounting system directly engages the attachment portion to block radially inward movement of the blade track segment relative to the carrier segment,
wherein the shaft of the brace extends from the substantially radially-extending portion of the bracket and the bracket is formed to include at least one load pad spaced apart from the shaft that extends radially from the substantially axially-extending portion to directly engage the attachment portion of the blade track segment, and
wherein the bracket is formed to include a plurality of circumferentially spaced apart load pads that are each spaced apart from the substantially radially-extending portion and that each extend radially from the substantially axially-extending portion to directly engage the attachment portion of the blade track segment.

US Pat. No. 10,508,660

APPARATUS AND METHOD FOR POSITIONING A VARIABLE VANE

Rolls-Royce Corporation, ...

1. A variable vane positioning apparatus for controlling the position of a variable vane of a stator in a compressor of a turbine engine, comprising:a unison ring, wherein the vane is coupled with the unison ring via a vane arm;
a torque tube;
a cam having one end secured to a torque tube and a terminal end opposite the one end;
a cam follower secured to the unison ring, the cam follower having a face with a non-circular profile in contact with the terminal end of the cam;
wherein rotation of the torque tube causes movement of the cam, and movement of the cam moves the cam follower, thereby moving the unison ring and the variable vane coupled thereto.

US Pat. No. 10,491,145

GAS TURBINE GENERATOR SPEED DC TO DC CONVERTER CONTROL SYSTEM

Rolls-Royce North America...

1. A method comprising:controlling a gas turbine to maintain a constant rated speed;
monitoring the actual speed of the gas turbine;
driving a generator with the gas turbine at the constant rated speed to output electric power;
supplying a DC to DC converter with electric power output from the generator;
outputting electric power from the DC to DC converter to a load bus;
supplying the load bus with electric power from an energy storage device; and
controlling an output current of the DC to DC converter to supply the load on the load bus, the output current of the DC to DC converter controlled based on deviation of the actual speed of the gas turbine from the constant rated speed.

US Pat. No. 10,487,672

AIRFOIL FOR A GAS TURBINE ENGINE HAVING INSULATING MATERIALS

Rolls-Royce Corporation, ...

1. An airfoil adapted for use in a gas turbine engine, the airfoil comprisinga ceramic matrix composite skin that provides a pressure side and a suction side of the airfoil that extends from a leading edge to a trailing edge of the airfoil, the ceramic matrix composite skin shaped to define an internal cavity between the pressure side and the suction side of the airfoil sized to carry a cooling air flow,
a ceramic matrix composite reinforcement rib that extends across the internal cavity between the pressure side and the suction side of the airfoil to reinforce the ceramic matrix composite skin,
a metallic support spar that extends through the internal cavity, and
an insulating layer having a thermal conductivity lower than that of the ceramic matrix composite reinforcement rib that engages at least one side of the ceramic matrix composite reinforcement rib to thermally insulate the ceramic matrix composite reinforcement rib from temperatures in the internal cavity,
wherein the metallic support spar engages the insulating layer to block movement of the insulating layer in at least one direction, and wherein the metallic support spar includes a channel that receives the insulating layer to block movement of the insulating layer in at least two directions.

US Pat. No. 10,458,263

TURBINE SHROUD WITH SEALING FEATURES

Rolls-Royce North America...

17. A method of assembling a segmented turbine shroud that extends around a central axis, the method comprisingmounting a retainer formed to include flanges and attachment posts to a blade track segment formed to include a runner and hangers so that the flanges engage the hangers,
attaching the retainer to a carrier segment by coupling the attachment posts of the retainer to the carrier segment so that the retainer and the hangers are received in a radially inwardly-opening cavity formed by the carrier segment, and
positioning at least one compartment seal in a recess of the carrier segment to engage a radially outer side of the runner of the blade track segment to seal the radially inwardly-opening cavity.

US Pat. No. 10,451,826

SYSTEM FOR FIBER OPTIC COMMUNICATION CONNECTIONS

Rolls-Royce Corporation, ...

1. A system comprising:a housing;
a circuit board disposed in the housing;
a port included in the housing, the port forming an opening in the housing that is configured to receive a fiber optic cable;
circuitry included in the port, the circuitry positioned for electrical communication with printed circuitry included on the fiber optic cable received in the port;
a keyway included in the port to axially and rotationally align the fiber optic cable so that the printed circuitry aligns with the circuitry included in the port;
a locking stay included on the port, the locking stay configured to fixedly couple and hold the fiber optic cable in place in the port; and
an output connector configured for detachable connection and output of electrical signals.

US Pat. No. 10,443,493

EXHAUST MIXER FOR WAVE ROTOR ASSEMBLY

Rolls-Royce North America...

1. A wave rotor assembly comprisinga wave rotor combustor including an aft plate formed to include an exit port that extends circumferentially along an arc about a central axis of the wave rotor combustor and a rotor drum mounted for rotation relative to the aft plate about the central axis, the rotor drum formed to include a plurality of rotor passages that extend along the central axis and are arranged so that the rotor passages align with the exit port at predetermined intervals when the rotor drum rotates about the central axis to allow exhaust gasses in the rotor passages to flow through the exit port,
an exit duct coupled to the aft plate that defines a duct passage arranged to receive the exhaust gasses flowing through the exit port, the exit duct including a first duct wall and a second duct wall spaced apart circumferentially from the first duct wall to define a portion of the duct passage, and
an exhaust mixer located in the duct passage and arranged to circumferentially redirect relatively-high velocity exhaust gasses received by the exhaust mixer in the exit duct toward relatively-low velocity exhaust gasses received by a remainder of the exit duct so that the relatively-high velocity exhaust gasses are mixed with the relatively-low velocity exhaust gasses in the duct passage downstream of the exhaust mixer,
wherein the exhaust mixer includes a first radial sidewall that extends axially and circumferentially at an angle away from the first duct wall toward the second duct wall only partway into the duct passage relative to the central axis, the exhaust mixer being angled with respect to the remainder of the exit duct to redirect the relatively-high velocity exhaust gasses toward the relatively-low velocity exhaust gasses, the exhaust mixer having a converging area to accelerate the relatively-high velocity exhaust gasses.

US Pat. No. 10,443,539

HYBRID EXHAUST NOZZLE

Rolls-Royce North America...

