US Pat. No. 9,234,463

THERMAL MANAGEMENT SYSTEM FOR A GAS TURBINE ENGINE

United Technologies Corpo...

1. A thermal management system for a gas turbine engine comprising:
means for delivering a first stream of cooling air to at least one forward cavity for controlling disk temperature and shaft
temperature;

means for allowing said first stream of cooling air to exit said at least one forward cavity;
means for delivering a second stream of cooling air to at least one aft cavity for providing disk web cooling; and
means for mixing said first stream of cooling air exiting said at least one forward cavity with an intershaft flow of cooling
air, wherein said at least one forward cavity comprises a plurality of forward cavities formed by a plurality of disks, hubs
and an outer diameter of a shaft and said at least one aft cavity comprises a plurality of aft cavities formed by a plurality
of disks and the outer diameter of said shaft.

US Pat. No. 9,121,280

TIE SHAFT ARRANGEMENT FOR TURBOMACHINE

UNITED TECHNOLOGIES CORPO...

1. A turbomachine comprising:
a tie shaft extending along an axis;
multiple rotors mounted on the tie shaft;
first and second clamping members secured to the tie shaft and exerting a clamping load between the rotors and clamping members
at multiple interfaces, the clamping load at one of the interfaces including a radial clamping load of greater than 5% of
a total design clamping load at the one interface; and

a friction modifier is provided at the interface, wherein the interface includes at least one of a rough surface finish, a
grit blasted surface, a coating, a spray, a plasma, colloidal particles, adhesives, pastes and additives.

US Pat. No. 9,359,903

GAS TURBINE AND GUIDE BLADE FOR A HOUSING OF A GAS TURBINE

MTU Aero Engines AG, Mun...

13. A method for manufacturing a guide blade for a housing of a gas turbine, the guide blade including a shroud configuration
having a radially outer shroud in the installed state and having a shroud holding device with the aid of which the shroud
is securable on the housing; the guide blade including a turbine blade extending radially inward from the shroud configuration,
the method comprising:
forming at least one air passage channel in the shroud holding device for permitting a flow passage through the shroud holding
device in the installed state of the guide blade,

the shroud holding device including at least one shroud hook on a side of the shroud opposite the turbine blade, with the
aid of which the shroud is secured on the housing, and the shroud hook, which is situated in the area of a profile inlet edge
of the turbine blade, including the air passage channel.

US Pat. No. 9,334,747

APPARATUS AND METHOD FOR SECURING SEALING ELEMENTS

MTU AERO ENGINES AG, Mun...

1. An apparatus for securing sealing elements (12, 14) in an installation slot formed in a turbomachine, having at least one sealing support element (18) comprising at least one sealing element (12, 14) wherein at least one fixing piece (20) is disposed between the sealing support element (18) and a clamping piece (22) associated with the sealing support element (18), wherein the fixing piece (20) can be joined integrally to the sealing support element (18) and to the clamping piece (22) by strut-like connections (24) that can be broken off by means of applying force.

US Pat. No. 9,726,019

LOW NOISE COMPRESSOR ROTOR FOR GEARED TURBOFAN ENGINE

UNITED TECHNOLOGIES CORPO...

1. A gas turbine engine comprising:
a fan, a turbine section having a fan drive turbine rotor, and a compressor rotor;
a gear reduction effecting a reduction in a speed of said fan relative to an input speed from said fan drive turbine rotor;
said compressor rotor having a number of compressor blades in at least one of a plurality of blade rows of said compressor
rotor, and said blades configured to operate at least some of the time at a rotational speed, and said number of compressor
blades in said at least one of said blade rows and said rotational speed being such that the following formula holds true
for said at least one of said plurality of blade rows of the compressor rotor:

(said number of blades×said rotational speed)/60 sec?about 5500 Hz; and
said rotational speed being an approach speed in revolutions per minute.

US Pat. No. 9,470,103

TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A turbomachine, comprising:
a mid turbine vane frame (1) configured to direct a flow of hot gas (4) into a path of hot gas (6) from a high-pressure turbine to a low-pressure turbine, the mid turbine vane frame (1) having a groove walls (32, 34) forming a circumferential groove (12) therein;

a portion of a low-pressure turbine casing (2) delimiting an annular compartment (8);

a radially inner seal (16) in contact with the mid turbine vane frame (1); and

a radially outer ring seal (14) in contact with the portion of the low-pressure turbine casing;

the radially inner seal (16) contacting the radially outer ring seal (14) forming an annular overlap (18) between the radially inner seal (16) and the radially outer ring seal (14);

the radially inner seal (16) and radially outer ring seal (14) arranged in the circumferential groove (12) of the mid turbine vane frame (1); the ring seals (14, 16) being pressed against one of the groove walls (34) in proximity to the path of hot gas by a spring element (48); and

wherein, cooling air is prevented from being drawn out of the annular compartment (8) into the path of hot gas (6).

US Pat. No. 9,470,094

BLADE CASCADE WITH SIDE WALL CONTOURS AND CONTINUOUS-FLOW MACHINE

MTU Aero Engines AG, Mun...

1. A blade cascade for a continuous-flow machine, comprising:
at least one blade channel delimited in a circumferential direction by a pressure-side wall of a first blade and by an opposite
suction-side wall of an adjacent second blade and delimited in a radial direction by two opposite delimiting walls,

in the area of the pressure-side wall, at least one of the delimiting walls being provided with at least one pressure-side
elevation and, in the area of the suction-side wall, the at least one delimiting wall having at least one suction-side depression,
a highest section of the at least one elevation and a lowest section of the at least one depression being located over an
area of 30% to 60% of an extension of the blades in an axial direction, and wherein an axial position of the highest section
and an axial position of the lowest section differ from each other at a maximum by 10% in the axial direction; and

wherein, in the area of the suction-side wall, a second suction-side depression is arranged downstream from the first depression,
the second depression being at a distance from the first depression, separated by a non-contoured section of the delimiting
wall.

US Pat. No. 9,427,937

ANTI-WEAR COATING

MTU AERO ENGINES AG, Mun...

1. A component having a coating thereon, wherein the coating is an anti-wear coating and comprises at least two different
individual layers which alternate, the at least two individual layers comprising a ceramic main layer A and a non-metallic
intermediate layer B configured in such a way that energy is withdrawn from cracks which grow in a direction of the component
by crack branching in layer B so that crack growth is slowed or stopped, one of the layers B being arranged directly on the
component and one of the layers A being arranged at a surface of the coating, and wherein layer B at least one of
comprises a material having preferential sliding planes which are arranged parallel to a surface of the component;
is a multilayer system which comprises ceramic layers as well as layers comprising a material having preferential sliding
planes which are arranged parallel to a surface of the component;

comprises a ceramic material having pores deliberately introduced therein;
comprises a ceramic material comprising deliberately introduced microcracks running parallel to a surface of the substrate;
comprises a ceramic material comprising a deliberately introduced foreign phase.
US Pat. No. 9,845,526

SLIP AND PROCESS FOR PRODUCING AN ALUMINUM DIFFUSION LAYER

MTU AERO ENGINES AG, Mun...

1. A slurry for producing an aluminum diffusion layer, wherein the slurry comprises Al-containing powder, Si-containing powder,
and a binder, the Al-containing powder comprising uncoated powder particles and powder particles which are coated with Si.

US Pat. No. 9,726,038

METHOD OF PRODUCING AN INSULATION ELEMENT AND INSULATION ELEMENT FOR A HOUSING OF AN AERO ENGINE

MTU AERO ENGINES AG, Mun...

1. An insulation element, wherein the element is configured to be capable of being arranged radially above at least one guide
vane in a housing of a thermal gas turbine and consists of a solid body provided with a metallic shell, the solid body consisting
at least partially of a ceramic material, and wherein the insulation element comprises at least one sealing element for being
arranged in a corresponding receptacle of an adjacent insulation element, the at least one sealing element being formed in
such a manner that it undergoes reversible and/or anisotropic deformation upon thermal loading.

US Pat. No. 9,657,395

OXIDATION-RESISTANT LAYER FOR TIAL MATERIALS AND METHOD FOR THE PRODUCTION THEREOF

MTU AERO ENGINES AG, Mun...

1. A protective layer for protecting a TiAl material against oxidation, wherein the protective layer has a layer sequence
which, proceeding from an inner side facing toward the TiAl material, comprises an inner aluminum oxide layer, a first gradient
layer comprising aluminum and a base metal with a base metal content increasing outward toward a surface side, a base metal
layer, a second gradient layer comprising aluminum and a base metal with an aluminum content increasing outward toward the
surface side, and an outer aluminum oxide layer.

US Pat. No. 9,535,012

METHOD FOR THE NON-DESTRUCTIVE TESTING OF WORKPIECE SURFACES

MTU AERO ENGINES AG, Mun...

1. A method for the non-destructive testing of a workpiece surface of a workpiece by fluorescent penetrant testing or dye
penetrant testing, wherein the method comprises:
(a) cleaning a region of the workpiece surface which is to be examined;
(b) applying a liquid fluorescent penetrant or a liquid non-fluorescent dye penetrant to the region of the workpiece surface
which is to be examined, thereby allowing the penetrant to penetrate into possible recesses in the workpiece surface;

(c) removing excess penetrant from the workpiece surface;
(d) applying a developer to the region of the workpiece surface which is to be examined and from which excess penetrant has
been removed;

(e) bleaching the penetrant by a gaseous or liquid oxidant in a layer formed by applied developer on the workpiece surface;
and

(f) visually assessing penetrant that has remained in the recesses present in the workpiece surface.

US Pat. No. 9,500,230

BEARING CAGE AND BEARING MEANS HAVING THIS TYPE OF BEARING CAGE AS WELL AS METHOD FOR DESIGNING, REPAIRING AND/OR REPLACING SUCH A BEARING CAGE

MTU AERO ENGINES AG, Mun...

1. A bearing cage (1) of a bearing, which is produced by means of a generative manufacturing method, comprising:
an outer flange (2) and a bearing seat (8), which are joined together via a plurality of spring beams (3, 5, 7);

a first intermediate ring (4) joined to the outer flange (2) via first spring beams (3);

a second intermediate ring (6) joined to the first intermediate ring (4) via second spring beams (5); and

the second intermediate ring (6) being joined to the bearing seat (8) via third spring beams (7).

US Pat. No. 9,488,069

COOLING-AIR GUIDANCE IN A HOUSING STRUCTURE OF A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A housing structure of a turbomachine, the housing structure annularly surrounding, at least partially, a flow channel,
guide vanes and rotor vanes being arranged in the flow channel, the housing structure comprising:
an inner wall delimiting the flow channel;
an outer wall closed off vis-à-vis the environment;
a cooling-air channel formed at a distance from the inner wall and having at least one air-guiding element, the cooling-air
channel guiding the cooling air exclusively at a distance from the inner wall;

a heat-protection plate and insulation, wherein the insulation is arranged between the heat-protection plate and the cooling-air
channel, wherein the cooling-air channel is formed between the insulation and a radial inside of the outer wall, the cooling-air
channel having the at least one air-guiding element resting against the insulation.

US Pat. No. 9,194,331

FLOW CONDUCTING ASSEMBLY FOR COOLING THE LOW-PRESSURE TURBINE HOUSING OF A GAS TURBINE JET ENGINE

MTU Aero Engines AG, Mun...

1. A gas turbine jet engine comprising:
a main flow channel;
a housing structure radially surrounding the main flow channel, a housing gas flow flowing in the housing structure in a same
direction as a main flow in the main flow channel, the housing structure having an adjustable flow conducting assembly including
at least one flow conducting sheet dividing the housing gas flow into a first partial flow running near the main flow channel
and a second partial flow running at a distance from the main flow channel, the flow conducting sheet being adjustable or
having variable closable openings so that a gas mass flowing in the first partial flow or the second partial flow is adjustable;
wherein the flow conducting assembly is situated in the area of a low-pressure turbine; and

a radially outer bypass flow channel having an outer housing, so that the housing structure, together with the flow conducting
assembly, is situated between the bypass flow channel and the main flow channel.

US Pat. No. 9,551,231

DUCTILE COMPENSATION LAYER FOR BRITTLE COMPONENTS

MTU AERO ENGINES AG, Mun...

1. A turbomachine, wherein the turbomachine comprises a blade element and a receptacle in which the blade element is arranged,
the blade element comprising a fastening element with which the blade element is arranged in the receptacle of the turbomachine,
and the blade element comprising a blade base material in a region of the fastening element, and wherein in a region of a
contact surface with the fastening element, the receptacle comprises a receptacle surface which is formed from a receptacle
surface material that has a higher ductility than the blade base material.
US Pat. No. 9,346,131

NICKEL-BASED SOLDER ALLOY

MTU Aero Engines AG, Mun...

1. Method for producing a solder alloy for soldering a component made of a nickel-based material in which nickel is the largest
alloy component by weight, the method comprising the following steps:
providing three components composed of a first soldering material, a second soldering material, and a base material, wherein
the base material is a nickel-based powder material, which has the same composition as the nickel-based material of the component
to be soldered and is present in a proportion of 45-70% by weight in the total weight of the three components,

the first soldering material is a nickel-based powder material which has nickel as the largest alloy component by weight and
further comprises chromium, cobalt, tantalum, aluminum and boron and is present in a proportion of 15-30% by weight in the
total weight of the three components, and

the second soldering material is a nickel-based powder material which has nickel as the largest alloy component by weight
and further comprises chromium, cobalt, molybdenum, tungsten, boron and hafnium and is present in a proportion of 15-25% by
weight in the total weight of the three components,

wherein both the chromium content and also the cobalt content are higher in the case of the first soldering material than
in the case of the second soldering material, and

mixing together the base material powder, the first soldering material powder and the second soldering material powder, thereby
forming the solder alloy in a powder form.

US Pat. No. 9,694,568

METHOD FOR COATING COMPONENTS

MTU Aero Engines AG, Mun...

1. A method for coating a component of a turbomachine, comprising the steps of:
covering a first surface of the component with a covering device, wherein the covering device is profiled in a zigzag shape;
and

applying a coating material via cold kinetic compaction or kinetic cold gas spraying on the component such that a second surface
of the component is coated with the coating material and such that particles of the coating material are deflected off of
the covering device so that the particles do not adhere to the covering device.

US Pat. No. 9,682,397

DEVICE FOR THE GENERATIVE PRODUCTION OF A COMPONENT

MTU AERO ENGINES AG, Mun...

1. A device (10) for the generative production of a component (12), comprising:
a first supply tank (14a) for the uptake of powder-form material (16);

a first overflow tank (22a) for the uptake of excess powder-form material (16), wherein a first closing device (24a) is assigned to the first overflow tank (22a), the first closing device being switchable between a closed position, in which powder-form material (16) cannot be transported into the first overflow tank (22a), and an open position, in which powder-form material (16) can be transported into the first overflow tank (22a);

a working chamber (26), in which the component (12) is producible layerwise from the powder-form material (16);

the first overflow tank (22a) located between the first supply tank (14a) and the working chamber (26);

a second overflow tank (22b) for the uptake of excess powder-form material (16), wherein a second closing device (24b) is assigned to the second overflow tank (22b), the second closing device being switchable between a closed position, in which powder-form material (16) cannot be transported into the second overflow tank (22b), and an open position, in which powder-form material (16) can be transported into the second overflow tank (22b);

a second supply tank (14b) for uptake of powder-form material (16);

the second overflow tank (22b) located between the second supply tank (14b) and the working chamber (26); and

a transport device (20) by which the powder-form material (16) is transported at least from the first supply tank (14a) over the first overflow tank (22a) to the working chamber (26) and from the working chamber (26) to the second overflow tank (22b) and by which the powder-form material (16) is transported at least from the second supply tank (14b) over the second overflow tank (22b) to the working chamber (26) and from the working chamber (26) to the first overflow tank (22a);

wherein the transport device (20):

is movable from the first supply tank (14a) over the first overflow tank (22a) to the working chamber (26) and from the working chamber (26) over the second overflow tank (22b) to the second supply tank (14b); and

is movable from the second supply tank (14b) over the second overflow tank (22b) to the working chamber (26) and from the working chamber (26) over the first overflow tank (22a) to the first supply tank (14a).

US Pat. No. 9,664,059

SEALING DEVICE AND TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A sealing device (1) for sealing a radially inner gas channel (2) between a guide vane ring (4) and a rotor (6) of a turbomachine, wherein the sealing device (1) has a sealing ring (8) for forming a sealed space (10) with a rear segment, when considered in the direction of a principal flow, of an integral inner ring (12) of the guide vane ring (4), into which penetrates a front platform overhang (14) of a downstream row of rotating blades (16), and wherein the sealing device (1) defines an S-shaped cross-section and has an outer radial flange (18) for connecting to the integral inner ring (12) and a double-walled cylinder (26) with an outer wall structure (28) oriented in a first direction and with an inner wall structure (30) oriented in the opposite direction, which are joined to each other via an a first annular radial spring element (32), wherein the radial flange (18) transitions into the outer wall structure (28) and the cylinder (26) forms the sealing ring (8);
wherein the sealing ring (8) is positioned approximately parallel to the rear segment of the inner ring (12);

wherein the inner wall structure (30) transitions, via a second annular radial spring element (36), into at least one inner body segment (38, 50), which is oriented parallel to the first direction or to the opposite direction, the at least one inner body segment has
a radially inwardmost side that is configured to face a rotor drum having at least one seal fin thereon, and the at least
one inner body segment supports a sealing structure (40) on its radially inward side, the sealing structure in combination with the at least one seal fin on the rotor drum forming
a labyrinth seal directly against the drum by which a flow of the guide vane ring in the region of its vane tips facing the
rotor drum is prevented; the sealing device (1), rotor drum (39), rotor (6) and inner ring (12) defining a sealed space (10) therebetween; and

wherein the sealing device (1) has a uniform, preferably relatively reduced wall thickness over its individual segments integrally formed with one another,
so that the sealing device (1) is resilient within certain limits, wherein, in particular, first annular radial spring element (32) and the second annular radial spring element (36) act as radial spring elements, whereby the sealing device forms a spring configured and arranged for equilibrating a radial
thermal expansion of the guide vane ring.

US Pat. No. 9,506,360

CONTINUOUS-FLOW MACHINE WITH AT LEAST ONE GUIDE VANE RING

MTU Aero Engines AG, Mun...

19. An axial compressor comprising a continuous-flow machine, the continuous flow machine further including:
at least one guide vane ring, the guide vane ring defining radial, circumferential and axial directions,
the guide vane ring including at least one row of a plurality of guide vanes spaced in the circumferential direction, each
guide vane of the plurality of guide vanes having a vane body with a longitudinal axis in the radial direction between a base
and a tip, the plurality of guide vanes including first guide vanes and second guide vanes, each first guide vane being tapered
along the longitudinal axis so that the vane body becomes narrower at the tip than the base and each second guide vane being
tapered in an opposite direction along the longitudinal axis so that the vane body becomes wider at the tip than the base.