11. An exhaust system for a gas turbine engine, the exhaust system comprisinga inner duct arranged around a central axis of the gas turbine engine, the inner duct defining an inner passageway arranged to receive a first stream of pressurized air discharged from an engine core of the gas turbine engine and a second stream of pressurized air discharged from a fan of the gas turbine engine and passed around the engine core,
an outer duct arranged radially outward of the inner duct around the central axis of the gas turbine engine, the outer duct cooperating with the inner duct to define an outer passageway arranged to receive a third stream of pressurized air discharged from the fan of the gas turbine engine and passed around the engine core,
a plurality of aft adjustment flaps mounted to pivot relative to the inner duct, each of the aft adjustment flaps having an inner portion extending radially inward from the inner duct toward the central axis and an outer portion extending radially outward from the inner duct away from the central axis, and
a plurality of forward adjustment flaps positioned forward of the aft adjustment flaps along the central axis, each of the forward adjustment flaps having an aft end pivotally coupled to the outer portion of a corresponding aft adjustment flap and a forward end mounted to the inner duct to slide relative to the inner duct.

US Pat. No. 10,443,856

WAVE ROTORS WITH TEMPERATURE CONTROL FEATURES

Rolls-Royce North America...

16. A method of regulating a temperature distribution of a wave rotor combustor, the method comprisingproviding a rotating rotor drum assembly arranged to rotate about a central axis of the wave rotor combustor, the rotor drum assembly including a forward end, an aft end spaced apart from the forward end along the central axis, an outer tube, with radial flow exhaust openings, that extends along the central axis between the forward and aft ends, a drum sleeve surrounding the outer tube to form an outer air compartment, an inner tube positioned radially between the outer tube and the central axis, and a plurality of webs extending radially between and interconnecting the outer and inner tubes to form a plurality of rotor passages, each web formed to include a plurality of fluid flow passages extending radially entirely through each web and adapted to receive conditioned air,
supplying conditioned air to the plurality of fluid flow passages near the aft end of the rotor drum assembly to cool the aft end of the rotor drum assembly and provide heated conditioned air, and
collecting in the outer air compartment the heated conditioned air exiting the fluid flow passage which passes through the radial flow exhaust openings near the aft end of the rotor drum assembly and supplying the heated collected conditioned air to the plurality of fluid flow passages in the webs near the forward end of the rotor drum assembly,
ejecting an additional supply of heated conditioned air from additional cooling passages that are used to cool an aft trailing edge of the plurality of webs and are ejected in a generally axial direction, relative to the central axis.

US Pat. No. 10,429,154

ENERGY WEAPON HAVING A FAST START TURBINE FOR A HIGH POWER GENERATOR

Rolls-Royce North America...

1. A weapon system platform comprising:a high-energy beam unit configured to discharge high-energy beams,
a gas turbine engine configured to provide power for the high-energy beam unit, the gas turbine engine including a first shaft coupled to a compressor and a high pressure turbine rotor, a second shaft concentric with and independently rotatable relative to the first shaft and coupled to a low pressure turbine rotor, a starter adapted to rotate the first shaft, and a combustor adapted to combine air received from the compressor with fuel and to burn the fuel to supply high pressure gasses toward the high pressure turbine rotor and low pressure turbine rotor to rotate the first and second shafts,
a generator coupled to the second shaft of the gas turbine engine and adapted to generate electricity when driven by the gas turbine engine,
an energy storage unit coupled to the generator and configured to store the electricity generated by the generator, and
a generator control system configured to selectively operate the starter and to selectively deliver fuel to the combustor such that the first shaft is continuously rotated by at least one of the starter and high pressure gasses from the combustor and the second shaft is selectively rotated by high pressure gasses from the combustor,
wherein fuel is selectively delivered to the combustor when a power demand signal of the high-energy beam unit is received by the generator control system, and
wherein fuel is selectively cut from the combustor when no power demand signal of the high-energy beam unit is received by the generator control system while the first shaft continues to be rotated by the starter in order to maintain a rapid-start mode of the weapon system platform.

US Pat. No. 10,428,674

GAS TURBINE ENGINE FEATURES FOR TIP CLEARANCE INSPECTION

Rolls-Royce North America...

1. A turbine assembly comprisinga turbine wheel assembly including a disk and a plurality of blades that extend outwardly from the disk in a radial direction away from an axis, and
a turbine shroud that extends around the blades of the turbine wheel assembly to block gases from passing over the blades during operation of the turbine assembly, the turbine shroud including a plurality of blade track segments arranged circumferentially adjacent to one another about the axis to form a ring, each blade track segment having a runner that forms a primary track surface facing the axis and spaced from the axis in the radial direction,
wherein (i) at least one of the plurality of blade track segments includes a first set of rub depth indicators spaced from one another and each having a first depth measured from the primary track surface and a second set of rub depth indicators spaced from one another and each having a second depth measured from the primary track surface, (ii) the first set of rub depth indicators and the second set of rub depth indicators are configured such that approximate rub depths of the turbine wheel assembly into the turbine shroud caused by turbine wheel assembly rotation within the turbine shroud during operation of the turbine assembly may be determined based on visual observation of the first set of rub depth indicators and the second set of rub depth indicators, (iii) the first set of rub depth indicators are arranged along a first pathway in a first direction such that the first depths successively increase as the first set of rub depth indicators are located adjacent to one another in the first direction, and (iv) the second set of rub depth indicators are arranged in the first direction along a second pathway such that the second depths successively decrease as the second set of rub depth indicators are located adjacent to one another in the first direction.

US Pat. No. 10,415,403

COOLED BLISK FOR GAS TURBINE ENGINE

Rolls-Royce North America...

1. A blisk for a gas turbine engine, the blisk comprising:a disk disposed about an axis, the disk having an upstream side and a downstream side;
a plurality of blades, each blade comprising a platform, a cooled airfoil extending radially outward from the platform relative to the axis, and a shank portion extending between the platform and the disk, the disk having a cooling cavity defined therein between the shank portions of the blades that are adjacent to one another, the cooling cavity in fluid communication with the cooled airfoil, wherein a radial lip extends inward relative to the axis from the downstream side of the disk, the cooling cavity configured to receive cooling air provided upstream of the blisk; and
a coverplate disposed over the cooling cavity, the coverplate including an inner tab extending radially inward from an inner edge of the coverplate, wherein the inner tab is coupled to the radial lip to securely couple the coverplate to the disk.

US Pat. No. 10,415,415

TURBINE SHROUD WITH FORWARD CASE AND FULL HOOP BLADE TRACK

Rolls-Royce North America...