US Pat. No. 9,416,676

GAS TURBINE

MTU AERO ENGINES AG, Mun...

1. A gas turbine comprising a housing (1), an outer sealing ring (2) that can be detachably fastened to the housing, and a clamping member (3) for radially clamping the outer sealing ring and the housing together, further comprising a rotation locking member that
has at least one housing groove (10) and a radial flange (20) of the outer sealing ring that is locked against rotation in the housing groove in form-fitting manner with play (sa) in the axial and/or peripheral direction;
wherein an axial length (t1) of the housing groove from a front side of the housing (11) in the direction of through-flow is larger than an axial wall thickness (t2) of the radial flange, resulting in the play in the axial direction such that the sealing ring can move axially with respect
to the housing in an axial range equal to the axial length (t1) of the housing groove; and

wherein the play in the peripheral direction is provided by the housing groove being wider than the rotation locking member
in the peripheral direction.

US Pat. No. 9,726,042

GAS TURBINE DUCT CASING

MTU Aero Engines AG, Mun...

1. A gas turbine duct casing, comprising:
a first wall segment;
a second wall segment; and
a clamp arrangement, wherein the first wall segment is connectable to the second wall segment by the clamp arrangement;
wherein the clamp arrangement has a clamp and a bolt which is passable through a hole defined by the clamp and is screwable
to a nut on a side of the clamp facing away from the first and second wall segments;

wherein a first leg of the clamp is braceable against the first wall segment by the nut and a second leg of the clamp is braceable
against the second wall segment by the nut;

wherein the bolt has a head which is guidable in a form-fitting manner in a groove on an outside of the first wall segment
facing away from a gas duct;

and wherein the clamp has a first stop and a second stop which extend around the groove on opposing sides.

US Pat. No. 9,617,861

GUIDE VANE ARRANGEMENT AND METHOD FOR MOUNTING A GUIDE VANE

MTU AERO ENGINES AG, Mun...

1. A guide vane arrangement having at least one guide vane (110) with a radial inner journal (111), which engages in a borehole of a bushing (3), which is disposed in a second borehole of an inner ring segment (2), wherein a rotation-resistant connection is present between the journal (111) and the bushing (3) and a slide mounting is present between the bushing (3) and the borehole of the inner ring segment (2), wherein the journal (111) has an outer thread and the bushing (3) has an inner thread screwed therewith providing a screw connection; and a screw locking device to secure the screw connection
of the journal (111) to the bushing (3); a threaded sleeve disposed between the journal (111) and the bushing (3), and the journal (111) and the bushing (3) being connected to the threaded sleeve; the bushing (3), with journal (111) secured thereto, with threaded sleeve disposed therebetween, being freely rotatable within the second borehole of the inner
ring segment (2) thereby permitting adjustment of the guide vane (110), attached to the radial inner journal (111), about its longitudinal axis.

US Pat. No. 9,506,368

SEAL CARRIER ATTACHMENT FOR A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A housing structure for a turbomachine comprising:
a stationary blade ring and a seal carrier ring annularly surrounding a flow channel of the turbomachine, the stationary blade
ring and the seal carrier ring being connected to each other,

either the stationary blade ring or the seal carrier ring having a one-sidedly open receptacle in a radial direction in relation
to the flow channel; and

a radially projecting engagement area corresponding to the other of the seal carrier ring or stationary blade ring and accommodatable
in the receptacle so as to be held in a form-fitting manner in the axial direction and in the radial direction in the direction
of a closed side of the receptacle but freely movable with respect to the receptacle and the seal carrier ring or stationary
blade ring having the receptacle in the radially opposite direction in the direction of the receptacle opening, wherein the
seal carrier ring has a run-in coating, moving blade tips rubbing against the run in coating.

US Pat. No. 9,068,826

CHECKING A BLADE CONTOUR OF A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A method for checking a blade contour of a turbomachine, comprising:
registering an actual contour of a blade;
scaling a setpoint contour of the blade, the scaling being a shape-invariant or geometrically similar change of the setpoint
contour by enlargement or reduction of the setpoint contour as a whole, the setpoint contour being scaled so that a distance
to the actual contour becomes minimal; and

comparing the actual contour to the scaled setpoint contour, in order to determine if the actual contour is within a predetermined
tolerance band around the scaled setpoint contour.

US Pat. No. 10,294,814

ELLIPSOIDAL INNER CENTRAL BLADE STORAGE SPACE

MTU Aero Engines AG, Mun...

1. An inner ring for an adjustable guide vane assembly, comprising:an inner ring having a plurality of uptakes, each for the bearing of an adjustable guide vane, wherein the uptakes each create an essentially ellipsoidally shaped uptake space for a guide vane head of the respective guide vane.

US Pat. No. 9,765,625

TURBOMACHINE BLADE

MTU Aero Engines AG, Mun...

1. A blade for a turbomachine, comprising:
an impact chamber; and
a single impulse body movably situated in the impact chamber, a clearance between the impulse body and the impact chamber
in all directions being at least 0.1 mm and at most 1.5 mm and wherein a mass of the impulse body is at least 10 mg and at
most 1.5 g.

US Pat. No. 9,759,231

MID-FRAME FOR A GAS TURBINE AND GAS TURBINE

MTU AERO ENGINES AG, Mun...

1. A mid-frame (10) for a gas turbine, having at least one outer casing element (12) and having a plurality of duct segments (16), which are arranged in succession in the radial direction on the inside of the outer casing element (12) and in the peripheral direction of the outer casing element (12) and by which segments, in each case, at least one duct (18) through which a first gas can flow is delimited at least in the radial direction at least partially,
comprising a ring element (22) that is common to the duct segments (16) and is formed in one piece and the duct segments (16) are held at the outer casing element (12);

the ring element (22) having a ring body (24), from which tabs (26) protrude in the radial direction inwardly, where the duct segments (16) are retained at the ring element (22) by the tabs (26), with passage openings (28) defined between the tabs (26) in the peripheral direction, openings between the tabs (26) are spaced apart from one another in the peripheral direction.

US Pat. No. 9,416,672

HOUSING STRUCTURE WITH IMPROVED SEAL AND COOLING

MTU AERO ENGINES AG, Mun...

1. A housing structure for a turbo-engine having an outer housing wall and an inner housing wall, the inner and outer housing
walls being spaced apart radially from one another and annularly enclosing an axial flow channel for the turbo-engine, the
housing structure comprising:
a first heat shield and a second heat shield, both heat shields being disposed radially between the inner housing wall and
the outer housing wall, and

a bar projecting from an inner surface of the inner housing wall at least somewhat in a radial direction and extending to
a radially outer end directly abutting an inner surface of the outer housing wall, the bar including first and second broadening
elements disposed, respectively, on opposing sides of the bar that face away from one another,

the first broadening element being an expansion of a first cross-section of the bar at a first radial location in a first
direction transverse to the radial direction;

the second broadening element being an expansion of a second cross-section of the bar at a second radial location in a second
direction transverse to the radial direction; and

wherein each of the first and second broadening elements includes a respective sealing face against which the respective first
and second heat shield is positioned in a sealing manner; and

wherein each of the first and second broadening elements is disposed radially between the outer housing wall and the inner
housing wall such that neither of the broadening elements contacts the outer housing wall or the inner housing wall;

wherein an inner radial side of each of the first and the second heat shields and the inner housing wall delimit, respectively,
at least one inner chamber that is spaced apart from the outer housing wall; and

wherein an outer radial side of each of the first and the second heat shields and the outer housing wall delimit at least
one respective outer chamber that is spaced apart from the inner housing wall; and the respective inner chambers and the respective
outer chambers are sealed off from one another.

US Pat. No. 9,095,900

GENERATIVE PRODUCTION METHOD AND POWDER THEREFOR

MTU AERO ENGINES AG, Mun...

1. A generative production method for producing a component by repeated, consecutive selective melting and/or sintering of
a powder by heat introduced by beam energy, such that powder particles fuse and/or sinter in layers, wherein the method comprises
using powder particles of a first material which are surrounded in part or over an entire surface thereof by a second material
that has a lower melting point than that of the first material and/or has a property which lowers the melting point of the
first material when it is mixed with the first material, the first material being selected from Ni-based alloys, Fe-based
alloys, Co-based alloys, alloys of formula MCrAl where M is Ni or Co, tungsten, and tungsten alloys, and the second material
comprising on or more of boron, germanium and silicon; and wherein
(i) the first material has a melting point of more than 1500° C.; and/or
(ii) the method further comprises subjecting the component to a heat treatment with and without pressure, such that the second
material is distributed uniformly in the first material; and/or

(iii) the method further comprises preheating the powder or heating the powder by energy in addition to the beam energy.

US Pat. No. 9,850,766

BLADE CASCADE

MTU Aero Engines AG, Mun...

1. A blade cascade for a turbomachine, the blade cascade comprising:
a plurality of blades including a monocrystalline material, each blade having a crystal orientation value dependent on a crystal
orientation of the monocrystalline material of the blade; and the respective crystal orientation values of first blades of
the plurality of blades being less than a first limiting value and the respective crystal orientation values of second blades
of the plurality of blades being at least equal to the first limiting value, each of the first blades having a first body
arrangement having at least one first body mounted movably on said each first blade; and each of the second blades having
a second body arrangement differing from the first body arrangement, and having at least one second body mounted movably on
said each second blade, wherein each of the first body arrangement and the second body arrangement is implemented in accordance
with the crystal orientation of the first blade and the second blade respectively, and wherein the first limiting value is
at least 10% or at most 90% of a maximum crystal orientation value of the blade cascade.

US Pat. No. 9,839,977

APPARATUS FOR LASER MATERIALS PROCESSING

MTU Aero Engines AG, Mun...

1. An apparatus for laser materials processing comprising:
a stationary laser for generating a laser beam;
a movable laser head, the laser head translationally movable along at least two independent spatial directions and within
or along a plane and connected to the stationary laser via cable, the cable containing a light guide, the laser head emitting
a laser beam capable of processing a material; the laser head including a suction device creating a suction flow parallel
to a direction of the laser beam; and

a temperature conditioning device including an induction coil.

US Pat. No. 9,784,120

TURBOMACHINE STAGE AND METHOD FOR DETECTING A SEALING GAP OF SUCH A TURBOMACHINE STAGE

MTU Aero Engines AG, Mun...

1. A turbomachine stage, comprising:
a housing;
a rotor blade arrangement disposed within the housing, wherein the rotor blade arrangement has an exterior shroud band section
with a sealing flange; and

a sensor arrangement including a first sensor arranged on the housing, wherein a radial clearance to a circumferential surface
of the sealing flange is detectable by the first sensor;

wherein the sealing flange has a recess arrangement with a radial recess and a radial projection;
wherein the sensor arrangement further includes a second sensor;
wherein respective sensing surfaces of the first sensor and the second sensor form equally sized angles in opposite directions
with an axis of rotation of the turbomachine stage.

US Pat. No. 9,745,850

BLADE CASCADE AND CONTINUOUS-FLOW MACHINE

MTU Aero Engines AG, Mun...

1. A blade cascade of a continuous-flow machine, comprising:
a first blade and a second blade;
a first side wall and a second side wall; and
a blade channel, wherein the blade channel is defined in a peripheral direction by a pressure side of the first blade and
by a suction side of the second blade and is defined in a radial direction by the first side wall and the second side wall;

wherein at least one of the first side wall and the second side wall includes a side wall contouring, wherein the side wall
contouring is wave-like in the peripheral direction and has an elevation, a depression, and a rib that has a blade profile
with a pressure side and a suction side, wherein the elevation is a pressure side elevation which extends to a pressure-side
blade wall of the first blade and starts from a leading edge of the first blade and extends to a trailing edge of the first
blade, wherein the elevation has a smaller amplitude than the rib, and wherein the depression extends to a suction-side blade
wall of the second blade.

US Pat. No. 9,771,827

DAMPING DEVICE FOR BEING SITUATED BETWEEN A HOUSING WALL AND A CASING RING OF A HOUSING OF A THERMAL GAS TURBINE

MTU Aero Engines AG, Mun...

1. A damping device for being situated between a housing wall of a housing of a thermal gas turbine and a casing ring, the
casing ring having an area radially internal with regard to a rotation axis of a rotor of the thermal gas turbine and facing
rotating moving blades of the thermal gas turbine, the damping device comprising:
at least sectionally a porous damping structure, the porous damping structure being radially elastic and gas permeable in
a peripheral direction with respect to the rotation axis of the rotor, the porous damping structure being axially gas impermeable.

US Pat. No. 9,752,455

COMPONENT SUPPORT AND TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A component support of a turbomachine, the component support comprising:
at least first and second annular components on a stator side of the turbomachine, the first and second annular components
in axial contact with each other, the first annular component having a support ring surface, and the second annular component
having a plurality of bearing sections distributed over a circumference, the second annular component being in contact with
the support ring surface via the bearing sections distributed over the circumference,

the support ring surface having a plurality of radial groove recesses, the recesses have a depth-to-width ratio of 1:5 to
1:20, each bearing section having two side edge areas each overlapping one of the recesses and an uncontoured support ring
surface area extending between adjacent bearing sections, the uncontoured support ring surface continuing an original contour
of the support ring surface.

US Pat. No. 9,644,488

TURBINE STAGE WITH A BLOW-OUT ARRANGEMENT AND METHOD FOR BLOWING OUT A SEALING GAS FLOW

MTU Aero Engines AG, Mun...

1. A turbine stage, comprising:
a gas channel;
a rotor assembly disposed in the gas channel;
a cavity, wherein the cavity is in communication with the gas channel and wherein the cavity is defined by a front side of
a rotor element and by a rotor element-mounted seal; and

a blow-out arrangement with a gas passage, wherein a sealing gas flow is blowable out of the gas passage into the cavity;
wherein the gas passage is configured with a swirl in a circumferential direction and includes an outlet opening which is
offset radially outwardly from the rotor element-mounted seal;

and wherein the gas passage is a through-borehole of a pipe and wherein the pipe is fastened to the rotor element or integrally
configured with the rotor element.

US Pat. No. 9,605,551

AXIAL SEAL IN A CASING STRUCTURE FOR A FLUID FLOW MACHINE

MTU Aero Engines AG, Mun...

1. A casing structure for a fluid flow machine comprising:
an outer casing wall;
an inner casing wall, the inner and outer casing walls annularly surrounding a flow channel of the fluid flow machine and
being spaced apart in a radial direction with respect to the flow channel, and at least one cavity being formed between said
inner and outer casing walls, the cavity annularly surrounding the flow channel and having no defined openings to the flow
channel; and

an axial seal dividing the cavity axially into at least two regions so that different pressure conditions are created according
to an axial position of the regions, the different pressure conditions corresponding to pressure conditions in the flow channel.

US Pat. No. 9,605,554

TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A turbomachine, wherein the turbomachine comprises a rotor and a stator and wherein, in at least one radial gap between
the rotor and the stator, there is arranged a seal for reducing the radial gap, which seal has two opposite coatings, a first
coating on a stator section delimiting the at least one radial gap radially to an outside and a second coating on a rotor
section delimiting the at least one radial gap radially to an inside, the first and second coatings being built up from a
ceramic powder having a particle size smaller than 1.0 ?m.

US Pat. No. 9,580,774

CREEP-RESISTANT, RHENIUM-FREE NICKEL BASE SUPERALLOY

MTU AERO ENGINES AG, Mun...

1. A nickel base alloy, wherein the alloy is substantially free of rhenium and has a solidus temperature of more than 1320°
C., wherein precipitates of a ??-phase are present in a ?-matrix with a fraction of from 40 to 50 vol % at temperatures of
from 1050° C. to 1100° C., and a ?/?? mismatch at temperatures of from 1050° C. to 1100° C. is from ?0.15% to ?0.25%, and
wherein the alloy comprises:
aluminum from 11 to 12 at %,
cobalt from 8 to 10 at %,
chromium from 6 to 8 at %,
molybdenum from 0.5 to 1.5 at %,
tantalum from 2 to 3.5 at %,
titanium from 1 to 2 at %,
tungsten from 2 to 3 at %,
hafnium from 0.05 to 0.3 wt % or absent,
remainder nickel and unavoidable impurities,a tungsten content of the ?-matrix being greater than a tungsten content of the precipitates of the ??-phase.

US Pat. No. 10,006,300

ARMORING SEALING FINS OF TIAL VANES BY INDUCTION BRAZING HARD-MATERIAL PARTICLES

MTU Aero Engines AG, Mun...

1. A method of armoring a titanium aluminum (TiAl) vane of a turbomachine, comprising the steps of:applying a mixture of a hard material and a braze material to the TiAl vane, wherein the TiAl vane comprises intermetallic phases of one or both of ?-TiAl and ?2-Ti3Al; and
brazing the mixture on the TiAl vane by an inductive heating process, wherein the braze material is a nickel (Ni)-based braze material.

US Pat. No. 10,001,016

TURBOMACHINE BLADE HAVING A MAIN BODY INCLUDING A PLANE FIRST ATTACHMENT SURFACE

MTU Aero Engines AG, Mun...

1. A turbomachine blade comprising:a main body including a plane, first attachment surface having a first rim contour; and
a cover including a plane, second attachment surface having a second rim contour and welded to the first attachment surface; and
a tuning body configuration having at least one tuning body for contacting an inner wall of the cover by impact therewith, a gap being formed between the first and second rim contour;
wherein the first attachment surface is configured radially inwardly or outwardly from an airfoil for flow deflection on a surface of an inner or outer shroud facing away from the airfoil.

US Pat. No. 9,862,059

SURFACING OF ADDITIVELY MANUFACTURED COMPONENTS AND CORRESPONDING MANUFACTURED COMPONENTS OF A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A method for manufacturing a component, wherein an additive method is used at least partially for the manufacture of a
component from a plurality of layers each having edges, the plurality of layers defining a first surface region having a rough
surface defined by the edges of the plurality of layers of the layerwise construction of the plurality of layers, wherein
the first surface region of the additively manufactured part of the component is provided with a smoothing layer, which is
deposited by vapor deposition, the smoothing layer including a diffusion layer that at least partially diffuses into the plurality
of layers of the first surface region adhering thereto, whereby the smoothing layer forms a part of and remains on the component,
and whereby the smoothing layer provides a protective coating to the first surface region during use of the component.

US Pat. No. 9,856,736

FISH MOUTH SEAL CARRIER

MTU Aero Engines AG, Mun...