1. A turbine shroud assembly for a gas turbine engine, the turbine shroud assembly comprisingan annular turbine case arranged around a central axis of the turbine shroud assembly,
a full hoop blade track comprising ceramic matrix composite materials located radially between the turbine case and the central axis, the blade track having a leading edge and a trailing edge axially spaced apart from the leading edge, and
a forward case coupled to the turbine case and arranged to extend radially inwardly away from the turbine case toward the central axis and interlocked with the leading edge of the blade track to block circumferential and axial movement of the blade track relative to the turbine case while allowing radial movement of the blade track relative to the turbine case,
further comprising a carrier located radially between the turbine case and the blade track and a hollow cross-key pin that extends through the turbine case into the carrier, the carrier is arranged to define an inwardly facing thermal management chamber, and the hollow cross-key pin is configured to direct airflow through the turbine case and the carrier into the thermal management chamber toward the blade track.

US Pat. No. 10,612,468

GAS TURBINE ENGINE WITH THERMOELECTRIC INTERCOOLER

Rolls-Royce North America...

1. A gas turbine engine for generating drive from combustion of fuel, comprisinga compressor comprising a plurality of rotating stages each adapted to generate compressed air,
a cooling source adapted to provide coolant to the compressor of the gas turbine engine, and
a thermoelectric intercooler located axially between rotating stages of the plurality of rotating stages of the compressor along a central engine axis of the gas turbine engine, the thermoelectric intercooler comprising a compressed air passageway fluidly coupled to the compressor to pass the compressed air of the compressor between first and second rotating stages of the rotating stages of the compressor, a coolant passageway fluidly coupled to the cooling source to pass coolant of the cooling source therethrough, and a thermoelectric section configured in thermal communication with each of the compressed air passageway and the coolant passageway,
wherein the compressed air passageway comprises a plurality of compressed air conduits, and the coolant passageway comprises a plurality of coolant conduits each defining a coolant flow path that extends radially in communication with a turnaround passage formed between two adjacent coolant conduits of the plurality of coolant conduits;
wherein a compressed air flow path through the plurality of compressed air conduits is transverse to the coolant flow path through the plurality of coolant conduits; and
wherein the thermoelectric section is disposed between one of the plurality of compressed air conduits and one of the plurality of coolant conduits.

US Pat. No. 10,590,993

BEARING RACE COOLING

Rolls-Royce North America...

1. A cooling architecture comprising:a longitudinally extending radially inner shaft comprising an inner race, the inner race defining an inner circumferential chamber configured to carry an inner working fluid, wherein the inner shaft is at least partially hollow;
a fluid line extending through the hollow inner shaft and configured to deliver inner working fluid to the inner circumferential chamber;
a radially outer support comprising an outer race that defines an outer circumferential chamber configured to carry an outer working fluid;
a bearing assembly comprising a plurality of roller bearings disposed radially between the inner race and the outer race, the bearing assembly configured to radially align the inner shaft with respect to the outer support;
a plurality of co-circumferential inner inlets configured to deliver inner working fluid to the inner circumferential chamber; and
a processing system configured to independently adjust a flow rate of inner working fluid through each of the inner inlets such that a flow rate of inner working fluid through a first inner inlet is lower than a flow rate of inner working fluid through a second co-circumferential inner inlet.

US Pat. No. 10,563,528

TURBINE VANE WITH CERAMIC MATRIX COMPOSITE AIRFOIL

Rolls-Royce North America...

1. A turbine engine assembly, the assembly comprisinga turbine case, and
a turbine vane adapted for use in a gas turbine engine, the turbine vane comprising
a ceramic matrix composite component formed to include an airfoil shaped to interact with hot gasses moving along a primary gas path defined by the gas turbine and an attachment feature that extends from a radial end of the airfoil,
an inner end wall coupled to the ceramic matrix composite component and shaped to define an inner boundary of the primary gas path near an inner radial end of the airfoil, and
a metallic outer end wall coupled to the ceramic matrix composite component and shaped to define a boundary of the primary gas path near an outer radial end of the airfoil, the metallic outer end wall including a first part and a second part coupled to the first part,
wherein the first part and the second part are each shaped to include a gas path panel that defines part of the boundary of the primary gas path and an attachment receiver that forms a pocket into which a portion of the attachment feature included in the ceramic matrix composite component extends,
wherein the first part and the second part of the metallic outer end wall each include at least one case hanger that extends in a radial direction from the gas path panel away from the primary gas path and that is configured to be coupled to the turbine case, and
wherein the pockets formed by the attachment receivers of the first and the second part of the metallic end wall cooperate to define an attachment-feature receiving space shaped to block removal of the attachment feature from the pockets while the first part and the second part of the metallic outer end wall are coupled to one another so that the ceramic matrix composite component is mounted to the metallic outer end wall.

US Pat. No. 10,563,670

VANE ACTUATION SYSTEM FOR A GAS TURBINE ENGINE

Rolls-Royce Corporation, ...

1. A compressor comprisinga case extending circumferentially about a central axis,
a plurality of vanes arranged circumferentially adjacent to one another about the central axis inside the case, each one of the vanes configured for rotation about a vane axis generally perpendicular to the central axis, and
a vane actuation system including an actuator and an actuation ring coupled to one of the vanes and arranged radially inward of a portion of the case relative to the central axis, the actuator configured to drive movement of the actuation ring to cause rotation of the one of the vanes about the vane axis,
wherein the actuation ring is formed to include a plurality of circumferential channels opening radially outward toward the case and sized to receive a plurality of rollers configured to engage the case to permit movement of the actuation ring relative to the case.

US Pat. No. 10,530,163

MICRO GRID CONTROL SYSTEM

Rolls-Royce Corporation, ...

1. A micro grid control system comprising:a plurality of sources of electrical energy comprising a generator driven by a gas turbine engine, a renewable energy source and a stored energy source;
each of the sources of electrical energy comprising a source controller configured to optimize generation of electrical energy according to a source objective independently associated with each of the respective sources; and
a central power controller configured to receive operational parameters from each of the plurality of sources and selectively and dynamically provide a power demand signal and a mode signal to each of the sources independently, the central power controller further configured to provide electrical energy to a micro grid from any two or more of the sources according to a system objective indicated by the mode signal, the two or more sources operable in accordance with their respective source objective, after their respective source objective is independently adjusted by a feed forward weighting, to supply electrical power to the central power controller for allocation to the micro grid, the feed forward weighting determined based on the power demand signal and the mode signal.

US Pat. No. 10,502,131

WAVE ROTOR WITH PISTON ASSEMBLY

Rolls-Royce North America...

13. A method of operating a wave rotor assembly, the method comprisingproviding a rotor drum having a forward end and an aft end, an outlet plate offset from the aft end of the rotor drum by a clearance gap, and a piston assembly arranged to engage the outlet plate,
introducing gasses into the rotor drum to cause a first force to be applied onto the outlet plate at the aft end of the rotor drum, and
applying a pressure from the gasses in the rotor drum to the piston assembly to cause the piston assembly to apply a second force onto the outlet plate to counteract the first force,
wherein the outlet plate is formed to include an outlet port spaced apart circumferentially from the piston assembly and a pressurizing passage aligned circumferentially with the piston assembly to allow a portion of the gasses to communicate with the piston assembly and wherein a piston included in the piston assembly is formed to include a piston passage that is aligned with the pressurizing passage formed in the outlet plate.

US Pat. No. 10,494,116

HEAT SHIELD FOR SIGNATURE SUPPRESSION SYSTEM

Rolls-Royce North America...

1. An aircraft comprising a gas turbine engine including an exhaust system,a heat suppression system fluidly connected with the exhaust system and adapted to inhibit line of sight therein, the heat suppression system including:
an outer skin defining a cavity and including at least one mount,
an exhaust conduit arranged within the cavity of the outer casing having an exhaust passageway defined therethrough for receiving exhaust of the gas turbine engine,
a vane diffuser arranged within the exhaust passageway of the exhaust conduit, and
a shield system arranged within the cavity and secured with the outer skin, the shield system including a heat shield and an insulation layer, the heat shield arranged between the outer skin and the exhaust conduit, the insulation layer disposed between the heat shield and the outer skin,
wherein the heat shield is supported to float relative to the outer skin on the at least one mount,
wherein the at least one mount defines slanted surfaces for engagement with the heat shield,
wherein the heat shield includes a forward sheet and an aft sheet, the forward sheet secured to a first surface of the slanted surfaces and the aft sheet secured to a second surface of the slanted surfaces, wherein the first and second slanted surfaces have different pitch.

US Pat. No. 10,498,275

SYNCHRONOUS ELECTRICAL POWER DISTRIBUTION EXCITATION CONTROL SYSTEM

Rolls-Royce North America...

1. A system comprising:an exciter configured to output an excitation signal to control a magnetic field of a synchronous generator; and
a controller configured to control the exciter with an exciter voltage to control the magnetic field and electric power output of the synchronous generator over a range of rotational speed of the exciter;
a waveform of the exciter voltage selectively including at least one of an AC component or a DC component, and the controller configured to transition a level of the AC component lower and transition a level of the DC component higher based on an increase in rotational speed of the exciter within the range of rotational speed and electric power output of the synchronous generator, wherein the synchronous generator is electrically coupled with a load at a zero speed condition of the synchronous generator, and the controller is configured to transition the level of the AC component lower and transition the level of the DC component higher during the increase in rotational speed of the exciter within the range of rotational speed of the synchronous generator between the zero speed condition and rated speed of the synchronous generator while the electric power output is supplied to the load.

US Pat. No. 10,480,337

TURBINE SHROUD ASSEMBLY WITH MULTI-PIECE SEALS

Rolls-Royce North America...

1. A turbine shroud assembly adapted for use in a gas turbine engine, the assembly comprisinga carrier comprising metallic materials,
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 that mounts to the carrier to couple the blade track segment to the carrier, and
a multi-piece seal including components received in a first radially-inwardly opening channel formed in the carrier, the multi-piece seal engages the runner of the blade track segment to resist the flow of gasses through an interface between the carrier and the runner of the blade track segment,
wherein the multi-piece seal includes a forward wire arranged along a forward side of the first radially-inwardly opening channel, an aft wire arranged along an aft side of the first radially-inwardly opening channel, and a rope seal arranged in the first radially-inwardly opening channel between the forward wire seal and the aft wire seal, and wherein the rope seal is sized to engage the forward wire seal and the aft wire seal such that the forward wire seal and the aft wire seal are pushed away from one another and radially inward into engagement with the carrier and the runner of the blade track segment,
wherein the first radially-inwardly opening channel is arranged along a forward side of the blade track segment axially forward of the attachment portion of the blade track segment, and
wherein the carrier is formed to include a second radially-inwardly opening channel that opens to face the runner of the blade track segment, wherein the multi-piece seal includes (i) a second forward wire arranged along a forward side of the second radially-inwardly opening channel, (ii) a second aft wire arranged along an aft side of the second radially-inwardly opening channel, and (iii) a second rope seal arranged in the second radially-inwardly opening channel between the second forward wire seal and the second aft wire seal, and wherein the second rope seal sized to engage the second forward wire seal and the second aft wire seal such that the second forward wire seal and the second aft wire seal are pushed away from one another and radially inward into engagement with the carrier along a side of the second radially-inwardly opening channel and the runner of the blade track segment.

US Pat. No. 10,476,417

GAS TURBINE GENERATOR TORQUE DC TO DC CONVERTER CONTROL SYSTEM

Rolls-Royce North America...

1. A method comprising:controlling a gas turbine to rotate at a constant rated speed and a variable torque to drive a generator in accordance with a load being supplied by the generator;
outputting electric power with the generator to a power source bus;
energizing a DC to DC converter coupled to the power source bus as at least part of the load being supplied by the generator;
controlling a level of a DC output power output by the DC to DC converter on a load bus;
supplying the load bus with DC power supplied from an energy storage device;
receiving a transient load control signal indicative of an anticipated future occurrence of a step change load event on the load bus;
receiving the step change load event on the load bus after receipt of the transient load control signal;
adjusting the level of the DC output power of the DC to DC converter in response to the step change load event; and
applying a constraint on a rate of change of the level of the DC output power of the DC to DC converter, the constraint applied based on the transient load signal to maintain the variable torque of the generator below a predetermined threshold as a load ramp rate of the DC to DC converter changes on the power source bus.

US Pat. No. 10,458,429

IMPELLER SHROUD WITH SLIDABLE COUPLING FOR CLEARANCE CONTROL IN A CENTRIFUGAL COMPRESSOR

ROLLS-ROYCE CORPORATION, ...

1. A compressor shroud assembly in a turbine engine, the compressor shroud assembly comprising:a static compressor casing;
an actuator mounted to said casing; and
an impeller shroud for encasing a rotatable centrifugal compressor, said impeller shroud coupled at a forward end to said casing by a slidable coupling that maintains an air boundary during the full range of axial movement of said impeller shroud, said impeller shroud mounted proximate an aft end to said actuator, said impeller shroud moving relative to the rotatable centrifugal compressor in an axial direction while maintaining a radial alignment when said actuator is actuated.

US Pat. No. 10,458,268

TURBINE SHROUD WITH SEALED BOX SEGMENTS

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 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 the attachment-receiving space formed by the carrier segment, and
at least one strip seal configured to resist movement of gases after into the attachment-receiving space, the at least one strip seal arranged to extend across a gap formed radially between the carrier segment and the runner into a radially-outwardly opening runner slot formed in a radially-outwardly facing surface of the runner,
wherein a first strip seal is a one-piece component and extends at least most of the way around the attachment portion of the blade track segment, and
wherein the first strip seal is formed to include a gap as it extends around the attachment portion of the blade track segment sized to accommodate thermal expansion induced during use of the turbine shroud segment.