1. A fish mouth seal carrier for a guide vane arrangement of a gas turbine, the gas turbine having a guide vane with a radial
inner platform defining a first axial extent and having a single spoke extending radially inward from the radial inner platform,
a plurality of moving blades disposed axially adjacent to the guide vane, at least one of the moving blades having an axial
flange disposed radially inward from the radial inner platform, the axial flange extending axially from the moving blade defining
a second axial extent to a position radially across from the radial inner platform of the guide vane and within the first
axial extent such that a first portion of the first axial extent of the radial inner platform axially overlaps at least some
of the second axial extent of the axial flange of the moving blade on a radially outward side of the axial flange of the moving
blade, the seal carrier comprising:
a first half-shell element having
a radial inner axial arm section,
a radial arm and
a radial outer axial arm section;
a second half-shell element having
a radial inner axial arm section,
a radial arm and
a radial outer axial arm section;
the first half-shell element connected to the second half-shell element so as to together form a box profile, the box profile
including

a radial inner axial arm formed by connecting adjacent ends of the respective radial inner axial arm sections,
the two radial arms, the radial inner ends of the radial arms being connected, respectively, to non-adjacent ends of the respective
radial inner axial arm sections, and

a radial outer axial arm formed by connecting non-adjacent ends of the respective radial outer axial arm sections, respectively,
to radial outer ends of the two radial arms;

the box profile being mounted to the single spoke extending radially inward from the radial inner platform;
a sealing element mounted on the radial inner axial arm of the box profile; and
at least one of the first and second half-shell elements including an integrally formed axial flange,
the integrally formed axial flange being configured to extend axially toward the at least one of the plurality of moving blades
disposed axially adjacent to the guide vane having the axial flange extending axially therefrom,

the integrally formed axial flange being disposed radially inward from the axial flange of the moving blade,
the integrally formed axial flange defining a third axial extent and extending axially to a position radially across from
the axial flange extending from the moving blade such that a second portion of the third axial extent of the integrally formed
axial flange axially overlaps at least some of the second axial extent of the axial flange of the moving blade on a radially
inward side of the axial flange of the moving blade,

which together with a radial inner platform of a guide vane of a gas turbine, forms a fish mouth seal accommodating the axial
flange extending from the adjoining moving blade radially between the radial inner platform and the integrally formed axial
flange such that the radial inner platform axially overlaps at least the first portion of the axial flange of the moving blade
on the radial outward side and the integrally formed axial flange axially overlaps at least the second portion of the axial
flange of the moving blade on the radial inward side.

US Pat. No. 9,840,916

TURBOMACHINE BLADE

MTU Aero Engines AG, Mun...

1. A blade for a turbomachine, comprising:
a turbine blade having a channel extending over an entire height of the turbine blade in a channel longitudinal direction,
an impact chamber having a constricted cross section being situated in the channel and the turbine blade having solely a single
impulse body in the impact chamber, a channel height, running in a thickness direction of the turbine blade, being reduced
with respect to an upstream and downstream channel section in the channel longitudinal direction of the impact chamber, the
impact chamber being situated in an area between 10 percent and 90 percent of the blade height, measured from a blade root.

US Pat. No. 9,671,372

METHOD AND DEVICE FOR ASCERTAINING AN EDGE LAYER CHARACTERISTIC OF A COMPONENT

MTU AERO ENGINES AG, Mun...

1. A method for ascertaining an edge layer characteristic of a component (12) for an aircraft engine, comprising the steps of:
disposing a reference object (22) with a known edge layer characteristic on a surface of the component (12);

introducing at least one ultrasonic wave (18) into the surfaces of the component (12) and the reference object (22) by an ultrasonic transmitter (16);

detecting at least one ultrasonic wave (18) resulting from an interaction with the component (12) and the reference object (22) by an ultrasonic detector (20); and

ascertaining an edge layer characteristic of the component (12) by an ascertaining device (28), based on a difference between the at least one ultrasonic wave (18) that is introduced and the at least one ultrasonic wave (18) that is detected and interference of the at least one ultrasonic wave with the reference object (22).

US Pat. No. 9,771,830

HOUSING SECTION OF A TURBINE ENGINE COMPRESSOR STAGE OR TURBINE ENGINE TURBINE STAGE

MTU Aero Engines AG, Mun...

3. The housing section as recited in claim 1 wherein, in a cross section, the webs form a grid or a honeycomb structure.

US Pat. No. 9,689,270

DUPLEX-PHASE CRAL COATING FOR IMPROVED CORROSION/OXIDATION PROTECTION

MTU AERO ENGINES AG, Mun...

1. A process for producing a coating for protecting a component against high temperatures and aggressive media, the component
being formed by an alloy having one or more metallic main constituents which make up the largest proportion of the alloy,
wherein the process comprises chromizing a surface to be coated and subsequently aluminizing a chromium-rich layer produced
during chromizing, the chromizing being carried out with a chemical chromium activity of at least 0.4, and wherein the process
affords a coating that has an outer zone and an inner zone, the outer zone comprising ?-chromium phases in a matrix of a mixture
of mixed crystals comprising essentially chromium, aluminum, and the one or more metallic main constituents of the alloy,
and the inner zone comprising a mixed crystal zone comprising essentially chromium, aluminum, and the one or more metallic
main constituents of the alloy, the proportion of chromium in a total coating being greater than 30% by weight and a proportion
of aluminum in a total coating being at least 5% by weight, and wherein at least one of:
(i) a proportion of chromium in the outer zone is from 30% by weight to 95% by weight of chromium;
(ii) a proportion of chromium in the ?-chromium phases is at least 70% by weight;
(iii) a proportion of aluminum in the outer zone is from 10% to 40% by weight of aluminum;
(iv) the one or more metallic main constituents in the outer zone are present in a proportion of not higher than 40% by weight;
(v) in the inner zone a proportion of chromium is not higher than 30% by weight, a proportion of aluminum is not higher than
30% by weight, and a proportion of the one or more main constituents is at least 30% by weight;

(vi) a proportion of chromium in the total coating is from greater than 30% by weight to 90% by weight;
(vii) a proportion of aluminum in the total coating is from 10% to 40% by weight;
(viii) the outer zone of the coating makes up a proportion of at least 50% of the total coating;
(ix) the coating has up to 10% by volume of pores having average diameters of from 2 ?m to 20 ?m;
(x) the coating comprises from 1% to 15% by weight of oxides;
(xi) the one or more metallic main constituents of the alloy are one or more of nickel, iron and cobalt;
(xii) the chromizing is carried out using a Cr-rich slip containing a liquid phase.

US Pat. No. 9,617,863

GAS TURBINE STAGE

MTU Aero Engines AG, Mun...

1. A gas turbine stage comprising:
a rotor blade array having a plurality of rotor blades; and
an adjacent stator vane array having a plurality of stator vanes having leading edges facing the rotor blade array,
wherein in a first radial position of a rear face of the rotor blade array, a minimum axial gap between the rotor blade array
and the stator vane array is formed between said rear face and an opposite first contact region of at least one of the stator
vane leading edges, and
in a second radial position of the rear face different from the first position, the minimum axial gap is formed between the
rear face and an opposite second contact region of the stator vane leading edge, and wherein between the first and second
contact regions, the stator vane leading edge has an axial offset of no more than 0.6% of a radial height of the stator vane
leading edge.

US Pat. No. 9,822,706

GAS TURBINE SUBASSEMBLY

MTU AERO ENGINES AG, Mun...

1. A subassembly for a gas turbine, comprising:
a turbine casing;
a midframe, which is adjacent downstream to the turbine casing and has a plurality of support ribs spaced apart in the peripheral
direction;

wherein the turbine casing and the midframe define a flow duct for a working gas exiting a combustion chamber of the gas turbine,
and wherein a cooling air duct, with an opening on the flow duct side, is formed between the turbine casing and the midframe;

wherein an edge contour of the opening on the turbine casing side varies along the periphery radially and/or axially; and
wherein the edge contour is formed on a separate, two-arm, flange joined to the turbine casing.

US Pat. No. 9,605,541

BLADED ROTOR FOR A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A bladed rotor for a turbomachine, comprising
a plurality of blades distributed around a circumference of the bladed rotor; wherein a blade distribution of the plurality
of blades is a natural-number sector count (Z) of first angular sectors and second angular sectors; a first blade count of
first blades being arranged in the first angular sectors, and a second blade count of second blades being arranged in the
second angular sectors, said second blade count being different from said first blade count,

each of the first and second angular sectors extending around an angle equal to 360 degrees divided by 2Z and where each first
angular sector adjoining on both sides with the second angular sectors, and each second angular sector adjoining on both sides
with the first angular sectors;

wherein the natural number sector count is odd (Z=2n+1, n=1, 2, 3, . . . ) and/or greater than two (Z>2).

US Pat. No. 9,597,799

METHOD AND DEVICE FOR MACHINING ROBOT-GUIDED COMPONENTS

MTU Aero Engines AG, Mun...

1. A method for machining a robot-guided component, comprising:
detecting a change in position of a tool with respect to a tool holder from a desired position wherein the tool is fastened
in an articulated manner to the tool holder via an extension arm; and

changing a pose of a robot that is guiding the robot-guided component on a basis of the detected change in position;
wherein the changing the pose of the robot guides the robot-guided component in a first direction against the tool holder
to compensate for the detected change in position in a first degree of freedom and guides the robot-guided component in a
second direction against the tool holder to compensate for the detected change in position in a second degree of freedom.

US Pat. No. 10,043,257

METHOD AND DEVICE FOR THE QUALITY EVALUATION OF A COMPONENT PRODUCED BY MEANS OF AN ADDITIVE MANUFACTURING METHOD

MTU Aero Engines AG, Mun...

1. A method for the quality evaluation of a component manufactured by an additive manufacturing method, comprising:providing a component that was produced by an additive manufacturing method; the component having at least one component site;
providing an optical imaging device;
capturing image data from the at least one component site by the optical imaging device; the image data characterizing the at least one component site of the component produced by an additive manufacturing method;
converting the image data into a binary image;
providing a structure mask having dimensions less than the binary image;
eroding the binary image into a structure image using the structure mask, wherein the erosion comprises a pixel-by-pixel displacement of the structure mask over the binary image and determining whether the structure mask fully fits at a respective site of the binary image, and including the eroded data of the binary image if the structure mask fully fits and excluding the eroded data of the binary image if the structure mask does not fully fit;
determining contour data of the structure image;
determining at least one image section of the image data that is delimited by the contour data;
inspecting the at least one image section for the presence of an image region corresponding to a quality defect; and
classifying the component as being qualitatively OK if no quality defect is present, or classifying the component as being qualitatively not OK if a quality defect is present.

US Pat. No. 9,957,833

TURBOMACHINE STAGE AND METHOD FOR DETERMINING A SEAL GAP AND/OR AN AXIAL POSITION OF SUCH A TURBOMACHINE STAGE

MTU Aero Engines AG, Mun...

1. A turbomachine stage, comprising:a housing having an external surface;
a moving vane arrangement disposed within the housing, wherein the moving vane arrangement disposed along an axial rotation axis and including at least one radially disposed vane disposed along a radial axis transverse to the axial rotation axis, the at least one vane having an exterior shroud band section including a sealing flange, the sealing flange having a recess arrangement including a radial recess and a radial projection, that together, define a peripheral surface of the sealing flange; and
a sensor arrangement including a sensor having a sensing surface, the sensor arrangement being arranged on the housing so that the sensing surface of the sensor positionally converges towards or positionally diverges away in a direction relative to the radial axis in a direction that is different from the direction of the radial axis, and the converging or the diverging sensing surface of the sensor and the peripheral surface of the sealing flange, in combination, further define a radial clearance therebetween, and a radial distance of the radial clearance is detectable by the positionally converged or the positionally diverged sensor.
US Pat. No. 9,932,661

PROCESS FOR PRODUCING A HIGH-TEMPERATURE PROTECTIVE COATING

MTU AERO ENGINES AG, Mun...

1. A process for producing a high-temperature protective coating for metallic components, wherein the process comprises:producing a slip comprising (i) a powder of Cr metal and (ii) a powder of MCrAlY alloy, in which M represents at least one metal;
applying the slip to the component to be coated; and
alitizing the component provided with the slip.

US Pat. No. 9,657,576

AIRFOIL HAVING A PROFILED TRAILING EDGE FOR A FLUID FLOW MACHINE, BLADE, AND INTEGRALLY BLADE ROTOR

MTU Aero Engines AG, Mun...

1. An airfoil for a fluid flow machine, comprising:
a suction side;
a pressure side;
an airfoil trailing edge; and
a profile in at least part of a region of the airfoil trailing edge, the profile extending over the suction side and the pressure
side of the airfoil trailing edge;

wherein the profile has depressions, the depressions being wedge-shaped with a flat base and having a taper in a direction
opposite to the direction of flow.

US Pat. No. 9,617,866

BLADE OUTER AIR SEAL FOR A GAS TURBINE ENGINE

United Technologies Corpo...

1. A gas turbine engine, comprising:
a casing;
a blade outer air seal (BOAS) attached to said casing and including a seal body including a radially inner face and a radially
outer face that axially extend between a leading edge portion and a trailing edge portion, a trough at said radially inner
face, and an abradable seal received within said trough, said trough configured to expose a leading-most edge of said abradable
seal; and

a thermal barrier coating disposed on a first surface and a second surface of said casing, a portion of said first surface
being axially and radially displaced from a portion of said second surface such that said thermal barrier coating of said
first surface axially overlaps said thermal barrier coating of said second surface.

US Pat. No. 9,605,549

STATIONARY BLADE RING, ASSEMBLY METHOD AND TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A stationary blade ring for a turbomachine comprising:
a plurality of rotatable stationary blades;
an inner ring for radial inner stabilization of the stationary blades;
a seal for sealing a radial clearance between the inner ring and a diametrically opposed rotor section; and
bearing bodies for supporting the stationary blades in the inner ring, each stationary blade being supported via a bearing
journal in a bearing hole of an individual bearing body of the bearing bodies, each bearing body having a cuboid main body
in which the bearing hole is formed;

the inner ring being composed of two half rings supporting the bearing bodies, each of the half rings spanning a circumferential
angle of 180 degrees, and each having a U-shaped profile including, with respect to a flow direction of a main flow, a front
wall, a rear wall and a circumferential wall connecting the front and rear walls.

US Pat. No. 10,094,230

BRUSH SEAL SYSTEM FOR SEALING A CLEARANCE BETWEEN COMPONENTS OF A TURBO ENGINE THAT ARE MOVABLE IN RELATION TO ONE ANOTHER

MTU Aero Engines AG, Mun...

1. A brush seal system for sealing a clearance between components of a turbo engine that are movable in relation to one another, comprising:a brush seal with a brush head; and
a brush seal housing, wherein the brush head of the brush seal is accommodated in the brush seal housing, wherein the brush seal housing includes a first component with a cover plate section on a first end of the first component and a second component with a support plate section on a first end of the second component;
wherein the first component has a portion that is disposed radially over the brush seal, has a radial flange that extends radially outward from the portion, and has an axial flange forming a fish mouth seal integrally formed in the first component at a radially outermost end of the radial flange, wherein the first end of the first component is opposite from the radially outermost end of the radial flange.

US Pat. No. 10,070,069

METHOD AND DEVICE FOR DETERMINING A CONTOUR, AT LEAST IN REGIONS, OF AT LEAST ONE ADDITIVELY MANUFACTURED COMPONENT LAYER

MTU AERO ENGINES AG, Mun...

1. A method for the determination, at least in regions, of a contour of at least one additively manufactured component layer, comprising the steps of:creating a component layer by additive manufacturing;
traveling over a contour line of the component layer, at least in regions, by a laser beam;
producing a time exposure by capturing a single image of the laser beam travel over the contour line by a camera system; and
evaluating the quality of the contour line of the component layer from the single image of the laser beam,
wherein a mean speed with which the laser beam travels over the contour line is adjusted to a higher value than a mean speed that was used for the additive manufacture of the component layer, and/or in that the speed of the laser beam is varied one or more times during its travel over the contour line, and
wherein the mean speed with which the laser beam travels over the contour line is adjusted to a value between 900 mm/s and 3000 m/s and/or to a value that corresponds to at least 1.1 times the speed that was used for the additive manufacture of the component layer.

US Pat. No. 10,060,275

TURBOMACHINE BLADE ARRANGEMENT WITH FIRST AND SECOND GUIDES WITH RESPECTIVE MOVABLE FIRST AND SECOND ELEMENTS TO REDUCE VIBRATIONAL RESPONSE

MTU AERO ENGINES AG, Mun...

1. A turbomachine blade arrangement having:a first blade having a base element, which has a blade part for flow diversion and a blade root;
a first guide, fixed on the base element of the first blade, in which a first element is movably guided; and
a second guide, fixed on the base element of the first blade, in which a second element is movably guided;
wherein a dynamic of the first element in the first guide and a dynamic of the second element in the second guide of the first blade are designed differently;
a second blade adjacent to the first blade, the second blade having a base element, which has a blade part for flow diversion and a blade root;
a first guide, fixed on the base element of the second blade, in which a first element is movably guided, the first guide of the second blade arranged substantially perpendicular to the first guide of the first blade;
a second guide, fixed on the base element of the second blade, in which a second element is movably guided, the second guide of the second blade arranged substantially perpendicular to the second guide of the first blade;
wherein a dynamic of the first element in the first guide and a dynamic of the second element in the second guide of the second blade are designed differently;
wherein
the first guide of the first blade and the first guide of the second blade are arranged in a half, nearer to the blade root, of a radial height (H) of the base element, respectively, and the second guide of the first blade and the second guide of the second blade are arranged in a half, more remote from the blade root, of the radial height of the base element, respectively.

US Pat. No. 10,060,278

GUIDE VANE FOR A TURBOMACHINE HAVING A SEALING DEVICE; STATOR, AS WELL AS TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A rotatable guide vane for a turbomachine, the guide vane comprising:a sealing device at a radially inner end region of the guide vane for sealing leakage flows between the rotatable guide vane and an inner ring joined to the guide vane, the guide vane being rotatable in the inner ring;
the sealing device being movably configured relative to the guide vane, and the sealing device positionable in at least one open or in a closed configuration for sealing the leakage flows, the sealing device being movable between the open configuration and the closed configuration orthogonally to a longitudinal axis of the guide vane, wherein in the open configuration a cross section for the leakage flows is not or at least not completely sealed by the sealing device and in the closed configuration the cross section for the leakage flows is at least reduced by the dealing device.

US Pat. No. 10,035,223

REPAIR METHOD FOR THE ADDITIVE REPAIR OF A COMPONENT

MTU AERO ENGINES AG, Mun...

1. A repair method for a component, comprising the steps of:removing a damaged region of the component with the formation of at least one separating surface;
arranging the component in a processing chamber of a device for the additive restoration of at least the region of the component that has been removed;
determining first structural data of the component disposed in the processing chamber by a measurement system of the device, wherein the first structural data characterize an actual geometry of the component;
providing second structural data of the component by a computing device of the device, wherein the second structural data characterize a target geometry of the component;
determining third structural data based on the first and the second structural data by the computing device, wherein the third structural data characterize a target geometry of the region of the component that has been removed; and
additively restoring the component region that has been removed on the at least one separating surface of the component based on the third structural data by a construction apparatus of the device.
US Pat. No. 9,850,765

RHENIUM-FREE OR RHENIUM-REDUCED NICKEL-BASE SUPERALLOY

MTU Aero Engines AG, Mun...