US Pat. No. 10,443,419

SEAL FOR A GAS TURBINE ENGINE ASSEMBLY

Rolls-Royce North America...

1. A seal for a gas turbine engine comprising:a first component having a face and a second component having a face abutting the face of the first component, the first and second components separating a region of high pressure from a region of low pressure, the faces of the first and second components each including a discontinuity configured such that when the faces are placed in a confronting relationship, the discontinuities form a cavity, and
a seal member positioned in the cavity, the seal member cooperating with the cavity such that high pressure gas in the region of high pressure that traverses the interface between the confronted faces urges the seal member against a portion of the discontinuities to seal the interface between the seal member and those portions of the faces engaged by the seal member,
wherein the discontinuity in the face of the first component and the discontinuity in the face of the second component form an angle with an apex of the angle positioned nearer the region of low pressure as compared the region of high pressure, the seal member positioned such that the high pressure that traverses the interface between the confronted faces urges the seal member into contact with the faces defining the angle to control the flow of gas from the region of high pressure to the region of low pressure.

US Pat. No. 10,443,420

SEAL ASSEMBLY FOR GAS TURBINE ENGINE COMPONENTS

Rolls-Royce North America...

1. A gas turbine engine assembly comprisinga first component comprising ceramic matrix materials, the first component including a first panel formed to include a first chamfer surface and a first attachment feature that extends from the first panel to mount the first panel relative to other components within the gas turbine engine assembly,
a second component comprising ceramic matrix materials, the second component including a second panel formed to include a second chamfer surface and a second attachment feature that extends from the second panel to mount the second panel relative to other components within the gas turbine engine assembly, the second component being circumferentially spaced apart from the first component about a center axis of the gas turbine engine assembly to form a gap therebetween, and
a seal assembly arranged between the first component and the second component to block gasses from passing through the gap, the seal assembly including a first strip seal that extends circumferentially into the first attachment feature and the second attachment feature to block gasses from passing between the first attachment feature and the second attachment feature and a rod seal arranged in a channel formed by the first chamfer surface and the second chamfer surface to block gasses from passing between the first panel and the second panel and the first strip seal is interlocked with the rod seal for radial movement therewith to block gasses from passing between the rod seal and the first strip seal.

US Pat. No. 10,428,676

TIP CLEARANCE CONTROL WITH VARIABLE SPEED BLOWER

ROLLS-ROYCE CORPORATION, ...

1. A system, comprising:a distribution manifold positioned along an outer surface of an engine case of a gas turbine engine, the distribution manifold including a plurality of outlets defined on the distribution manifold to direct a thermal fluid received by the distribution manifold onto an outer surface of the engine case of the gas turbine engine; and
a variable blower configured to blow the thermal fluid into the distribution manifold at a flow rate controlled by the variable blower, wherein the variable blower comprises a plurality of blower blades and the variable blower is configured to adjust a rotational speed of the plurality of blower blades to vary the flow rate of the thermal fluid, wherein the flow rate through the variable blower is adjustable over a range of non-zero target flow rates, wherein the flow rate adjusted by the variable blower is not adjusted by a physical valve, wherein the thermal fluid from the variable blower is applied to the plurality of outlets.

US Pat. No. 10,622,813

LOAD ALIGNMENT ASSISTANCE DURING STARTUP OF SYNCHRONOUS GRID

ROLLS-ROYCE NORTH AMERICA...

1. An apparatus for load alignment assistance, 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 during a startup of the synchronous AC motor, wherein the synchronous AC motor receives polyphase AC power through the n-phase supply line from a synchronous AC grid during the startup, 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 is selected to cause a rotor of the synchronous AC motor to align with a rotor of a generator that powers the synchronous AC grid.

US Pat. No. 10,612,385

TURBINE BLADE WITH HEAT SHIELD

Rolls-Royce Corporation, ...

1. An airfoil adapted for use in a gas turbine engine, the airfoil comprisinga spar comprising metallic materials and formed to include a core body and a tail, the tail shaped to form an aft portion and a trailing edge of the airfoil, the aft portion defining an aft portion of a pressure side of the airfoil and an aft portion of a suction side of the airfoil, the tail defining an outermost surface of the airfoil adapted to be exposed to a gas path on the pressure side of the airfoil and the suction side of the airfoil, and
a single-piece heat shield comprising ceramic matrix composite materials and shaped to extend around the core body to form a leading edge of the airfoil, a forward portion of the pressure side of the airfoil, and a forward portion of the suction side of the airfoil,
wherein the forward portion of the pressure side of the airfoil is greater than half of the pressure side of the airfoil and the forward portion of the suction side of the airfoil is greater than half of the suction side of the airfoil,
wherein the core body includes a forward spar portion and an aft spar portion that is spaced-apart from the forward spar portion to define a heat shield channel therebetween, the tail is integrally formed with the aft spar portion, the single-piece heat shield is formed to include a suction side segment, a pressure side segment, a leading edge segment, and a web, the suction side segment and the pressure side segment are integrally formed with the web and extend in an aft direction from the web toward the tail such that the suction side segment and the pressure side segment have free ends that are cantilevered from the web, the leading edge segment extends from the pressure side segment to the suction side segment, and the web extends through the heat shield channel between the forward spar portion and the aft spar portion to interconnect the suction side segment with the pressure side segment.

US Pat. No. 10,612,399

TURBINE VANE ASSEMBLY WITH CERAMIC MATRIX COMPOSITE COMPONENTS

Rolls-Royce North America...

1. A turbine vane assembly adapted for use in a gas turbine engine, the turbine vane assembly comprisingan airfoil comprising ceramic matrix composite materials and shaped to interact with hot gases moving axially along a primary gas path of the gas turbine engine relative to an axis, the airfoil formed to include a radial-inner wall and a sidewall that extends radially outward and away from a perimeter of the radial-inner wall to define an interior region of the airfoil, and the radial-inner wall formed to define an airfoil passageway that extends radially through the radial-inner wall and opens into the interior region,
an endwall comprising ceramic matrix composite materials and shaped to define a boundary of the primary gas path near a radial end of the airfoil and the endwall formed to define an end-wall passageway that extends radially through the endwall, and
a spar comprising metallic materials and located in the interior region of the airfoil to carry loads that act on the airfoil, the spar including a spar body that engages the radial-inner wall of the airfoil so that an interface between the spar and the airfoil is located radially inward toward the boundary of the primary gas path, a spar tail that extends radially inward away from the spar body through the airfoil passageway and the end-wall passageway, and a retainer coupled to the spar tail to block movement of the radial-inner wall and the endwall away from the spar body.