1. A nickel-base superalloy comprising aluminum in an amount of 4% to 6% by weight, cobalt in an amount of 8% to 10% by weight,
chromium in an amount of 5% to 8% by weight, molybdenum in an amount of 2% to 5.5% by weight, tantalum in an amount of 4%
to 8% by weight, rhenium in an amount of 0% to 2% by weight, titanium in an amount of 1.5% to 5.5% by weight, and tungsten
in an amount of 3.5% to 8% by weight, as well as a remainder of nickel and unavoidable impurities.

US Pat. No. 9,784,131

SEALING ARRANGEMENT FOR A TURBOMACHINE, A GUIDE VANE ARRANGEMENT, AND A TURBOMACHINE WITH SUCH A SEALING ARRANGEMENT

MTU AERO ENGINES AG, Mun...

1. A sealing arrangement (15) for a guide vane ring (60) of a turbomachine (11),
wherein the sealing arrangement (15) comprises a thin-walled annular structure (80) that is substantially closed on all sides, and

wherein the annular structure (80) delimits an annular interior space (105),

wherein
a hollow cell structure (109), which is designed to mechanically support the annular structure (80), is provided in the annular interior space (105),

wherein the annular structure (80) comprises at least one passage opening (115, 135), which connects the annular interior space (105) to the surroundings of the sealing arrangement (15), wherein the at least one passage opening (115, 135) is designed for pressure equalization of the annular interior space (105) with the surroundings,

wherein at least one other passage opening (135) is provided for pressure equalization of the annular interior space (105), wherein the passage opening (115) and the other passage opening (135) are arranged on the same side (120) of the annular structure (80).

US Pat. No. 9,500,082

BLADE RING SEGMENT HAVING AN ANNULAR SPACE DELIMITING SURFACE HAVING A WAVY HEIGHT PROFILE

MTU Aero Engines AG, Mun...

1. A blade ring segment for a blade ring of a turbomachine comprising:
at least one shroud, the at least one shroud delimiting an annular space with a side of the shroud facing a blade profile
of a blade of the blade ring segment radially and defining at least one annular space delimiting surface on a pressure side
or an intake side of the blade profile, the at least one annular space delimiting surface being designed with a wavy height
profile, wherein the wavy height profile extends from a Z-shaped end edge of the at least one shroud, the wavy height profile
having peaks in the area of the protrusions of the Z-shaped end edge and having valleys in the area of the indentations of
the Z-shaped end edge, and the valleys and the peaks extending in the circumferential direction of the blade ring.

US Pat. No. 10,137,516

ELECTROCHEMICAL MACHINING OF A WORKPIECE

MTU AERO ENGINES AG, Mun...

1. A machine comprising a base and at least one work station which comprises a module for electrochemically machining a workpiece, wherein the module comprises:a frame; and
an electrode arrangement comprising
at least one electrode which is mechanically connected to the frame, and a drive for moving the at least one electrode, which drive is attached to the frame;
a workpiece holder for separably attaching the workpiece; and
a positioning device for displacing the workpiece holder and the module relative to each other,
and wherein
the drive comprises a drive axle at a distance from which a swivel axle is arranged, a drive arm being hinge-coupled to the swivel axle; and
an eccentric shaft is arranged on the drive axle or is integral with an output shaft of the drive, a second axle of the eccentric shaft representing the swivel axel.

US Pat. No. 10,125,627

METHOD FOR DISASSEMBLY AND ASSEMBLY OF A ROTOR OF A GAS TURBINE

MTU Aero Engines AG, Mun...

1. A method for disassembly of a rotor of a gas turbine, wherein the rotor is disposed within a housing and a channel which diverges in a direction of flow, comprising the steps of:axially displacing an outer sealing ring that is radially opposite the rotor against the direction of flow, wherein a minimum inside diameter of the outer sealing ring is smaller than a maximum outside diameter of the rotor; and
following the step of axially displacing the outer sealing ring, axially displacing the rotor against the direction of flow out of the housing.

US Pat. No. 10,112,236

DEVICE AND METHOD FOR THE MANUFACTURE OR REPAIR OF A THREE-DIMENSIONAL OBJECT

MTU AERO ENGINES AG, Mun...

1. A device for the manufacture or repair of a three-dimensional object, comprising:at least one construction chamber for successive solidification of at least one solidifiable material layer by layer in predefined regions for the layer-by-layer buildup of the three-dimensional object or for layer-by-layer repair of individual regions of the three-dimensional object within the construction chamber, and
at least one inlet nozzle and at least one suction nozzle for a process gas, with the inlet nozzle and the suction nozzle being arranged where a gas flow that passes at least partially over a buildup and joining zone of a construction platform formed in the construction chamber is created,
wherein the ratio of the sums of the fluid-dynamically relevant cross-sectional areas of the at least one suction nozzle to the at least one inlet nozzle is 2.5:1 to 0.3:1, and
wherein the at least one suction nozzle comprises at least two suction orifices, wherein said at least two suction orifices are each connected to an associated suction channel in such a way as to carry the flow, and each suction channel has wall surfaces that have a contoured course that is curved at least in sections in the flow direction and the associated suction channels are separated from one another.

US Pat. No. 10,024,192

CERAMIC COMPONENT FOR A TURBOMACHINE

1. A sealing section for a turbomachine comprising:a brush seal having brushes for contacting a shaft;
a holding device holding the brush seal, the holding device extending radially to back brushes of the brush seal; and
a ceramic component contacting the holding device and extending radially beyond the holding device, a section of the ceramic component designed to be destroyed in response to contact with the shaft;
wherein the section is a U-shaped section and wherein legs of the U-shaped section extend axially parallel to an axis of rotation.

US Pat. No. 9,920,411

DEVICE AND METHOD FOR PARTIALLY MASKING A COMPONENT ZONE OF A COMPONENT

MTU Aero Engines AG, Mun...

1. A device for partially masking a component or assembly located in a coating vacuum chamber during a coating process, the
device comprising:
at least one masking plate for separating a region to be coated from a region to be masked, the masking plate having at least
a portion of at least one opening conforming to a contour of the component and is adapted to allow passage of the component
therethrough;

a gap being present along the opening; and
at least one seal element being disposed at the gap on a side of the masked region, the seal element sealing the opening with
respect to the coating vacuum chamber when a vacuum is present in the coating vacuum chamber;

wherein the seal element is hollow.

US Pat. No. 9,835,040

TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A turbomachine comprising:
an annular space seal for the purpose of at least reducing a fluid exchange between an annular space, a main stream flowing
through the annular space, and at least one cavity situated radially on an inside or radially on an outside of the annular
space, the annular space seal having a plurality of sheet-shaped elastic elements, the elastic elements extending over shroud
edges of a guide blade row in an axial direction of the turbomachine, the elastic elements being oriented in the radial direction
of the turbomachine and in a flow direction of the main stream flowing through the guide blade row;

wherein the plurality of elastic elements are oriented axially and spaced apart circumferentially to define channels extending
in the flow direction; and

wherein the channels are open in the radial direction.

US Pat. No. 9,822,657

GAS TURBINE

MTU AERO ENGINES AG, Mun...

1. An aircraft-engine gas turbine, having a housing that has a flow channel inlet, a first array of rotor blades in a direction
of through-flow arranged in the housing, an outer sealing ring for sealing the first array of rotor blades attached to the
housing by a clamp in a friction fit, and a plurality of ring segments; wherein
a free axial path length of each of the plurality of ring segments counter to the direction of through-flow is at least as
large as

a quotient of a clearance sum
of the outer sealing ring attached to the housing in a friction fit and pi (?) is at least as large as a difference between
a maximum outer diameter of the outer sealing ring, attached to the housing in a friction fit, and a minimum inner diameter
of the flow channel inlet of the housing

US Pat. No. 9,822,462

OPTICAL MEASUREMENT SYSTEM FOR DETERMINING THE POSITION OF AN ELECTRODE DURING THE ELECTROCHEMICAL PROCESSING OF A COMPONENT

MTU AERO ENGINES AG, Mun...

1. A method for the electrochemical processing of a component, in which an electrode is moved during the electrochemical processing,
wherein the method comprises detecting a position of the electrode by an optical measurement system comprising a single light
source, a detection instrument, and deviation optics which deviate light from the light source with positional accuracy onto
the detection instrument, the light source and the detection instrument being arranged on at least one stationary carrier
and/or the deviating optics being arranged on an electrode holder or on the electrode.

US Pat. No. 9,822,797

TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A turbomachine (1), comprising:
a cooling device for supplying cooling air from an air distribution chamber (6) between an outlet diffuser (4) of a compressor (2) and at least one combustion chamber onto a compressor region, wherein the cooling device has cooling air pipes (68) for conveying cooling air through the cooling air pipes (68) from the air distribution chamber (6) into a cavity (36) between the outlet diffuser (4) and a rotor segment (11, 12) of the compressor (2) through an intervening chamber, wherein the cooling air pipes (68) have a first end connected to a wall delimiting a portion of the air distribution chamber (6) and a second end connected to a radial wall proximate to the cavity (36).

US Pat. No. 9,777,588

BRUSH SEAL SYSTEM FOR SEALING A GAP BETWEEN COMPONENTS OF A THERMAL GAS TURBINE THAT MAY BE MOVED RELATIVE TO ONE ANOTHER

MTU AERO ENGINES AG, Mun...

1. A brush seal system for sealing a gap between relatively movable components of a thermal gas turbine, the sealing system
comprising:
a brush seal having a brush head and a brush packet, the brush packet projecting from the brush head in a first direction
toward a brush packet end area disposed at an end of the brush seal opposing the brush head;

a brush seal housing connected to a first component of the thermal gas turbine, the brush seal housing receiving the brush
head of the brush seal so the brush head is not movable relative to the brush seal housing and so the brush packet projects
in the first direction toward a second component of the thermal gas turbine that is movable relative to the first component
across a gap between the relatively movable first and second components; and

a support element connected to the first component so the support element is not movable relative to the brush seal housing,
the support element projecting in the first direction and ending at a support element end area, the support element contacting
the brush packet projecting from the brush head to support the brush packet against flexing, but the support element not contacting
the brush head;

wherein the support element and the brush seal housing are embodied as separate components.

US Pat. No. 9,733,266

LOW NOISE COMPRESSOR AND TURBINE FOR GEARED TURBOFAN ENGINE

UNITED TECHNOLOGIES CORPO...

1. A gas turbine engine comprising:
a fan;
a turbine section having a first turbine including a first turbine rotor;
a compressor rotor;
a gear reduction configured to effect a reduction in a speed of the fan relative to an input speed from the first turbine
rotor; and

wherein each of the compressor rotor and the first turbine rotor includes a number of blades in each of a plurality of blade
rows, the number of blades configured to operate at least some of the time at a rotational speed, and the number of blades
and the rotational speed being such that the following formula holds true for at least a majority of the blade rows of the
first turbine rotor, but does not hold true for any of the blade rows of the compressor rotor:

(number of blades×rotational speed)/60?5500, and
the rotational speed being an approach speed in revolutions per minute.

US Pat. No. 10,013,524

METHOD FOR DESIGNING A TURBINE WITH AN IMPROVED VANE-TO-BLADE RATIO IN THE LAST STAGE OF THE TURBINE

MTU AERO ENGINES AG, Mun...

1. A method for designing a turbine of a gas turbine aircraft engine, which has a last stage with a rotating last rotor grid, having a plurality of rotating blades, and an adjacent upstream, stationary exit guide grid, having a plurality of guide vanes; with the last stage being characterized by a vane-to-blade ratio characteristic quantity, which indicates the ratio of the number of guide vanes to the number of rotating blades;wherein the last stage is configured and arranged where the vane-to-blade ratio characteristic quantity lies, in a predetermined operating condition of the turbine, above an upper cut-off limit (ok=?1) for the mode k=?1 or between a lower cut-off limit (uk=?1) for the mode k=?1 and an upper cut-off limit (ok=?2) for the mode k=?2 or between a lower cut-off limit (uk=?2) for the mode k=?2 and an upper cut-off limit (ok=?3) for the mode k=?3 of a blade-passing frequency of the last stage, with which its rotating blades rotate past one of it guide vanes;
that, in the predetermined operating condition, a flow of exhaust gas passing through the last rotor grid is reduced to a minimum flow cross section, which is at most 80% of its minimum flow cross section in the last rotor grid, in front of the exit guide grid, in a region between the last rotor grid and the exit guide grid, in the exit guide grid, and/or after the exit guide grid, in a region with an axial length that corresponds to an axial length of the exit guide grid; and
manufacturing the turbine using the results of the design steps.

US Pat. No. 10,001,083

TURBOFAN AIRCRAFT ENGINE

MTU Aero Engines AG, Mun...

1. A turbofan aircraft engine comprising:a primary duct including a combustion chamber;
a first turbine disposed downstream of the combustion chamber;
a compressor disposed upstream of the combustion chamber and coupled to the first turbine; and
a second turbine disposed downstream of the first turbine and coupled via a speed reduction mechanism to a fan for feeding a secondary duct of the turbofan aircraft engine,
a square of a ratio of a maximum blade diameter of the fan to a maximum blade diameter of the second turbine being at least 3.5;
wherein a sum of (i) a product of a square of the maximum blade diameter of the fan in [m2] and 0.15 m, (ii) a product of the maximum blade diameter of fan in [m] and ?0.28 m2, and (iii) 0.2 m3 is, in absolute value, equal to at least a volume of the primary duct bounded by an outer wall of the second turbine between the entrance cross section and the exit cross section in [m3] thereof.

US Pat. No. 9,821,368

METHOD FOR PRODUCING A CYLINDRICAL COMPONENT

MTU AERO ENGINES AG, Mun...

1. A method for producing a cylindrical component for turbomachines, comprising the steps of:
(a) providing a heated blank;
(b) press forming the heated blank to form a flat blank;
(c) bending or rolling the flat blank into a cylindrical half-shell;
(d) forming at least two half-cylindrical-shells with steps (a) through (c); and
(e) configuring and arranging the at least two half-cylindrical-shells formed in step (d) into a housing.

US Pat. No. 9,790,813

TWIST PREVENTION FOR TURBOMACHINERY

MTU Aero Engines AG, Mun...

1. A twist prevention system for a turbomachine, comprising:
a first ring segment pair with a first radially inner ring segment and a first radially outer ring segment; and
a second ring segment pair with a second radially inner ring segment and a second radially outer ring segment;
wherein the second ring segment pair is offset in a radial direction from the first ring segment pair such that the first
radially inner ring segment of the first ring segment pair is at a same radial distance from a central axis of the turbomachine
as the second radially outer ring segment of the second ring segment pair.

US Pat. No. 9,696,142

METHOD AND APPARATUS FOR DETERMINING RESIDUAL STRESSES OF A COMPONENT

MTU Aero Engines AG, Mun...

1. A method for determining residual stresses of a component while being manufactured by an additive manufacturing process,
the method comprising the steps of:
creating at least one local melt pool via laser energy in a surface of the component being manufactured after a predetermined
portion of the component is completed;

optically detecting surface distortions or elongations occurring at least in a region around the created melt pool, wherein
the optical detection of the surface distortions or elongations is carried out using an optical distortion-based method, the
optical distortion-based method being a speckle interferometry method; and

determining residual stresses of the component present at least in the region around the created melt pool based on the optically
detected surface distortions or elongations.

US Pat. No. 9,636,769

IRRADIATION IN GENERATIVE FABRICATION

MTU Aero Engines AG, Mun...

1. A method for the generative production of components for turbomachines, in which the component is constructed in layers
on a substrate or a previously produced part of the component (3), wherein a construction in layers results by melting of powder material in layers with a high-energy beam (13) and solidification of the powder melt (16),
wherein
the high-energy beam has a beam cross section (19) in the area of its impingement on the powder material that is altered in comparison to a circular or other symmetrical cross
section and/or the beam energy is distributed, over the beam cross section, in a shape that is selected from the group consisting
of non-uniform, asymmetrical and eccentric, and that

the powder material, after melting has occurred, is irradiated a second time, at a time delay to the melting, by a high-energy
beam with an energy input into the powder material that is altered in comparison to the melting, and

wherein the solidification of the powder melt occurs epitaxially.

US Pat. No. 9,664,065

CLAMPING RING FOR A TURBOMACHINE

MTU Aero Engines AG, Mun...

12. A turbomachine comprising:
a flow channel;
a housing radially surrounding the flow channel;
at least one liner being situated between the flow channel and the housing;
a stator shroud spaced apart from the at least one liner in a radial direction;
a rotor shroud spaced apart from the at least one liner in the radial direction; and
at least one clamping ring situated at least partially circumferentially around the flow channel to abut at least one contact
surface of the at least one liner;

wherein the clamping ring includes an outer ring and an inner ring spaced a distance apart in the radial direction;
wherein the outer ring and the inner ring are connected to each other via an annular web situated on a same side of front
edges of the outer and inner rings to form a cross-sectional C-shape;

the inner ring being freely supported by the outer ring via the annular web to permit the inner ring to move relatively toward
to the outer ring;

wherein the inner ring and the at least one liner separate the flow channel from the housing.

US Pat. No. 9,617,865

GUIDE VANE FOR A TURBOMACHINE, GUIDE VANE CASCADE, AND METHOD FOR MANUFACTURING A GUIDE VANE OR A GUIDE VANE CASCADE

MTU Aero Engines AG, Mun...

1. A guide vane for a turbomachine comprising:
an upper trailing edge and a lower trailing edge, the guide vane being axially pivotably coupled to a radially outwardly disposed
flow-limiting wall and to a radially inwardly disposed inner ring of the turbomachine, a trailing edge gap being formed between
the upper trailing edge of the guide vane and the flow-limiting wall or between the lower trailing edge of the guide vane
and the inner ring,

wherein the upper trailing edge or the lower trailing edge of the guide vane has at least one air outlet opening for an air
outflow for forming an air curtain for at least partially sealing the trailing edge gap in the area of the upper trailing
edge or the trailing edge gap in the area of the lower trailing edge;

wherein the guide vane has a blade surface, the air outlet opening communicating air-conductively via at least one channel
with at least one air inlet opening on the blade surface; and

wherein the air inlet opening is configured on a pressure side of the guide vane.

US Pat. No. 9,512,734

SEALING OF THE FLOW CHANNEL OF A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A turbomachine having a flow channel and comprising:
a housing radially surrounding the flow channel having at least one housing hook;
a plurality of stationary and moving blades situated in the flow channel, the stationary blades being situated adjacent to
the moving blades in an axial direction, and the stationary blades having at least one stationary blade hook engaging with
the housing hook for the purpose of connecting the stationary blades to the housing;

at least one liner situated in a radial direction between the moving blades adjacent to the stationary blades and the housing,
the liner having a seal interacting with blade tips of the moving blades;

at least one heat protector being in an area of the liner between the liner and the housing; and
a further seal situated on the heat protector to create sealing contact with the housing hook,
the heat protector having a higher thermal resistance than the further seal, wherein the heat protector is made of a material
having higher thermal resistance properties than the further seal.

US Pat. No. 10,066,486

METHOD FOR DESIGNING A TURBINE

MTU Aero Engines AG, Mun...

1. A method for designing a turbine having a plurality of stages disposed axially one behind the other in the direction of flow through the turbine, each stage being formed of a stationary row of a plurality of stator vanes and a rotating row of a plurality of rotor blades, the row of rotor blades of at least one of the stages having a plurality of rotor blade clusters each formed of at least two of the rotor blades having different airfoil profiles or different distances from adjacent rotor blades following in the direction of rotation, the stage having a vane-to-blade-cluster ratio (V/(B/P)) parameter indicating the ratio of the number of stator vanes (V) to the quotient (B/P) of the number of rotor blades (B) of the stage divided by the number of rotor blades per rotor blade cluster (P), the method comprising:designing the stage in such a way that, under a predetermined operating condition of the turbine, the vane-to-blade-cluster ratio parameter is above an upper cut-off limit for the mode k=?1 or between a lower cut-off limit for the mode k=?1 and an upper cut-off limit for the mode k=?2, of a frequency defined by a ratio of the blade-passing frequency to the number of rotor blades per rotor blade cluster or an integral multiple of this ratio.

US Pat. No. 9,896,940

BLADE FOR A GAS TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A blade for a gas turbomachine, the blade comprising:
an airfoil for deflecting a flow of working fluid;
a first platform connected thereto to radially bound a flow duct for the working fluid,
the airfoil having a suction side and a pressure side connected at a leading edge and at a trailing edge, the trailing edge
having a first minimum wall thickness in a first region of a radial longitudinal extent of the airfoil proximal to the first
platform, and a maximum wall thickness smaller than the first minimum wall thickness in a platform-distal region of the radial
longitudinal extent; and

a second platform radially opposite the first platform being connected to the airfoil to radially bound the flow duct, the
trailing edge having a second minimum wall thickness greater than the maximum wall thickness in a second region of the radial
longitudinal extent proximal to the second platform,

the maximum wall thickness of the trailing edge in the platform-distal region being no greater than 0.35 mm,
wherein each of the first and second platform proximal regions extends over 10% of the radial longitudinal extent in a direction
away from a platform;

wherein the platform-distal region extends over at least 25% of the radial longitudinal extent from a middle of the radial
longitudinal extent of the airfoil toward a platform-proximal region,

wherein the wall thickness of the trailing edge is equal to a maximum wall thickness of the airfoil in a region which extends
by 5% from an axially rear end of the airfoil in the upstream direction toward the leading edge,

whereby the maximum wall thickness of the trailing edge in the platform distal region is at least substantially constant,
wherein the wall thickness of the trailing edge decreases strictly monotonically respectively, from the minimum wall thickness
of the respective platform-proximal region to the maximum wall thickness of the platform-distal region in a transition region
of the radial longitudinal extent between the first and second platform-proximal region and the platform-distal region, and

wherein the transition regions extend respectively over 10% of the radial longitudinal direction.

US Pat. No. 9,803,494

SEALING ELEMENT, SEALING UNIT, AND TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A sealing element of a turbomachine for setting a radial gap between a moving blade row and a housing area, the sealing
element comprising:
a support structure having a radially outer side and a radially inner side;
an intake lining situated on the radially inner side;
at least one retaining element situated on the radially outer side for retaining the support structure in a setpoint position;
and

a safeguard situated on the radially outer side for securing the support structure in the setpoint position during mounting
or transport,

the at least one retaining element and the safeguard being situated one behind another in a mounting direction, the safeguard
having a catch section having a mounting surface inclined by a mounting angle in the mounting direction, and having a securing
surface inclined by a safety angle in an opposite direction opposite the mounting direction, the catch section elastically
mounted via a spring section, the mounting angle being smaller than the safety angle;

wherein the spring section is a profile with an outer leg and an inner leg situated one above the other, the outer leg merging
into the catch section and the inner leg merging into a fastening section attached on the radially outer side.

US Pat. No. 9,664,141

THRUST DEFLECTING DEVICE AND AIRCRAFT ENGINE

MTU Aero Engines AG, Mun...

1. A thrust deflecting device for deflecting a thrust stream of an aircraft engine, the thrust deflecting device comprising:
a housing of the aircraft engine;
a flap system having a plurality of deflecting flaps, each deflecting flap pivotable around a yaw axis extending orthogonally
with respect to a transverse axis of the thrust deflecting device; and

two parallel baffle plates extending downstream from the housing, the flap system situated between the two parallel baffle
plates, the two parallel baffle plates together with the flap system being pivotable around a pivot axis running in the direction
of the transverse axis for the purpose of deflecting the thrust stream in a pitch direction.

US Pat. No. 9,664,057

STATOR VANE SEGMENT OF A FLUID FLOW MACHINE AND TURBINE

MTU Aero Engines AG, Mun...

1. A stator vane segment of a fluid flow machine, the stator vane segment comprising:
a downstream casing-side attachment; and
a platform, the downstream attachment being provided on the platform, the platform having a first sealing section and extending
in the direction of flow beyond a connection region connecting the downstream attachment to the platform, the first sealing
section located in a section downstream of the connection region having a first slot, the first slot being continuous in the
circumferential direction and forming a passage through said section in the circumferential direction.

US Pat. No. 9,587,499

INNER RING OF A FLUID FLOW MACHINE AND STATOR VANE ARRAY

MTU Aero Engines AG, Mun...

1. An inner ring of a fluid flow machine for attachment to stator vanes of the fluid flow machine and for receiving seal segments,
the inner ring comprising:
a fixing ring and a seal carrier, the fixing ring being split into at least a first fixing ring segment and a second ring
segment and the seal carrier being split into at least a first seal carrier segment a second seal carrier segment in the circumferential
direction of the fluid flow machine, and the first and second fixing ring segments and the first and second seal carrier segments
being arranged in the circumferential direction with respective end faces facing each other,

the first fixing ring segment having a first ring segment shoulder and the second fixing ring segment having a second ring
segment shoulder offset with respect to the first ring segment shoulder;

the first seal carrier segment having a first carrier segment shoulder and the second seal carrier segment having a second
carrier segment shoulder offset with respect to the first carrier segment shoulder; and

in a separation plane of the inner ring, the first carrier segment shoulder at least partially abutting the second ring segment
shoulder at respective end faces; and, in the separation plane, the second carrier segment shoulder at least partially abuts
the first ring segment shoulder at the respective end face.

US Pat. No. 9,500,088

BLADE RIM SEGMENT FOR A TURBOMACHINE AND METHOD FOR MANUFACTURE

MTU AERO ENGINES AG, Mun...

1. A blade rim segment for a turbomachine having a rotational axis defining an axial direction, the blade rim segment comprising:
at least one shroud band extending along a segment of a circle;
at least three blades connected to the shroud band in single-piece fashion, each respective blade extending radially from
the shroud band from a respective connection area;

wherein each respective blade is configured to be hollow with at least one respective channel formed therethrough; and
wherein each respective channel has a respective outlet opening formed in the shroud band in the respective connection area
of the respective blade;

a respective first reinforcement rib formed on the shroud band at the respective outlet opening of the respective channel
of each respective blade, each respective first reinforcement rib having a surrounding portion surrounding the respective
outlet opening, the surrounding portion having a first height parallel to the longitudinal extension of the blades above the
shroud band, and an axial portion running axially, the axial portion having a second height above the shroud band and having
a width parallel to the circumferential direction;

wherein the first height of the surrounding portion is greater than the second height of the axial portion; and
at least one second reinforcement rib formed on the shroud band between two of the first reinforcement ribs and running parallel
to the axial portions of the first reinforcement ribs;

wherein the at least one second reinforcement rib is independent from any of the respective outlet openings; and
wherein the at least one second reinforcement rib is situated outside the respective connection areas of the shroud band and
the respective blades; and

wherein the blade rim segment is formed of a single piece directionally solidified cast material having a lattice of crystallites
set in a predetermined orientation.

US Pat. No. 9,169,734

SYSTEM FOR SPECIFYING AN INSTALLATION POSITION OF ROTOR BLADES, SECURING ELEMENT, ROTOR BLADE, TURBOMACHINE, AND METHOD

MTU Aero Engines AG, Mun...

1. A system for specifying an installation position of adjacent rotor blades of a blade row of a turbomachine, the turbomachine
including a rotor shaft having a plurality of adjacent rotor grooves with identical cross sections spaced apart at a groove
distance, the system comprising:
a plurality of rotor blades of a first type, each rotor blade of the first type having a blade root positionable in one of
the rotor grooves and having a counter contour of a first type formed in a base face of the blade root such that, when the
blade root is positioned in one of the rotor grooves, the counter contour of the first type faces a groove base of the rotor
groove;

a plurality of rotor blades of a second type, each rotor blade of the second type having a blade root positionable in one
of the rotor grooves and having a counter contour of a second type formed in a base face of the blade root such that, when
the blade root is positioned in one of the rotor shaft grooves, the counter contour of the second type faces the groove base
of the rotor groove; and

a plurality of axial securing elements, each securing element having at least a first section and a second section spaced
apart a groove distance from one another and joined together by a connecting web;

the first section including a profile area of a first type that is configured to form a positive-fit pair with the counter
contour of the first type when positioned between the groove base of a rotor groove and the blade root of a rotor blade of
the first type;

the second section including a profile area of a second type that is configured to form a positive-fit pair with the counter
contour of the second type when positioned between the groove base of a rotor groove and the blade root of a rotor blade of
the second type;

whereby each axial securing element is positionble with the first section in a first rotor groove of the rotor shaft and the
second section in an adjacent rotor groove such that the profile area of the first type on the first section permits the positioning
of a rotor blade of the first type in the first rotor groove but blocks the positioning of a rotor blade of the second type
in the first rotor groove and the profile area of the second type on the second section permits the positioning of a rotor
blade of the second type in the adjacent rotor groove but blocks the positioning of a rotor blade of the first type in the
adjacent rotor groove.

US Pat. No. 10,066,493

ROTOR OF A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A rotor of a turbomachine comprising:at least one blade having a blade leaf and a blade root;
a rotor base body having an outwardly open, circumferential groove for receiving the blade root;
the circumferential groove and the blade root being shaped in a way to allow the blade root to be secured in the circumferential groove by rotation of the blade about an axis, and
a securing wire in the circumferential groove resting against a bottom of the blade root; and
a circumferential securing device having a head portion received radially inwardly in a pair of radially extending grooves in sidewalls of the circumferential groove.

US Pat. No. 10,041,353

BLADE CASCADE AND TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A blade cascade of a turbomachine comprising:a plurality of blade channels each circumferentially bounded by a pressure side of a blade and by an opposite suction side of an adjacent blade, each blade channel radially bounded by two opposing side walls, at least one of the two side walls of the blade channels being provided with a side-wall contour, the side-wall contour being circumferentially undulated and comprising at least two elevations relative to a non-contoured surface of the at least one of the two side walls and at least one depression relative to the non-contoured surface of the at least one of the two side walls;
wherein the amplitudes of the elevations differ from one another; and
wherein a plurality of depressions merge into the at least one depression.

US Pat. No. 9,765,633

BLADE CASCADE

MTU Aero Engines AG, Mun...

1. A blade cascade for a turbomachine, the blade cascade comprising:
a plurality of blades including a monocrystalline material, each blade having a crystal orientation value dependent on a crystal
orientation of the monocrystalline material of the blade; and the respective crystal orientation values of first blades of
the plurality of blades being less than a first limiting value and the respective crystal orientation values of second blades
of the plurality of blades being at least equal to the first limiting value, at least one first sector including at least
three successive first blades, and at least one second sector includes at least three successive second blades.

US Pat. No. 9,540,522

ANTIFOULING LAYER FOR COMPRESSOR BLADES

MTU AERO ENGINES AG, Mun...

1. A process for providing a coating having antifouling properties on a component of a turbomachine, wherein the process comprises
applying to the component a coating composition which comprises a binder comprising at least one silicon-organic constituent
that comprises at least one alkoxysilane, ceramic particles which comprise at least silicon dioxide, and a solvent; and subsequently
curing the coating composition by a heat treatment at a temperature of from 200° C. to 350° C. for from 10 minutes to 120
minutes.

US Pat. No. 9,522,435

PROCESS FOR PRODUCING INTERMETALLIC WEAR-RESISTANT LAYER FOR TITANIUM MATERIALS

MTU AERO ENGINES AG, Mun...

1. A process for producing a wear-resistant layer on a component, wherein the process comprises:
providing a component which comprises a titanium material on at least part of a surface of the component on which the wear-resistant
layer is to be produced,

contacting the titanium material with a solder in the form of a paste or semifinished product formed from a cobalt base material,
soldering the solder to the titanium material by applying heat to thereby produce one or more diffusion zones between solder
and titanium material, which one or more diffusion zones comprise one or more intermetallic phases and form the wear-resistant
layer, the soldering being carried out at a temperature of from about 1100° C. to about 1200° C.

US Pat. No. 10,047,618

COMPONENT SYSTEM OF A TURBO ENGINE

MTU Aero Engines AG, Mun...

1. A component system of a turbine engine, the component system comprising:a first component segment; and
a second component segment configurable in a ring segment shape about an axis of a rotor of the turbine engine, so that a first abutment surface of the first component segment abuts against a second abutment surface of the second component segment, wherein, together, the first component segment and the second component segment include at least first, second and third overlapping elements for sealing a gap between the first and second abutment surfaces;
the first and third overlapping elements being configured as first projections in a circumferential direction with respect to the first abutment surface on the first component segment and overlapping radially the second component segment;
the second overlapping element being configured as a second projection in the circumferential direction with respect to the second abutment surface on the second component segment and overlapping radially the first component segment; and,
the second overlapping element being axially configured between the first and third overlapping elements;
wherein, the first and third overlapping elements each have a radially inward surface overlapping radially the second component segment,
wherein the second overlapping element has a radially inward surface overlapping radially the first component segment, and
wherein at least one of the three radially inward surfaces is configured at a radially different distance or in a radially different plane relative to the axis than another of the three radially inward surfaces.

US Pat. No. 10,047,619

SEAL CONFIGURATION FOR A TURBO MACHINE

MTU Aero Engines AG, Mun...

1. A seal configuration for a turbo machine having a shaft, the seal configuration being ring-shaped and comprising:a seal element to provide a seal with respect to the shaft when installed; and
a ring-shaped support device, the support device having a material at a radial inner end, the material having a melting point between 800° C. and 1100° C. such that, when installed, the material melts when coming into contact with the rotating shaft;wherein the material is a self-fluxing alloy or a nickel-base brazing alloy.
US Pat. No. 10,029,309

PRODUCTION PROCESS FOR TIAL COMPONENTS

MTU AERO ENGINES AG, Mun...

1. A process for producing a component of a TiAl alloy, wherein the process comprises:introduction of a powder of the TiAl alloy into a capsule whose shape corresponds to a shape of the component to be produced and closing of the capsule,
hot isostatic pressing of the capsule together with the powder,
heat treatment of the hot isostatically pressed capsule,
removal of the capsule,
post-working of a contour of the component by removal of material.

US Pat. No. 9,989,093

ROLLER BEARINGS, IN PARTICULAR NEEDLE BEARINGS, FOR ARRANGING ON A PIVOT PIN OF A VARIABLE GUIDE VANE OF A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A roller bearing for arranging on a pivot pin of a variable guide vane of a turbomachine, comprising:a radially elastic needle which has two opposing end regions and a roller region arranged between the two end regions, wherein the roller region has an average cross-sectional thickness that is greater than an average cross-sectional thickness of the needle; and
a bearing housing;
wherein the two end regions roll off at respective raceways of the bearing housing and wherein the roller region is disposed in a region of a groove of the bearing housing.

US Pat. No. 9,982,559

BLADE OR VANE FOR A TURBOMACHINE AND AXIAL TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A blade or vane for a turbomachine, comprising:at least one impulse element housing including
a first impact cavity, in which an impulse element is arranged with play of movement,
at least one second impact cavity, which is in alignment with the first impact cavity in a first matrix direction and in which an impulse element is arranged with play of movement,
at least one third impact cavity, which is in alignment with the first impact cavity in a second matrix direction crosswise to the first matrix direction and in which an impulse element is arranged with play of movement, and
at least one fourth impact cavity, which is in alignment with the at least one third impact cavity in a first matrix direction and in which an impulse element is arranged with play of movement and is, respectively, in alignment with the respective at least one second impact cavity in the second matrix direction, and
at least one of the impact cavities in at least one first and one second matrix direction, which enclose between them an angle of at least 30° and at most 150° and each of which encloses an angle of at least 75° and at most 105° with a longitudinal axis of the blade or vane that is perpendicular to an axis of rotation of the turbomachine, in each case, has cavity walls lying opposite each other, wherein a distance between the cavity walls lying opposite each other in the first matrix direction differs by at most 10% in comparison to a distance between the cavity walls lying opposite each other in the second matrix direction;
whereby the first impact cavity, the at least one second impact cavity, the at least one third cavity and the at least one fourth cavity are arranged in a substantially equidistant matrix-like manner.
US Pat. No. 9,952,236

METHOD AND DEVICE FOR PROCESS MONITORING

MTU AERO ENGINES AG, Mun...

1. A generative fabrication process of a component, wherein the process comprises forming the component in an installation space from a multiplicity of layers by using a three-dimensional data model and fixing a following layer to a preceding layer by means of a high-energy beam, and wherein the process further comprises (i) carrying out an optical 3-D measurement of the component at least one of after a layer has been produced and after application of a powder, and (ii) carrying out thermography after application but before sintering or fusing of a powder.

US Pat. No. 9,945,241

SEALING ELEMENT OF AN AXIAL TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A sealing element of an axial turbomachine for the sealing of regions at or in at least two static components of the turbomachine against flow media, wherein the components can move axially relative to one another due to thermal expansion,consisting essentially of:
the sealing element having an inner radius, an outer radius and segmented in a peripheral direction in at least two segments, each segment having an arc length, the arc length having a curvature about a mid-point angle, the sealing element further having a first radial region, lying radially inward at least at inner radius, with a first stiffness or rigidity and a second radial region, lying radially outward at most at outer radius with a second stiffness or rigidity, wherein the first radial region is spiral-shaped with a first distal end curved inwardly thereon, the second radial region comprising a sealing surface and a diagonal transition region between the first radial region and the sealing surface, the diagonal transition region extending in an axial and radial direction offsetting the first radial region axially from the sealing surface, wherein the first stiffness and the second stiffness are different from one another, the first stiffness being greater than the second stiffness and the first radial region is configured and arranged for attachment in a static component of the turbomachine, the sealing surface having a second distal end, the sealing surface and second distal end aligned perpendicular in an axial direction, the sealing surface further configured and arranged to form a seal against a flow media.