US Pat. No. 10,612,402

METHOD OF ASSEMBLY OF BI-CAST TURBINE VANE

Rolls-Royce North America...

1. A method of assembling a gas turbine engine vane including 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 within the slot of the airfoil, the spar including a pair of extensions each with at least one bi-cast groove formed on opposing ends of the spar, wherein the extensions of the spar are configured to engage with corresponding apertures formed in a pair of opposing endwalls, the method comprising:inserting at least one of the pair of extensions into one of the corresponding apertures of the pair of opposing end walls,
injecting a pre-cursor into the at least one bi-cast groove of the at least one of the pair of extensions, and
heating the pre-cursor to form a fixed connection between the inserted extension of the spar and a circumferential surface defining the corresponding aperture of the corresponding end wall, wherein the fixed connection is formed between a lateral face of the at least one of the pair of extensions and the circumferential surface of the corresponding aperture, the lateral face of the at least one of the pair of extensions formed to face perpendicularly from a length of the at least one of the pair of extensions inserted into the one of the corresponding apertures, wherein the lateral face includes the at least one bi-cast groove and the circumferential surface includes at least one complimentary bi-cast groove.

US Pat. No. 10,577,939

TURBINE BLADE WITH THREE-DIMENSIONAL CMC CONSTRUCTION ELEMENTS

Rolls-Royce Corporation, ...

1. A ceramic matrix composite blade comprising a root, a platform, and an airfoil adapted for use in a gas turbine engine, a core having a proximal end and a distal end, the core comprising reinforcing fibers in a ceramic matrix material, and a wrap comprising three-dimensional reinforcing fibers in a ceramic matrix material, the wrap arranged to extend under the proximal end of the core to form an outer surface of the root and the wrap being configured to contact a disk when the ceramic matrix composite blade is assembled into a turbine wheel, and a platform ply arranged to overlie the platform without extending upwardly into the airfoil or downwardly into the root, wherein the wrap extends outwardly away from the core to form at least a portion of the platform, and wherein the core is exposed between a pressure-side section and suction side section of the wrap along a leading edge of the airfoil and wherein the core is exposed between a pressure-side section and suction side section of the wrap along a trailing edge of the airfoil.

US Pat. No. 10,563,535

KEYSTONED BLADE TRACK

Rolls-Royce Corporation, ...

1. A blade-track system for a gas turbine engine, the blade-track system comprisinga plurality of blade track segments positioned circumferentially around a central axis to form a ring providing a blade track, each blade track segment comprising ceramic-matrix composite materials and shaped to extend part-way around the central axis, each blade track segment including opposing circumferential end faces and a radially outer surface, and
a track biaser positioned to surround the blade track,
wherein the end faces of the blade track segments are engaged with one another and the track biaser is configured to bias the blade track segments toward the central axis such that each blade track segment acts as a keystone to maintain the form of the ring,
wherein the end faces of the blade track segments are configured to engage and form a resultant radially-outward force away from the central axis against the track biaser and wherein the track biaser is positioned to engage the radially outer surface of the blade track segments to provide a radially-inward force against the blade track segments,
wherein the track biaser includes at least one unitary ring sized to surround the blade track and engage the outer surfaces of the blade track segments, and
wherein the at least one unitary ring includes a first unitary ring substantially aligned with an axially-forward face of the blade track segments and a second unitary ring substantially aligned with an axially-aft face of the blade track segments.

US Pat. No. 10,513,981

HEAT EXCHANGER ASSEMBLY FOR A GAS TURBINE ENGINE PROPULSION SYSTEM

Rolls-Royce North America...

1. A propulsion system for an aircraft, the propulsion system comprisinga gas turbine engine including an engine core defining a central axis and a fan coupled to the engine core, the fan configured to discharge pressurized bypass air that is passed by the engine core through a fan duct that extends along the central axis coaxially with the engine core, and
a nacelle mounted to the gas turbine engine, the nacelle including an outer shroud that surrounds at least a portion of the engine core defining a portion of the fan duct and a strut that extends away from the engine core through the fan duct to the outer shroud,
a heat exchanger assembly fluidly coupled to the gas turbine engine to cool fluid or gas from the gas turbine engine and return the cooled fluid or gas to the gas turbine engine, the heat exchanger assembly including an inlet duct having at least a portion positioned in the strut, a heat exchanger housing coupled fluidly to the inlet duct and positioned radially inward of the strut relative to the central axis, heat exchangers housed by the heat exchanger housing, and
a valve system disposed downstream of the inlet duct and movable in the heat exchanger housing from a first position arranged to direct pressurized bypass air received from the inlet duct into contact with the heat exchangers to a second position arranged to divert pressurized bypass air received from the inlet duct through at least one bypass passageway disposed between the heat exchangers, away from the inlet duct without contacting the heat exchangers, and directly to the fan duct.

US Pat. No. 10,224,854

ACTIVE DAMPING OF SYNCHRONOUS GRID OSCILLATIONS USING PARTIAL POWER CONVERTER

ROLLS-ROYCE NORTH AMERICA...

1. A method to damp oscillations in a synchronous alternating current (AC) grid, the method comprising:receiving current from the synchronous AC grid through a phase of a n-phase supply line;
supplying the current received from the synchronous AC grid to a phase of a synchronous motor;
detecting a sub-harmonic oscillation in the current received from the synchronous AC grid; and
damping the sub-harmonic oscillation by:
shunting a portion of the current away from the phase of the synchronous motor during a first time period in an upper-half of the sub-harmonic oscillation, and/or
supplying compensation current from a partial power converter to the phase of the synchronous motor during a second time period in a lower-half of the sub-harmonic oscillation.

US Pat. No. 10,221,715

TURBINE SHROUD WITH AXIALLY SEPARATED PRESSURE COMPARTMENTS

Rolls-Royce North America...

1. A turbine shroud for use in a gas turbine engine, the turbine shroud comprisinga carrier comprising metallic materials and adapted to be coupled to a turbine case,
a blade track comprising ceramic-matrix composite materials and coupled to the carrier, the blade track including a runner extending at least a portion of the way around a central axis so that the runner is adapted to block hot gasses from passing over a turbine wheel surrounded by the turbine shroud, and
at least one bulkhead included in the carrier that radially interconnects the carrier and the runner to effect engagement between the carrier and the runner and to divide a space between the carrier and the runner into at least a first cooling-air cavity and a second cooling-air cavity, the at least one bulkhead formed to include a plurality of pressure-control holes sized to cause the first cooling-air cavity to have a first pressure and the second cooling-air cavity to have a second pressure, lower than the first pressure.

US Pat. No. 10,704,466

HIGH-MACH VEHICLE COOLING

Rolls-Royce North America...

10. A cooling assembly of a turbine-based combined cycle (TBCC) system of an aircraft, the TBCC system including a combined cycle power assembly having a gas turbine engine and a scramjet engine, and a flow valve selectively positionable to block flow into the gas turbine engine, the gas turbine engine and the scramjet engine each adapted for connection with an inlet to receive ambient air from the environment, the scramjet engine adapted for supersonic operation permitting cocooning of the gas turbine engine by a position of the flow valve blocking flow into the gas turbine engine, the cooling assembly comprisingan intake for receiving ambient air,
a turbine-generator arranged in communication with the intake to receive the ambient air from the intake for driving the turbine-generator to cool the ambient air, the turbine-generator directing exhausted ambient air into a first air channel and a second air channel separate from the first air channel, wherein the turbine-generator produces electric power for use by auxiliaries of the aircraft when the gas turbine engine is cocooned, and
a refrigeration cycle for further cooling the ambient air after driving the turbine, the refrigeration cycle directly thermally coupled with the first air channel and the second air channel.

US Pat. No. 10,654,576

GAS TURBINE ENGINE WITH THERMOELECTRIC COOLING AIR HEAT EXCHANGER

Rolls-Royce North America...

15. A method of operating a gas turbine engine for providing propulsion for an aircraft, the method comprisingdetermining an operational state of the gas turbine engine as one of ground idle, takeoff, climb, cruise, and flight idle,
comparing a thermal differential existing between a cooling air system and a coolant system of the gas turbine engine to a predetermined threshold corresponding to the operational state of the gas turbine engine,
based on the determined operational state and a result of the comparing of the thermal differential to the predetermined threshold corresponding to the operational state of the gas turbine engine, selectively:
applying voltage across a thermoelectric section of a thermoelectric heat exchanger of the gas turbine engine in response to determination that the thermal differential does not exceed the predetermined threshold, or
extracting electric power from the thermoelectric section of the thermoelectric heat exchanger in response to determination that the thermal differential does exceed the predetermined threshold.

US Pat. No. 10,654,079

HIGH POWER CLUTCH WITH CLEANING FEATURES

Rolls-Royce North America...

1. An aircraft comprisingan airframe,
a gas turbine engine configured to provide thrust for propelling the airframe,
a high power accessory system configured to be driven by the gas turbine engine, and
a high power clutch configured to selectively transmit rotation from the gas turbine engine to the high power accessory system, the high power clutch including a case, an input shaft, an output shaft, and clutch plates housed in the case and mounted for rotation about an axis,
wherein the case of the high power clutch is opened to air surrounding the high power clutch, and wherein the clutch plates include perforations formed through the clutch plates that are sized to carry a foamed cleaning agent through associated clutch plates and evacuation channels formed in surfaces of associated clutch plates that face along the axis that are sized to carry the foamed cleaning solution radially outward away from the axis upon rotation of the clutch plates.

US Pat. No. 10,655,479

TURBINE WHEEL ASSEMBLY WITH CERAMIC MATRIX COMPOSITE BLADES

Rolls-Royce Corporation, ...

1. A wheel assembly for a gas turbine engine, the wheel assembly comprisinga metallic disk that includes a body arranged around an axis and a plurality of disk posts that extend radially away from the body to define an axially extending first disk slot and an axially extending second disk slot spaced apart circumferentially relative to the first disk slot,
a plurality of ceramic matrix composite blades adapted to interact with gases during use of the gas turbine engine and each of the plurality of ceramic matrix composite blades having substantially similar radial lengths, and
a blade-attachment system that couples the plurality of ceramic matrix composite blades to the disk for rotation with the disk, the blade-attachment system includes a first attachment member received in the first disk slot to couple the first attachment member to the disk and a second attachment member received in the second disk slot to couple the second attachment member to the disk,
wherein the plurality of ceramic matrix composite blades includes a first blade, a second blade, and a third blade, each of the first blade, the second blade, and the third blade includes a root and an airfoil that extends radially away from the root, the root of the first blade is received in the first attachment member to couple the first blade to the first attachment member, the root of the second blade is received in the second attachment member to couple the second blade to the second attachment member, and the root of the third blade is received between the first attachment member and the second attachment member to couple the third blade to the blade-attachment system, and
wherein the radial length of each of the plurality of ceramic matrix composite blades is measured from a bottom of the root to a tip of the airfoil.

US Pat. No. 10,655,491

TURBINE SHROUD RING FOR A GAS TURBINE ENGINE WITH RADIAL RETENTION FEATURES

Rolls-Royce Corporation, ...

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 including a body plate that extends partway around a central axis of the gas turbine engine and a bracket that extends radially inward away from the body plate, the body plate defines a first pocket that extends radially outward into the body plate, and the bracket defines an axially opening channel located radially inward of the first pocket,
a blade track segment comprising ceramic matrix composite materials, the blade track segment including a runner shaped to extend partway around the central axis to define a primary gas path and a hanger that extends radially outward from the runner and into the channel of the carrier segment, the hanger being sized relative to the channel such that a gap is provided between a radially outer surface of the hanger and a radially inner surface of the carrier segment defining the first pocket to allow for different thermal expansion rates between the carrier segment and the blade track segment to avoid binding stresses, and
a first spring located in the first pocket and configured to positively position the blade track segment radially inward toward the central axis so that the hanger engages the bracket, the first spring sized to extend radially inward out of the first pocket into the channel and into engagement with the blade track segment in order to apply a retention force to the hanger of the blade track segment in order to position the blade track segment so that a clearance distance between the runner and turbine blades located in the primary gas path can be measured while the turbine shroud segment is not in operation,
wherein the body plate includes a radial inner surface that faces the hanger and an aft side surface that extends radially away from the radial inner surface and the first pocket extends axially through the aft side surface into the body plate.

US Pat. No. 10,584,605

SPLIT LINE FLOW PATH SEALS

Rolls-Royce Corporation, ...