US Pat. No. 9,920,655

GAS TURBINE

MTU Aero Engines AG, Mun...

1. A gas turbine comprising:
a housing having a circumferential groove;
a guide vane assembly having at least one vane extending radially inwardly with respect to the housing; and
a securing ring axially contacting and locking the guide vane assembly in position with respect to the housing, the securing
ring having a radially outer rim configured in the groove, a radially inner rim configured outside of the groove, and a slot
extending from the radially outer rim to the radially inner rim, a first flank of the slot having a radially inner rim section
and an undercut portion, the radially inner rim section forming an angle of at least 50° with the radially inner rim, a second
flank of the slot having a planar surface that is opposite both the undercut portion and the radially inner rim section of
the first flank, the planar surface parallel to the undercut portion.

US Pat. No. 9,695,699

SECURING BLADE ASSORTMENT

MTU Aero Engines AG, Mun...

1. A securing plate assortment for a blade assembly of a gas turbine, comprising:
a plurality of different securing plates, wherein the different securing plates have a first flange, a second flange, and
a connecting web, and wherein the different securing plates have respective different geometric codings situated in respective
coding areas;

wherein, in each of the different securing plates, the first flange or the second flange has the respective coding area;
and wherein each respective coding area is situated on a first surface of the first or second flange facing away from the
blade assembly and/or on an edge joining the first surface with a surface facing the blade assembly; and

wherein the different geometric codings include codings on or through the first surface, or codings on the edge.
US Pat. No. 9,333,576

METHOD FOR PRODUCING AND REPAIRING A PART, AND PART OF A GAS TURBINE

MTU AERO ENGINES AG, Mun...

1. A method for producing and repairing a part comprising at least two metal components joined together, the components being
components of a gas turbine, wherein the joining and connection of the corresponding joining surfaces of the components is
produced by means of a pressure welding method, and during the joining process, a machining allowance is upset in the region
of a joining zone of the two joining surfaces, is hereby characterized in that after joining the two components, the machining
allowance is processed by means of a precise electrochemical machining (PECM) until a pre-defined final contour of the part
is obtained, wherein the machining is conducted by means of an electrode;
the step of performing precise electrochemical machining further comprising a step of disposing an electrically non-conductive
guide element on the electrode for guiding an electrolyte;

whereby the guide element is disposed on the electrode in a way that provides a targeted guiding of the electrolyte along
the part so that removed material does not enter into regions of the part that are to be machined;

wherein, in precise electrochemical machining, the electrode conducts oscillating movements when sunk down, and the step of
performing precise electrochemical machining further comprises a step of emitting a current pulse only in a region of minimum
distance between the electrode and the part.

US Pat. No. 10,126,207

MAINTENANCE OF A USED GAS TURBINE

MTU AERO ENGINES AG, Mun...

1. A method for the maintenance of a used gas turbine comprising the at least partially automated steps of:determining the geometry of a flow-guiding component, configured as rotating blade or a guide vane, of the gas turbine;
prognosticating the aerodynamics and/or thermodynamics of the component based on the determined geometry;
classifying the component into one of several predetermined classes based on the prognosticated aerodynamics and/or thermodynamics, said predetermined classes denoting different properties and parameter ranges indicating unusable components to usable components with poor performance;
virtual variation of the geometry of the component;
prognosticating the aerodynamics and/or thermodynamics of the component based on this varied geometry;
classifying the component into one of the classes; and
repairing the component based upon a repair recommendation for the component which is output, in an at least partially automated way, on the basis of the varied geometry, if the class of the component fulfills a predetermined condition, wherein the component has prognosticated aerodynamics and/or thermodynamics increased over the class of the component with the determined geometry.

US Pat. No. 10,119,408

METHOD FOR CONNECTING A TURBINE BLADE OR VANE TO A TURBINE DISC OR A TURBINE RING

MTU AERO ENGINES AG, Mun...

1. A method for connecting a turbine blade or vane to a turbine disk or a turbine ring, wherein the method comprises:(a) supplying an additive suitable for fusion welding to a surface of the turbine blade or vane,
(b) melting the additive on the surface of the turbine blade or vane, with incipient melting of the surface,
(c) allowing the additive and the surface to solidify to form a connecting body on the turbine blade or vane; and
(d) connecting the turbine blade or vane to the turbine disk or the turbine ring by directly fusion welding the connecting body to the turbine disk or the turbine ring.

US Pat. No. 10,066,668

SPLIT INNER RING

MTU AERO ENGINES AG, Mun...

1. A split inner ring for an adjustable guide blade arrangement in a turbomachine, comprising:a ring segment arrangement including a first ring segment and a second ring segment adjacent thereto;
a bearing bush arrangement with a bearing bush configured and arranged for mounting an adjustable guide blade of the adjustable guide blade arrangement, the adjustable guide blade includes a journal received in the bearing bush of the bearing bush arrangement; the bearing bush having a radial projection in communication with the first ring segment and the second ring segment with a portion of the first ring segment and a portion of the second ring segment located between the radial projection and the journal thereby maintaining the first ring segment and the second ring segment adjacent to each other;
the first ring segment and the second ring segment of the ring segment arrangement being joined together by the bearing bush arrangement wherein the at least one bearing bush is positioned between the journal of the adjustable guide blade and the first and second ring segments of the ring segment arrangement.

US Pat. No. 9,982,566

TURBOMACHINE, SEALING SEGMENT, AND GUIDE VANE SEGMENT

MTU AERO ENGINES AG, Mun...

1. A turbomachine with a sealing segment ring between a front guide vane row and a back guide vane row for sealing a radial gap between a casing section and a rotor blade row rotating between the guide vane rows, with the sealing segment ring having a plurality of identical sealing segments, and at least one of the guide vane rows having a plurality of identical guide vane segments, wherein the sealing segments each have a plurality of engagement sites comprising slots lying adjacent to one another in the peripheral direction on a front edge portion of the sealing segment for interaction with securing elements on the guide vane row, with the engagement sites and securing elements being distributed uniformly over the periphery and the engagement sites being a multiple of the securing elements, wherein each guide vane segment of the guide vane row has only one securing element.

US Pat. No. 9,964,496

METHOD FOR THE QUALITY ASSESSMENT OF A COMPONENT PRODUCED BY MEANS OF AN ADDITIVE MANUFACTURING METHOD

MTU Aero Engines AG, Mun...

1. A method for the quality assessment or the quality class of a component for a machine produced by means of an additive manufacturing method, comprising:a) preparation of a first data set, wherein the first data set comprises absolute limit values that each characterize a maximum allowed range of values at an assigned component position of a machine component being produced;
a1) providing an optical tomography device as an acquisition device;
b) acquisition of a second data set by means of the acquisition device that captures images of the machine component being produced, wherein the second data set comprises actual values corresponding to the first data set, which characterize the assigned component position of the machine component being produced;
c) comparison of the first data set and the second data set by means of a computing device and
c1) classification of the machine component as being qualitatively fundamentally not OK if at least one actual value lies outside its assigned maximum allowed range of values; or
c2) classification of the component as being qualitatively fundamentally OK if no actual value lies outside its assigned maximum allowed range of values; and
if the machine component has been classified as being qualitatively fundamentally not OK, it can be concluded that there is a serious breakdown or malfunction;
if the machine component has been classified as being qualitatively fundamentally OK:
d) preparation of a third data set, which comprises mean values that are determined from a plurality of actual values of the second data set by means of the computing device, wherein the plurality of actual values characterize an interrelated machine component region composed of a plurality of component positions;
e) determination of at least one best-fit function that is dependent on a geometry of the machine component on the basis of the third data set by means of the computing device;
f) determination of threshold values that depend on the geometry of the machine component by means of the computing device, wherein the threshold values characterize an allowed range of scatter of the actual values around target values predetermined by the best-fit function;
g) checking by means of the computing device whether at least one actual value lies outside of the range of scatter characterized by the threshold values, and
g1) if no actual value lies outside of the range of scatter, the machine component is classified as being qualitatively OK; or
g2) if at least one actual value lies outside the range of scatter, then all actual values that lie outside of the range of scatter are compiled to obtain a fifth data set, and a quality of the machine component is assessed on the basis of the fifth data set and at least one predetermined quality criterion.

US Pat. No. 9,963,993

TURBINE RING AND TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A turbine ring of a turbomachine having a one-piece ring structure, the turbine ring comprising:an outer ring continuous in a circumferential direction of the turbine ring and having a fastening section for fastening the turbine ring on a housing section of the turbomachine;
an inner ring segmented in the circumferential direction and having a plurality of inner ring segments mutually displaceable in the circumferential direction for guiding hot gas; and
front sealing segments and rear sealing segments, with respect to an axial direction of the turbine ring, the front and rear sealing segments being spaced a distance apart from one another in the axial direction and extending radially between the inner ring and the outer ring, wherein the front and rear sealing segments are in fluid-tight contact with the inner ring and the outer ring.

US Pat. No. 9,869,184

GAS TURBINE BLADE

MTU AERO ENGINES AG, Mun...

1. A blade (1) for a gas turbine, comprising:
a leading edge and a trailing edge, which are connected by a pressure side and an intake side;
wherein, in at least one segment of a stacking axis, the blade has a cross section with a profile with
a common profile tangent at a leading edge region and at a trailing edge region,
a leading edge tangent at the leading edge, which is perpendicular to the common profile tangent, the leading edge tangent
having a contact point on the profile and resting on the profile in a flow direction;

a leading edge circle having the contact point of the leading edge tangent in common, the leading edge circle having a center
point that lies at a point of intersection of a normal line at the contact point of the leading edge tangent on the profile
with a profile center line that extends through the profile at an equal distance from the pressure side and the intake side
and includes this contact point;

a trailing edge tangent at the trailing edge, which is perpendicular to the common profile tangent, the trailing edge tangent
having a contact point on the profile and resting on the profile opposite to the flow direction;

a trailing edge circle having the contact point of the trailing edge tangent in common and having a center point that lies
at a point of intersection of a normal line at the contact point of the trailing edge tangent on the profile with said profile
center line and includes this contact point of the trailing edge tangent; and

a camber line, which extends, at an equal distance from the pressure side and the intake side, from the center point of the
leading edge circle, up to the center point of the trailing edge circle;

wherein the leading edge tangent is parallel to the trailing edge tangent;
wherein the blade thickness between the pressure side and the intake side increases, extending from a trailing edge thickness
at the center point of the trailing edge circle in a camber line segment, the length of which is at least 15% and at most
25% of a length of the camber line between the center point of the leading edge circle and the center point of the trailing
edge circle, toward the center point of the leading edge circle, increasing to at least two times the trailing edge thickness.

US Pat. No. 9,835,039

SLIDE RING SEAL

MTU Aero Engines AG, Mun...

1. A housing structure for a turbomachine, the housing structure surrounding a flow channel for a fluid and comprising:
an outer housing wall and an inner wall defining the flow channel, a hollow space being formed between the inner wall and
the outer housing wall, the hollow space having an indentation, the hollow space separable into at least two regions; and

a movable wire element adapted to rest against contact faces and being configured in the hollow space for purposes of separating
the at least two regions, wherein the indentation is one of the contact faces; and

an abradable coating for blade tips provided on one side of the inner wall, the movable wire element being configured on an
opposite, facing away side of the inner wall in an area of the abradable coating;

wherein at least another one of the contact faces is a surface of the opposite facing-away side of the inner wall, or a surface
of a shield configured on the opposite facing away side of the inner wall.

US Pat. No. 9,829,130

ADAPTER AND TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. An adapter for connecting a fluid line to a housing section of a turbomachine, comprising:
a core housing with a fluid channel for establishing a fluid communication between the fluid line and a housing inlet;
a plurality of flanges extending from the core housing, wherein each of the plurality of flanges has a feed-through for passage
of a fastener;

wherein a first two of the plurality of flanges form an only single lower pair of flanges for attachment of the adapter to
the housing section and wherein a second two of the plurality of flanges form an only single upper pair of flanges for connecting
a connecting element of the fluid line to the adapter;

wherein the lower and the upper pairs of flanges are arranged offset from each other around a channel axis of the fluid channel
by 90° and are arranged without a space between each other in a direction of the channel axis;

and wherein the feed-throughs of the upper pair of flanges are formed as circumferentially open slots.

US Pat. No. 9,790,615

IDENTIFYING AND REGULATING THE STARTING BEHAVIOR DURING ELECTROCHEMICAL MACHINING OF WORKPIECES

MTU AERO ENGINES AG, Mun...

1. A method for electrochemical machining of a workpiece, wherein at least one electrode is situated adjacent to a surface
to be machined and current pulses are generated in pulsed operation to ablate material from the workpiece, and wherein before
and/or at the beginning and/or during an electrochemical ablation, data of the current pulses are registered and analyzed
to identify a starting phase or a transient phase comparable to a starting phase and/or to regulate a spacing of the electrode
to the surface to be machined and/or a current flow during the starting phase or the transient phase comparable to a starting
phase, and wherein the spacing of the electrode to the surface to be machined and/or of the current flow is regulated by setting
a feed of the electrode in a direction of the surface to be machined and/or by setting an applied potential as manipulated
variables, a regulation comprising a linearization and decoupling of the manipulated variables, the applied potential U being
defined by

and/or the feed V being defined by
V=?v2?v1,

wherein R is a total resistance of a system, x is the spacing of the electrode from the surface to be machined, ? represents
electrical conductivity, A represents a working surface area of the electrode, ?U represents an overvoltage, and v1 and v2 are virtual manipulated variables, v1 corresponding to a control variable of the spacing of the electrode from the surface to be machined and v2 corresponding to a control variable of a current strength.

US Pat. No. 9,682,437

MULTIPART ELECTRODE ARRAY AND METHOD FOR THE ELECTROCHEMICAL TREATMENT OF BLADES HAVING SHROUDING BANDS

MTU AERO ENGINES AG, Mun...

1. A method for the electrochemical machining of a blade of a turbomachine, wherein the method comprises using at least three
multipart electrodes, each comprising at least three sub-electrodes which are lined up in a plane transversely to a plane
in which the multipart electrodes are arranged, which multipart electrodes are moved with a main direction of movement toward
the blade airfoil in a star-like manner from an initial position at a first distance from a surface to be machined to an end
position at a second distance from an end contour of the surface to be machined so that in the end position a closed working
surface of the electrodes with a negative shape lies opposite the end contour of the surface to be machined, the blade having
a gas passage region and comprising shrouds at a blade root and/or at a blade tip, which shrouds are inclined on one side
or on both sides so that undercuts exist in the gas passage region.

US Pat. No. 10,066,633

GAS TURBINE COMPRESSOR BLEED CHANNEL

MTU Aero Engines AG, Mun...

1. A gas turbine compressor comprising:a guide vane;
a moving vane; and
a bleed channel having a curved upstream channel wall merging into an annular space, a curved downstream channel wall downstream axially at a distance from the upstream channel wall and having an inlet edge, and a bleed channel outlet,
the curved downstream channel wall defining, with respect to an axis of rotation of the compressor, a first angle increasing in the flow direction;
the curved upstream channel wall defining, with respect to the axis of rotation, a second angle increasing in the flow direction, the second angle increasing more than the first angle in the flow direction; and wherein:
b1?rK/5; or
0.5·[(rK+b1)2?(rK+H)2]1/2?L?1.2·[(rK+b1)2?(rK+H)2]1/2 or
b1?0.5·b2 or
(b2?b1)/s?0.2
with the inlet channel height b1 at the inlet edge, the radius of curvature of the upstream channel wall rK, the radial distance between the inlet edge and the transition of the annular space into the upstream channel wall H, the axial distance between the inlet edge and the transition of the annular space into the upstream channel wall L, the outlet channel height b2 at the bleed channel outlet and the length of the downstream channel wall between the inlet edge and the bleed channel outlets.

US Pat. No. 10,060,439

GUIDE VANE FOR A TURBOMACHINE, GUIDE VANE CASCADE, AND METHOD FOR MANUFACTURING A GUIDE VANE OR A GUIDE VANE CASCADE

MTU Aero Engines AG, Mun...

1. A guide vane for a turbomachine comprising:an upper trailing edge on an upper surface of the guide vane and a lower trailing edge on a lower surface of the guide vane, the guide vane being axially pivotably coupled to a radially outwardly disposed flow-limiting wall and to a radially inwardly disposed inner ring of the turbomachine, a trailing edge gap being formed between the upper trailing edge of the guide vane and the flow-limiting wall or between the lower trailing edge of the guide vane and the inner ring,
wherein the upper trailing edge or the lower trailing edge of the guide vane has at least one air outlet opening for an air outflow for forming an air curtain for at least partially sealing the trailing edge gap in the area of the upper trailing edge or the trailing edge gap in the area of the lower trailing edge;
wherein the guide vane has a blade surface, the air outlet opening communicating air-conductively via at least one channel with at least one air inlet opening on the blade surface; and
wherein the air inlet opening is configured on an active surface of the guide vane, the active surface being a side of the guide vane extending between the upper and lower surfaces.

US Pat. No. 10,030,585

SHAFT SEAL SYSTEM AND A COMPRESSOR HAVING A CORRESPONDING SHAFT SEAL SYSTEM

MTU Aero Engines AG, Mun...

1. A shaft seal system which is disposed axially between a first chamber and a second chamber, comprising:a shaft;
a casing that surrounds the shaft; and
a seal, wherein the seal is disposed axially closer to the second chamber than the first chamber;
wherein the first chamber includes a fluid;
wherein the seal includes a mechanical pressure booster, wherein the mechanical pressure booster is shaped such that a pressure of a sealing gas flowing at the seal and in a direction of the first chamber is increased by the mechanical pressure booster, wherein the sealing gas is flowable from the second chamber to the first chamber along the shaft, and wherein the sealing gas is provided from a compressor;
wherein the mechanical pressure booster comprises a first pressure booster and a second pressure booster and wherein the first pressure booster comprises a helical groove on the casing and the second pressure booster comprises a radially opposite series of spaced apart longitudinal indentations on the shaft.

US Pat. No. 9,994,934

CREEP-RESISTANT TIA1 ALLOY

MTU Aero Engines AG, Mun...

1. A component made from a TiAl alloy, wherein the alloy comprisesnot more than 43 at. % of Al,
from 3 at. % to 8 at. % of Nb,
from 0.2 at. % to 3 at. % of Mo and/or Mn,
from 0.05 at. % to 0.5 at. % of B,
from 0.1 at. % to 0.5 at. % of C,
from 0.1 at. % to 0.5 at. % of Si and
Ti as balance;and wherein the component has a microstructure comprising from about 70% to 80% by volume of ?-TiAl, from about 20% to 25% by volume of ?2-Ti3Al, and from about 1% to 3% by volume of ?o-Ti.