1. A gas turbine engine assembly, the assembly comprisinga first component comprising ceramic matrix materials, the first component including a panel that extends partway around a central axis, that is arranged to separate a high pressure zone located radially outward of the panel from a low pressure zone located radially inward of the panel, and that is formed to include a first chamfer surface that extends radially outward and circumferentially away from a first side of the first component,
a second component comprising ceramic matrix materials, the second component including a panel that extends partway around the central axis, that is arranged to separate the high pressure zone from the low pressure zone, and that is formed to include a second chamfer surface that extends radially outward and circumferentially away from a second side of the second component, and
a seal assembly arranged between the first chamfer and the second chamfer when the first side of the first component is arranged in confronting relation to the second side of the second component, the seal assembly including a rod engaged with the first chamfer and the second chamfer that is configured to block gasses from passing through the interface of the first side included in the first component with the second side included in the second component and a rod locator that is configured to engage the rod to hold the rod in place relative to the first component and the second component.

US Pat. No. 10,577,960

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, the carrier segment including a seal channel that is divided into a plurality of sections adapted to be pressurized at different pressures,
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 the 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 at least a portion of the seal member is received in the seal channel formed in the carrier segment.

US Pat. No. 10,563,613

COANDA DEVICE FOR A ROUND EXHAUST NOZZLE

Rolls-Royce North America...

1. A gas turbine engine system comprising:an engine core configured to produce discharge air directed through a round exhaust nozzle along a central axis, and
a thrust director arranged near the round exhaust nozzle and configured to redirect the discharge air by applying flow to the discharge air near the exhaust nozzle, the thrust director including an arcuate momentum nozzle an arcuate coanda nozzle, and a round, annular manifold extending around the central axis and directly coupled to the arcuate momentum nozzle and the arcuate coanda nozzle and configured to provide flow to the arcuate momentum nozzle and the arcuate coanda nozzle,
wherein the arcuate momentum nozzle extends along a constant radius from the central axis and is configured to discharge flow generally perpendicular to and toward the central axis through a momentum nozzle passageway defined between a first momentum nozzle component and a second momentum nozzle component, wherein the first momentum nozzle component is coupled to the manifold and has a generally C-shaped cross section, and the second momentum nozzle component is coupled to the manifold and has a generally L-shaped cross section, and
wherein the arcuate coanda nozzle extends along a constant radius from the central axis and includes a first coanda nozzle component coupled to the manifold and having a generally C-shaped cross section that is configured to cause the arcuate coanda nozzle to discharge flow generally parallel to and along the central axis, and the arcuate coanda nozzle includes a second coanda nozzle component coupled to the manifold that defines an arcuate coanda surface that influences the discharge air to turn and exit the exhaust nozzle perpendicular to the central axis.

US Pat. No. 10,544,834

BEARING FOR USE IN HIGH SPEED APPLICATION

Rolls-Royce North America...

1. A bearing system comprisingan outer race that extends around a central axis,
an inner race that extends around and rotates in a first direction about the central axis relative to the outer race and spaced apart radially from the outer race to define a bearing cavity therebetween,
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 lubrication system configured to provide lubrication to the plurality of internal rotating components during rotation of the inner race, the lubrication system including a side-jet injector located in spaced-apart relation to the outer race in a fixed position relative to the outer race and configured to inject a stream of lubrication from an outlet formed in the side-jet injector in an axial direction toward the plurality of internal rotating components, and a windage barrier located adjacent to the outlet of the side-jet injector in a fixed position relative to the side-jet injector and the outer race upstream from the outlet and configured to establish a zone of stagnant fluid downstream of the windage barrier such that the stream moves through the zone from the outlet to the plurality of rotating components.

US Pat. No. 10,358,977

PHASE CHANGE MATERIAL COOLING SYSTEM FOR A VEHICLE

Rolls-Royce North America...

1. An apparatus comprising:an aircraft having one or more propulsion engines and an external pod including a turbine apart from the one or more propulsion engines, the turbine in power communication with an electrical generator and a heat exchange system that includes a working fluid in communication with a phase change heat exchanger, the aircraft further including an implement in thermal communication with the heat exchange system, the external pod including an electrical line in communication with the implement.

US Pat. No. 10,221,713

SHROUD CARTRIDGE HAVING A CERAMIC MATRIX COMPOSITE SEAL SEGMENT

ROLLS-ROYCE CORPORATION, ...

1. A segmented turbine shroud for radially encasing a rotatable turbine in a gas turbine engine, said shroud including a plurality of cartridges, one or more cartridges comprising:a carrier segment; and
a ceramic matrix composite seal segment carried by said carrier segment,
said carrier segment and said CMC seal segment being positioned so that a surface of said carrier segment and a surface of said CMC seal segment form a mating region proximate an entire perimeter of said CMC seal segment to thereby form a cavity bounded by surfaces of said carrier segment and said CMC seal segment,
wherein said mating region extends both laterally and axially around the entire perimeter of said CMC seal segment.

US Pat. No. 10,654,162

THERMAL MANAGEMENT SYSTEM

Rolls-Royce North America...

1. A thermal management system comprising:thermal management components including an electrical source configured to supply electricity to an electrical apparatus, the electrical source further configured to store thermal energy, the thermal management components further including a cooling source; and
a controller configured to:
cause the electrical apparatus to be cooled with the cooling source, and
delay cooling the electrical source with the cooling source in response to detection of a dynamic thermal load and a state of charge of the electrical source being greater than a predetermined charge threshold of the electrical source, the state of charge of the electrical source comprising a measure of cooling capacity available in the electrical source, wherein the dynamic thermal load is at least partially created by the electrical apparatus and the dynamic thermal load varies in response to an increased amount of power provided to the electrical apparatus by the electrical source.

US Pat. No. 10,655,482

VANE ASSEMBLIES FOR GAS TURBINE ENGINES

Rolls-Royce Corporation, ...

1. A vane ring for use in a gas turbine engine, the vane ring comprisinga plurality of metal spars, each spar including a web section having an airfoil shape, a first end connector coupled to a radially outer portion of the web section, and a second end connector coupled to a radially inner portion of the web section,
a plurality of inner end wall segments positioned to function as a continuous hoop and engage the metal spars, each inner end wall segment including a first flow surface positioned to guide expanding hot gases along a flow path through the gas turbine engine and at least one locator hole sized to receive the second end connector of the metal spars,
a unitary outer end wall forming a one-piece continuous hoop, the outer end wall comprising ceramic-matrix materials and including a second flow surface positioned to cooperate with the first flow surface of the inner end wall segments to form the flow path and a plurality of locator holes sized to receive the first end wall connectors to locate the metal spars circumferentially along the outer end wall, wherein the locator holes are sized to allow the metal spars to tilt relative to the inner end wall segment, and
a plurality of inner end caps having an opening to receive the second end connectors and positioned to engage a radially inner surface of the inner end wall, wherein the opening has a geometry that is the same as a geometry of one of the plurality of locator holes.