US Pat. No. 9,988,927

HOUSING FOR A GAS TURBINE, AIRCRAFT ENGINE, AND A PROCESS FOR OPERATING A GAS TURBINE

MTU Aero Engines AG, Mun...

1. A housing for a gas turbine, comprising:a wall element movably disposed on the housing, wherein the wall element limits a flow channel of the housing in a radial direction from a rotational axis of a rotor of the gas turbine; and
a variably adjustable guide blade which extends through the wall element into the flow channel;
wherein the wall element defines an opening and wherein a blade disk of the guide blade is disposed in the opening;
wherein the wall element is movable between a sealing setting in which the wall element makes contact with at least a partial area of a side of a blade leaf of the guide blade facing toward the wall element and an open setting in which the blade leaf and the wall element are spaced apart.

US Pat. No. 9,982,547

GUIDE MECHANISM FOR A GAS TURBINE AND GAS TURBINE HAVING SUCH A GUIDE MECHANISM

MTU AERO ENGINES AG, Mun...

25. A guide mechanism for a gas turbine, comprising:at least one casing element, the casing element having a passage opening therethrough;
at least one first duct segment arranged in the radial direction on the inside of the casing element, the first duct segment defining a portion of at least one duct through which gas can flow, at least partially delimited outward in the radial direction; the at least one first duct segment is mounted relative to the at least one casing element via at least two tabs and at least two grooves, which are spaced apart in the axial direction;
at least one second duct segment arranged in the radial direction on the inside of the first duct segment, the second duct segment defining a portion of the duct, at least partially delimited inward in the radial direction;
at least one guide vane arranged at least partially in the duct, the guide vane configured and arranged to rotate around an axis of rotation relative to the at least one casing element and relative to the duct segments; and
at least one bushing inserted in the radial direction from the outside to the inside into the passage opening of the casing element, the bushing projecting radially inward from the casing element; the bushing has a wall thickness that is greater than the length of at least one of the tabs of the first duct segment projecting into one of the grooves such that removal of the bushing allows sufficient mobility to move the tabs from an uninstalled state to an installed state within the grooves;
wherein the guide vane is mounted on the casing element via the bushing;
wherein the second duct segment is fastened to the guide vane and is retained on the at least one casing element via the guide vane.

US Pat. No. 9,943,907

APPARATUS AND METHOD FOR GENERATIVE PRODUCTION OF A COMPONENT

MTU AERO ENGINES AG, Mun...

1. A method for the generative production of a component, comprising the steps of:disposal of a material layer by movement of a dispensing device;
local bonding of the material layer to a cross section of the component being produced; and
displacement of a platform supporting the component being produced counter to a direction of layer buildup;
sensing the position and change in position of the platform by scanning the position of the platform over time in the direction of layer build up;
determining whether the movement of the dispensing device is not free of interference with the component being produced based on the sensed position or change in position of the platform in the direction of layer buildup; and
triggering a response based on the determination of whether the movement of the dispensing device is not free of interference.

US Pat. No. 9,920,838

BRUSH SEAL

MTU Aero Engines AG, Mun...

1. A brush seal for a gas turbine, comprising:
a support ring, wherein the support ring is comprised of a mounting plate and a support element, wherein a radial inner end
of the mounting plate is integral with or integrally molded onto a radial outer end of the support element and wherein the
support element has a support plate and a support structure, wherein the support structure is arranged downstream in a direction
of flow with respect to the support plate and wherein a clearance is defined between the support structure and the support
plate; and

bristles which are arranged upstream in the direction of flow with respect to the support ring and wherein ends of the bristles
protrude radially inward beyond the support plate.

US Pat. No. 9,920,640

EXTRUDED PROFILE FOR MANUFACTURING A BLADE OF AN OUTLET GUIDE VANE

MTU Aero Engines AG, Mun...

1. An extruded profile for manufacturing a blade of an outlet guide vane of a turbine engine, the extruded profile comprising:
a cross-sectional area having an axial length LAX and a thickness D/LAX relative to axial length LAX, along axial length LAX
thereof, the cross-sectional area having the following mutually adjoining regions:

an at least nearly axisymmetric leading edge region;
a first transition region having a relative thickness D/LAX varying along the first transition region;
a first constant region having a relative thickness D/LAX at least substantially constant along the first constant region,
the first constant region, relative to a leading edge of the extruded profile, beginning at the closest at 10% LAX and ends
at the furthest at 50% LAX;

a second transition region having a relative thickness D/LAX varying along the second transition region, the second transition
region, relative to the leading edge, beginning at the closest at 30% LAX and ending at the furthest at 90% LAX;

a second constant region having a relative thickness D/LAX at least substantially constant along the second constant region,
the second constant region having an axial length X of 40% LAX at most; and

an at least nearly axisymmetric trailing edge region, wherein the first constant region has a relative thickness D/LAX of
between 3% and 6% LAX or, relative to the leading edge of the extruded profile, the first constant region extends within the
range of approximately 20% to 40% LAX, and wherein the first constant region relative thickness D/LAX is approximately 4%
LAX.

US Pat. No. 10,197,472

METHOD FOR PERFORMING MAINTENANCE ON AN ENGINE

MTU Aero Engines AG, Mun...

1. A method for performing maintenance on an engine, comprising the steps of:providing an engine maintenance system including a database system having a database and a database management device;
providing at least one first performance parameter stored in the database and characterizing an engine performance before an engine maintenance procedure;
providing at least one maintenance parameter stored in the database and characterizing a scope of a maintenance measure performed on the engine during the engine maintenance procedure;
providing at least one second performance parameter stored in the database and characterizing the engine performance after the engine maintenance procedure;
determining, using the database management device, a functional relationship between the maintenance parameter and the contribution of the maintenance parameter to a difference between the first performance parameter and the second performance parameter;
outputting the functional relationship via the engine maintenance system; and
performing maintenance on the engine taking the functional relationship into account;
wherein the functional relationship comprises
?Ymodel,j=f(Xi,j,ai)
wherein ?Ymodel,j is a difference between the first and second performance parameters for an engine j;
wherein Xi,j is a maintenance parameter for a maintenance measure i and the engine j; and
wherein ai is a function parameter for the maintenance measure i.

US Pat. No. 9,951,623

BLADE CASCADE FOR A TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A blade cascade for a gas turbine, comprisinga first group with
at least two,
first blade arrangements that are of the same type to one another, wherein each blade of the first blade arrangements comprises a
first detuner arrangement having at least one first detuner and having only one detuner type movably mounted therein, wherein each first detuner arrangement on each blade of the first blade arrangements is identical to one another; and
a second group with
at least two, second blade arrangements, which are of the same type to one another and of different type from the first blade arrangements, wherein each blade of the second blade arrangements comprises a
second detuner arrangement having at least one second detuner and having only one detuner type movably mounted therein, wherein each second detuner arrangements on each blade of the second blade arrangements is identical to one another and of different type than the first detuner arrangement on the blades of the first blade arrangements.

US Pat. No. 9,938,858

MID-FRAME FOR A GAS TURBINE AND GAS TURBINE HAVING SUCH A MID-FRAME

MTU AERO ENGINES AG, Mun...

1. A mid-frame for a gas turbine, having at least one outer casing element, having at least one hub element arranged on the inside of the outer casing element in the radial direction, having at least one strut connecting the outer casing element to the hub element, and having at least one fairing element that delimits at least partially a duct at least in the radial direction, through which a gas can flow, said fairing element having a passage opening, through which the strut passes, the passage opening cladding the strut on an outer peripheral side of the strut that passes through the fairing element, wherein the fairing element is coupled in the radial direction to at least one support member of the hub element and at least one guide vane, the at least one guide vane for at least partial guiding of the gas flowing through the duct.

US Pat. No. 9,926,786

METHOD FOR JOINING AT LEAST TWO ROTOR ELEMENTS OF A TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A method for joining at least two rotor elements of at least one rotor of a turbomachine, the method comprising at least the following steps:detecting any radial runout of at least one radially outer-lying cylindrical surface of the rotor elements at each of at least two points that are spaced axially apart from each other by a measuring device;
determining a distance and an angular position of a center of mass with respect to an axis of rotation of the respective rotor element as a function of the respectively detected radial runout;
determining a respective distance of a total center of mass of the rotor assembled from the rotor elements with respect to its total axis of rotation for various relative mounting alignments of the rotor elements with respect to one another as a function of the previously determined centers of mass of the respective rotor elements by an analysis device;
determining of those alignments of these various relative mounting alignments of the rotor elements with respect to one another for which the distance of the total center of mass of the rotor is minimized relative to its total axis of rotation; and
joining the rotor elements to one another by that mounting alignment for which the previously determined distance of the total center of mass of the rotor to be manufactured is minimized relative to its total axis of rotation,
wherein the step of detecting any radial runout of the radially outer-lying cylindrical surface of the rotor elements occurs optically by at least one optical sensor element of the measuring device, and
wherein, by the optical sensor element, an eccentricity of an entire respective cylindrical surface of the rotor elements is captured, and wherein, by the analysis device, the distance and the angular position of the center of mass from the axis of rotation of the respective rotor element are determined as a function of this detected eccentricity of the entire respective cylindrical surface.

US Pat. No. 9,840,919

METHOD FOR PRODUCING A RUN-IN COATING, A RUN-IN SYSTEM, A TURBOMACHINE, AS WELL AS A GUIDE VANE

MTU Aero Engines AG, Mun...

1. A method for producing a run-in coating for a turbomachine for braking a rotor in the event of a shaft breakage, comprising:
forming the run-in coating as an integral blade portion during a generative manufacture of a blade, wherein the generative
manufacture includes forming the blade and the run-in coating through deposition and melting of metal powder in layers.

US Pat. No. 9,835,166

ARRAY OF FLOW-DIRECTING ELEMENTS FOR A GAS TURBINE COMPRESSOR

MTU Aero Engines AG, Mun...

1. An array of flow-directing elements for a compressor of a gas turbine, the array comprising:
at least one first flow-directing element; and
at least one second flow-directing element different from the first flow-directing element; the first and second flow-directing
elements each having a leading edge facing an inlet of the gas, a trailing edge facing away from the gas turbine inlet, a
pressure side connecting the leading edge and the trailing edge and located ahead in a direction of operational rotation,
a suction side located opposite of the pressure side, and successive chords along a stacking axis; the first and second flow-directing
elements each extending between an airfoil root and an airfoil tip;

the trailing edge of the first flow-directing element being, at least in a portion thereof, axially offset from the trailing
edge of the second flow-directing element in a direction toward the leading edge of the first flow-directing element, at least
in a half proximate to the airfoil tip; and

at least one third flow-directing element having a third flow-directing trailing edge and a third flow-directing leading edge,
at least in a portion thereof, axially offset from the trailing edges of the first and second flow-directing elements in a
direction toward the third flow-directing leading edge;

wherein the trailing edge of the first flow-directing element is, at least in a portion thereof, offset from the trailing
edge of the second flow-directing element by at least 0.5% and no more than 15% of a chord length of the first flow-directing
element.

US Pat. No. 9,822,258

CR(VI)-FREE CORROSION PROTECTION LAYERS OR ADHESION PROMOTER LAYERS PRODUCED USING A SOLUTION COMPRISING PHOSPHATE IONS AND METAL POWDER, WHEREIN THE METAL POWDER IS COATED AT LEAST PARTLY WITH SI OR SI ALLOYS

MTU AERO ENGINES AG, Mun...

1. A coating material for producing a corrosion protection layer and/or adhesion promoter layer, wherein the coating material
comprises metal powder and a solution containing phosphate ions as binder, and wherein the metal powder comprises (a) from
60% to 99% of Al powder and (b) from 1% to 40% of Al powder coated with Si or Si alloys.

US Pat. No. 9,816,379

BALANCING BODY FOR A CONTINUOUS BLADE ARRANGEMENT

MTU AERO ENGINES AG, Mun...

1. A balancing body for fastening to a ring of a continuous blade arrangement of a compressor or turbine stage of a gas turbine,
wherein the balancing body comprises a first stop for form-fitting attachment of the balancing body in one peripheral direction
(R) to a first axial shoulder of the ring wherein the first axial shoulder is an axial material projection from an outer shroud
of the ring and at least one undercut section for the form-fitting attachment of the balancing body to the ring in a radial
direction;
wherein the balancing body is distanced from adjacent rotating blades of the continuous blade arrangement.

US Pat. No. 9,771,896

MIXING DEVICE AND TURBOFAN ENGINE HAVING SUCH MIXING DEVICE

MTU AERO ENGINES AG, Mun...

1. A mixing device for mixing a first gas flow with a second gas flow in a turbofan engine, comprising:
an actuating device and walls that bound a radially inner-lying channel for the first gas flow and a radially outer-lying
channel for the second gas flow,

the actuating device comprising a first coupling element, having a first end and a second end; the first end of the first
coupling element being coupled to the walls,

the actuating device being configured and arranged to pivot the walls between a first position and a second position disposed
radially outside relative to the first position, by the first coupling element, wherein

the actuating device comprises an adjusting ring that rotates between a first rotating position and a second rotating position
in a peripheral direction and that is joined to the first coupling element,

the first coupling element being configured and arranged as rigid and being coupled to the adjusting ring via the second end
of the first coupling element whereby rotating the adjusting ring between the first rotating position and the second rotating
position, the walls are pivoted by the first coupling element between the first position and the second position corresponding
to the rotation of the adjusting ring;

a support structure, which is disposed bounding the walls in an axial direction, the support structure comprising a fastening
element, where axially, the adjusting ring is disposed between the fastening element and the walls; and

a second coupling element having a first end and a second end; the first end of the second coupling element being coupled
to the fastening element of the support structure and the second end of the second coupling element being coupled to the adjusting
ring wherein the first end of the second coupling element and the first end of the first coupling element are axially on opposing
sides of and axially spaced apart from the adjusting ring.

US Pat. No. 10,507,526

METHOD AND DEVICE FOR THE ADDITIVE MANUFACTURE OF AT LEAST ONE COMPONENT REGION OF A COMPONENT

MTU Aero Engines AG, Mun...

1. A method for the additive manufacture of at least one region of a component, comprising the following steps:a) layer-wise application of at least one powder-form component material onto a component platform in the region of a build-up and joining zone;
b) layer-wise and local solidifying of the component material by selective exposure of the component material by at least one high-energy beam in the region of the build-up and joining zone, with the formation of a component layer;
c) layer-wise lowering of the component platform by a pre-defined layer thickness; and
d) repeating steps a) to c) until the component region or the component has been completely fabricated;
wherein during at least one step b), at least one exposure parameter of the high-energy beam from the group: power, velocity, and focal position, is adjusted as a function of at least one construction parameter from the group: component thickness, hatch distance to an adjacent exposure trace, angle of incidence of the high-energy beam relative to the surface of the component layer, angle of deflection of the high-energy beam with respect to a vertical axis of the component layer, overhang angle of the component layer, layer thickness of the component layer, and distance from a complete volume element of the component layer, and
wherein during at least one step b), the method further comprises determining, with a controller, at least the overhang angle of the component layer, and adjusting
at least the exposure parameter, power, of the high-energy beam to reduce the power, when compared to the power in an inskin region, if the construction parameter, overhang angle, of the component layer in the exposed region corresponds to a downskin region, and
at least the exposure parameter, power, of the high-energy beam to increase the power, when compared to the power in an inskin region, if the construction parameter, overhang angle, of the component layer in the exposed region corresponds to an upskin region.

US Pat. No. 10,107,112

METHOD FOR PRODUCING FORGED COMPONENTS FROM A TIAL ALLOY AND COMPONENT PRODUCED THEREBY

MTU AERO ENGINES AG, Mun...

1. A method for producing a component from a TiAl alloy, wherein the method comprises shaping the component by forging and subsequently subjecting the component to at least two heat treatments and wherein in a first heat treatment a temperature of from 1100° C. to 1200° C. is maintained for 6 to 10 hours, whereafter the component is cooled at a cooling rate of from 1° C./s to 5° C./s, and wherein in a second heat treatment the component is heated to a temperature above a solvus line of ?-TiAl.

US Pat. No. 10,081,877

METHOD FOR THE ELECTROPLATING OF TIAL ALLOYS

MTU AERO ENGINES AG, Mun...

1. A method for coating a surface of a TiAl alloy, comprising the steps of:providing a TiAl alloy having a surface;
roughening the surface of the TiAl alloy in a two-step surface treatment, including:
applying an electrochemical process to the surface; and
treating the surface with an electroless chemical process after the electrochemical process;
wherein the surface of the TiAl alloy is sufficiently roughened after the electrochemical process and electroless chemical process without mechanical roughening of the surface of the TiAl alloy;
chemically activating the surface of the TiAl alloy after roughening the surface of the TiAl alloy in the two-step surface treatment; and
electroplating at least one layer on the surface of the TiAl alloy after the step of chemically activating the surface of the TiAl alloy,
wherein the electrochemical processing is conducted by anodic etching in an acetic acid-hydrofluoric acid solution, wherein concentrations by weight of 800 to 900 g/L of acetic acid and 100 to 200 g/L of hydrofluoric acid are selected for the composition of the acetic acid-hydrofluoric acid solution.

US Pat. No. 10,066,508

METHOD FOR PRODUCING, REPAIRING AND/OR EXCHANGING A HOUSING, IN PARTICULAR AN ENGINE HOUSING, AND A CORRESPONDING HOUSING

MTU AERO ENGINES AG, Mun...

1. A method for producing a component, comprising:layer-by-layer constructing a housing of an aircraft engine, the housing having an inner shell part and an outer shell part together with a structural part therebetween, by a generative manufacturing method, wherein the generative manufacturing method comprises applying a powder layer and solidifying the powder layer by energy radiation, wherein the structural part has at least one porous structure and/or one honeycomb structure, wherein the step of the layer-by-layer constructing further comprises:
layer-by-layer constructing an auxiliary support structure by the generative manufacturing method on a base plate;
subsequently forming the housing on the auxiliary support structure by the generative manufacturing method, wherein the auxiliary support structure and housing are integrally connected; and
removing the auxiliary support structure from the housing after finishing the housing.
US Pat. No. 10,060,012

HIGH-TEMPERATURE TIAL ALLOY

MTU AERO ENGINES AG, Mun...

1. A TiAl alloy for use at high temperatures, wherein the alloy comprises:from 30 to 42 at. % of Al
from 5 to 25 at. % of Nb
from 2 to 10 at. % of Mo
from 0.1 to 10 at. % of Co
from 0.1 to 0.5 at. % of Si
from 0.1 to 0.5 at. % of Hf,balance Ti,and wherein the alloy comprises a matrix of ? phase and precipitates of ? phase embedded in the matrix, the ? phase and the ? phase together making up at least 55% by volume of a microstructure of the alloy.
US Pat. No. 10,005,128

PRODUCTION PROCESS FOR TIAL COMPONENTS

MTU AERO ENGINES AG, Mun...

1. A process for producing a component of a TiAl alloy, wherein the process comprises:introduction of a powder of the TiAl alloy into a capsule whose shape corresponds to a shape of the component to be produced and closing of the capsule,
hot isostatic pressing of the capsule together with the powder,
heat treatment of the hot isostatically pressed capsule,
removal of the capsule,
post-working of a contour of the component by removal of material.

US Pat. No. 9,932,847

GUIDE BLADE FOR A GAS TURBINE

MTU Aero Engines AG, Mun...

1. A guide blade for a gas turbine, comprisinga blade leaf with a receptacle, wherein a sealing element is disposed in the receptacle, wherein the sealing element is movable relative to the blade leaf between a sealing setting in which the sealing element is disposed at least partially out of the receptacle and a storage setting in which the sealing element is disposed in the receptacle;
wherein the blade leaf has a fluid channel, wherein a fluid under pressure is routable into the receptacle via the fluid channel in order to move the sealing element from the storage setting to the sealing setting;
and wherein an inlet opening of the fluid channel is formed on a pressure-side surface of the blade leaf.

US Pat. No. 9,931,719

METHOD FOR REPAIRING A RECEIVING HOOK FOR GUIDE VANES

MTU AERO ENGINES AG, Mun...

1. A method for repairing a receiving hook for a guide vane, which receiving hook is arranged in a housing of a turbomachine, wherein the method comprises:removing first material in a region of the receiving hook, which region extends over a circumference of the housing; followed by
applying second material in or on a region which extends over the circumference of the housing, and applying the second material using localized heat.

US Pat. No. 9,822,796

GAS TURBINE COMPRESSOR STATOR VANE ASSEMBLY

MTU Aero Engines AG, Mun...

1. A stator vane assembly for a compressor of a gas turbine comprising:
a plurality of stator vanes whose airfoil sections form a stagger angle with an axis of rotation of the compressor, the stagger
angle varying along a duct height of the stator vane assembly,

wherein along the duct height between a radially innermost initial value and a radially outermost final value, the stagger
angle increases to a local maximum in a second section adjoining a first, radially innermost section, and decreases to an
outer local minimum in a third section adjoining said second section, and

wherein, along the duct height from the inside to the outside, the stagger angle decreases from the initial value to an inner
local minimum in the first, radially innermost section or increases from the outer local minimum to the final value in a fourth,
radially outermost section adjoining the third section.

US Pat. No. 9,765,632

PROCESS FOR PRODUCING A TIAL GUIDE VANE RING FOR A GAS TURBINE AND A CORRESPONDING GUIDE VANE RING

MTU AERO ENGINES AG, Mun...

1. A process for producing a blade or vane ring segment for a gas turbine which comprises at least two adjacent main blade
or vane parts having a single common blade or vane root, wherein the process comprises:
(a) forging at least two blanks made of a TiAl material,
(b) joining the blanks to form the blade or vane ring segment by an integral connection process, and
(c) remachining of a blank composite thus obtained by a material-removing process; (b) being carried out to result in a joining
zone which extends through only a center or a central region of the common blade or vane root and each blank being formed
as a cuboid with protruding joining zones.

US Pat. No. 10,208,765

GAS TURBINE AXIAL COMPRESSOR

MTU Aero Engines AG, Mun...

1. An axial compressor for a gas turbine, in particular an aircraft engine, having at least one rotor blade or guide vane having a blade or vane element, which is arranged in the flow duct, and a leading edge and a trailing edge, which are joined to each other through a pressure side and a suction side, wherein,in each profile section of the blade or vane element in a range between 5% and 95% of a radial blade or vane element height from a blade or vane element root to a blade or vane tip,
a profile skeletal line has at most two points of inflection;
a metal angle at the leading edge lies in a range between 20° and 75°;
a metal angle at the trailing edge lies in a range between ?20° and 70° and is at most equal to the metal angle at the leading edge; and
a stacking angle lies in a range between 0° and 70°;
and wherein
a local or absolute minimum of a front load angle, which is equal to the difference obtained from the arithmetic mean of the metal angles at the leading edge and trailing edge as minuend and the stacking angle as subtrahend

lies in a range between 55% and 85% of the radial blade or vane element height; and a local or absolute maximum of the front load angle lies in a range between 95% and 100% of the radial blade or vane element height at a blade or vane tip.

US Pat. No. 10,100,655

BRUSH SEAL AND METHOD FOR PRODUCING A BRUSH SEAL

MTU AERO ENGINES AG, Mun...

1. A brush seal for the sealing of gaps occurring in turbomachines, comprising:a plurality of individual fibers or wires;
at least two of the individual fibers or wires are braided together, along their length, into at least two fiber bundles or wire bundles, wherein each of the at least two fiber bundles or wire bundles includes at least two fibers or wires braided together along their length;
the at least two fiber bundles or wire bundles are braided or twisted, along their length, into a fiber or wire package configured as a bristle for the brush seal; the diameter of the fiber package or wire package being between 0.05 and 2.0 mm; and
a plurality of fiber packages or wire packages being configured and arranged as the brush seal for bridging a large gap between components in the turbomachine.

US Pat. No. 10,041,363

BLADE-DISK ASSEMBLY, METHOD AND TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A blade-disk assembly of a turbomachine, the blade-disk assembly comprising:a plurality of adjacent rotor blades and a closure blade inserted with root portions in a circumferential anchoring groove of a rotor disk and cooperating with supporting portions of a forward wall and an aft wall of the circumferential anchoring groove in radially interlocking relationship therewith;
at least one circumferential retention body having surfaces interlockingly cooperating with at least one of the closure blade and the adjacent rotor blades; and
a plurality of tilt-out prevention elements disposed between a groove base and the root portions and, in a rest state, spacing the adjacent rotor blades from the groove base when in an upper position, wherein the tilt-out prevention elements are formed sheet-metal members supported, in the rest state, with supporting legs on the groove base; and
a locking element which provides locking and tilt-out prevention, the locking element including a threaded shank and a head wider than the threaded shank, the locking element supported in an inclined internally threaded bore formed in the closure blade of the plurality of adjacent rotor blades and extending radially and at an inclined angle in a circumferential direction.

US Pat. No. 10,013,814

DIAGNOSIS OF AIRCRAFT GAS TURBINE ENGINES

MTU AERO ENGINES AG, Mun...

1. A method for the at least partially automated diagnosis of aircraft gas turbine engines, comprising:detecting actual parameter values of an aircraft gas turbine engine for several operational segments with sensors disposed within the aircraft gas turbine engine, the sensor types include at least one of a temperature sensor, a pressure sensor, a mass flow sensor, and a rotational speed sensor, the parameter values comprising temperatures in one or more operational segments of the gas turbine, pressures in one or more stages of the gas turbine, fuel consumption and/or fuel mass flow, and rotational speed values of a rotor of the aircraft gas turbine engine;
determining deviation, of these actual parameter values from theoretical parameter values of the operation of the aircraft gas turbine engine;
determining damage pattern probabilities based on a similarity of at least one determined deviation to deviation patterns of various known damage patterns to aircraft gas turbine engines;
determining a damage pattern probability for an unknown damage pattern; and
generating an error message if at least one determined deviation exceeds a predetermined limit value,
wherein damage patterns and/or causes comprise components, subassemblies, and/or defects.

US Pat. No. 9,816,386

CASING ARRANGEMENT FOR A GAS TURBINE

MTU AERO ENGINES AG, Mun...

1. A gas turbine casing arrangement, comprising:
a gas turbine casing element;
a guide-vane ring with an outer ring; and
a coated ring, which lies radially opposite a rotor grid adjacent to the guide-vane ring; wherein
an intermediate ring, which,
by a downstream front face engages behind an upstream stop fixed in place on the casing element, and
by an upstream front face of a radial flange of the intermediate ring engages behind a downstream stop fixed in place on the
outer ring, in order to axially secure the guide-vane ring at the gas turbine casing element in the direction of through-flow;

wherein in an operating position, a surface area of the radial flange abuts the coated ring at a support wall disposed at
a seal.

US Pat. No. 10,285,222

METHOD AND DEVICE FOR GENERATIVELY PRODUCING AT LEAST ONE COMPONENT AREA

MTU AERO ENGINES AG, Mun...

1. A method for generatively producing or for repairing at least one area of a component which is made up of individual powder layers, wherein the method comprises(i) locally heating, by a first high-energy beam, a powder layer to a melting temperature (T2), whereby a molten bath is formed locally at a part of the component corresponding to the first high-energy beam, the first high-energy beam being moved across the component so that the molten bath is formed in consecutive parts of the component,
(ii) post-heating to a post-heating temperature (T3), by second high-energy beam which follows a movement of the first high-energy beam, a part arranged downstream of a current molten bath, which part has already been heated by the first high-energy beam, and
(iii) setting, by an additional heating device, a temperature of the component in its entirety to a base temperature (T1),
wherein T2>T3>T1 and wherein the part arranged downstream of a current molten bath adjoins the molten bath so that steep changes in temperature between the current molten bath and a post-heated part are avoided.

US Pat. No. 10,145,003

CMAS-INERT THERMAL BARRIER LAYER AND METHOD FOR PRODUCING THE SAME

MTU AERO ENGINES AG, Mun...

1. A method of modifying a thermal barrier layer arranged on a metallic component, wherein the thermal barrier layer comprises a ceramic coat and the method comprises incorporating aluminum oxide and titanium oxide in the ceramic coat by infiltrating the ceramic coat with aluminum-containing and titanium-containing particles or substances, the aluminum oxide and the titanium oxide being incorporated only in an outer region of the ceramic coat.

US Pat. No. 10,131,963

PNEUMATIC NEEDLING DEVICE

MTU Aero Engines AG, Mun...

1. A method for operating a pneumatic needling devicefor local surface treatment, including strengthening, of components, having a first needle (2), which is movable in a needle direction;
a first piston chamber (3) for the pneumatic pressurizing of the first needle in its needle direction;
a pressure supply (1), which can be connected to and separated from the first piston chamber, by a movement of the first needle in its needle direction;
a pressure detecting means (5) for determining a first pressure fluctuation in the first piston chamber;
a second needle (2), which is movable in a needle direction;
a second piston chamber (3) for the pneumatic pressurizing of the second needle in its needle direction;
a pressure detecting means (5) for determining a second pressure fluctuation in the second piston chamber;
a pressure supply (1), which can be connected to and separated from the second piston chamber, by a movement of the second needle in its needle direction; and
a control means (6),
wherein the method comprises the steps:
determining (S10) the first pressure fluctuation in the first piston chamber (3);
determining (S10) a second pressure fluctuation in the second piston chamber (3); and
executing (S20, S40) a response (R1, R2) based on the determined first and/or second pressure fluctuation.

US Pat. No. 10,107,194

GAS TURBINE

MTU Aero Engines AG, Mun...

1. A gas turbine comprising:at least one casing provided with at least one bleed duct and at least one bore;
at least one variable guide vane having a trunnion, a rotary plate, and a guide vane airfoil, the trunnion being disposed in the bore, and the guide vane airfoil being integrally formed with the rotary plate,
the guide vane airfoil extending beyond the rotary plate in such a way that, viewed in a direction of flow of the gas path of the gas turbine, a flag corner of the guide vane airfoil facing the bleed duct is located downstream of a beginning of the inlet opening of the bleed duct, wherein the bleed duct connects a primary gas path to a secondary gas path; and
at least one rotor blade located downstream of the guide vane, the at least one casing separating the primary flow path from the secondary flow path, the primary flow path and the secondary flow path extending axially past the at least one guide vane and the at least one rotor blade vane.

US Pat. No. 9,951,631

TURBOMACHINE ROTOR BLADE

MTU Aero Engines AG, Mun...

1. A rotor blade for a turbomachine, comprising:an airfoil for flow deflection;
a blade root for attachment to a rotor of the turbomachine;
an inner platform between the airfoil and the blade root; and
two axially spaced walls defining at least one pocket, the two axially spaced walls extending from the side of the platform opposite the airfoil toward the blade root;
the rotor blade being configured such that in the mounted state, facing pockets of adjacent rotor blades as well as the inner platforms, contact each other, so that their walls form two at least substantially closed, axially spaced sealing rings or flanges between the rotor and a ring defined by the inner platform,
a first of the walls having an outer side facing away from the pocket and, in at least one cross section perpendicular to a radial longitudinal axis of the rotor blade, slopes outwardly as it extends in a circumferential direction, so that the outer side of the first wall diverges in an axial direction with increasing distance as it extends in the circumferential direction from the rotor blade toward a rotor blade that is circumferentially adjacent in the mounted state.

US Pat. No. 9,561,556

PROCESS FOR PRODUCING INTERMETALLIC WEAR-RESISTANT LAYER FOR TITANIUM MATERIALS

MTU AERO ENGINES AG, Mun...

1. A process for producing a wear-resistant layer on a component, wherein the process comprises:
providing a component which comprises a titanium material on at least part of a surface of the component on which the wear-resistant
layer is to be produced,

contacting the titanium material with a solder in the form of a paste or semifinished product formed from a cobalt base material,
soldering the solder to the titanium material by applying heat to thereby produce one or more diffusion zones between solder
and titanium material, which one or more diffusion zones comprise one or more intermetallic phases and form the wear-resistant
layer, the soldering being carried out at a temperature of from about 1100° C. to about 1200° C.

US Pat. No. 10,287,989

SEAL SUPPORT OF TITANIUM ALUMINIDE FOR A TURBOMACHINE

MTU AERO ENGINES AG, Mun...

1. A turbomachine, wherein the turbomachine comprises an annular flow duct and a housing structure surrounding the flow duct and guide vanes and rotor blades which are arranged in the flow duct, the rotor blades being rotatably accommodated in the housing structure and the guide vanes being fixed in the housing structure, and a plurality of the guide vanes forming an annular guide vane ring, and wherein the housing structure has a seal in a region of a radially inner flow duct boundary to prevent hot gas escaping from the flow duct, said seal being arranged on guide vane roots of the guide vanes of the guide vane ring via a seal support and sealing against a rotatable seal surface, said seal support being formed from an intermetallic material.

US Pat. No. 10,287,910

ADJUSTABLE GUIDE VANE FOR A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. An adjustable guide vane for a turbomachine, comprisinga vane element and a turning disk, wherein the turning disk has a first impact chamber, in which an impulse element is arranged with play of movement,
wherein the first impact chamber is arranged on an arc, in the turning disk, around the axis of rotation of the guide vane and turning disk, and
wherein the impulse element is configured and arranged to dampen vibrations in the turning disk and the guide vane.

US Pat. No. 10,215,050

APPARATUS FOR ENERGY ABSORPTION, A TURBO-MACHINE AND A METHOD FOR ENERGY ABSORPTION

MTU Aero Engines AG, Mun...

1. An apparatus for absorption of an impact energy acting on an element from a blade fragment, wherein the element is disposed through an opening in a turbo-machine, wherein the opening passes radially through a housing section in a region of a row of rotating blades, and wherein the element has a section that protrudes into a flow channel of the turbo-machine, comprising:a holder, wherein the holder holds the element in a position and releases the element when a preset maximum load acting on the element is exceeded; and
a cage disposed at a rear of the apparatus, wherein the cage limits a radial displacement of the element and wherein the holder releases the element such that the released element runs onto a bottom wall of the cage;
wherein the element is a sensor or a probe.

US Pat. No. 10,180,206

PIPE ARRANGEMENT WITH SUPPORT SECTIONS ON THE OUTER PIPE

MTU Aero Engines AG, Mun...

1. A pipe arrangement with a fluid-carrying inner pipe and an outer pipe that surrounds the inner pipe, wherein spacers are arranged between the inner pipe and the outer pipe with a gap formed between the inner pipe and the outer pipe, wherein the outer pipe is joined at its first axial end section in a material-bonded or cohesive manner to a first spacer, which is joined to the inner pipe in a material-bonded manner, and wherein the outer pipe is arranged at its second axial end section at least partially with a radial distance to a second spacer, which is joined to the inner pipe in a material-bonded manner, whereinthe outer pipe has a plurality of support sections at its second end section, which are arranged in a distributed manner in the peripheral direction and project radially inward and which are in contact with the second spacer, and
the second spacer and a second axial end section of the inner pipe are configured and arranged to axially move relative to the outer pipe.

US Pat. No. 10,151,209

SEALING SYSTEM MADE OF CERAMIC FIBER COMPOSITE MATERIALS

MTU Aero Engines AG, Mun...

1. A sealing system comprising:a first component at least partially manufactured from ceramic fiber composite materials;
a second component at least partially manufactured from ceramic fiber composite materials;
a sealing element accommodated between the first component and the second component, the sealing element being designed as a sealing strip,
at least one recess accommodating the sealing element being formed on the first component or on the second component, the recess having a cross-sectional profile including an inner section concave toward the sealing element, a radius of curvature of the inner section being selected so as to be formable with continuous, curved fibers of the ceramic fiber composite material.

US Pat. No. 10,428,437

WEAR-RESISTANT COATING PRODUCED BY ELECTRODEPOSITION AND PROCESS THEREFOR

MTU AERO ENGINES AG, Mun...

1. A coating, wherein the coating is wear-resistant and comprises an electrodeposited matrix which comprises from 15% by weight to 50% by weight Co, from 15% by weight to 50% by weight Ni, from 10% by weight to 30% by weight Cr, and from 1% by weight to 10% by weight Al, and in which first particles comprising hard material particles and/or slip material particles are incorporated in a proportion of from 5% by volume to 40% by volume.

US Pat. No. 10,309,538

DEVICE AND METHOD FOR ATTACHING SEAL ELEMENTS

MTU Aero Engines AG, Mun...

1. A device for attaching a seal element in a recess of a machine, comprising:a seal carrier element;
a clamping body; and
a fixing body that is disposed between the seal carrier element and the clamping body,
wherein the fixing body is connected integrally to the seal carrier element and the clamping body by bar-like connections that are detachable when subjected to a force; and
wherein when disengaging the bar-like connections between the fixing body, the seal carrier element, and the clamping body, a force-fit results between the fixing body, the seal carrier element, the clamping body, and side walls of the recess of the machine.

US Pat. No. 10,280,775

GUIDE VANE RING FOR A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A guide vane ring for a turbomachine with at least one guide vane ring segment comprising at least one guide vane arranged radially around an axis of rotation, an outer shroud arranged radially on the outer side of the guide vane, and an inner shroud arranged radially on the inner side of the guide vane, with the outer shroud or the inner shroud having at least one expansion joint, wherein the guide vane ring segment with the guide vane, the outer shroud, and the inner shroud is formed in one piece and the at least one expansion joint is sealed by at least one sealing device arranged in the expansion joint,wherein the sealing device has an elastically and/or plastically deformable sealing element, which is configured in one piece with the outer shroud or inner shroud having the expansion joints.

US Pat. No. 10,247,005

BLADE OR VANE ARRANGEMENT FOR A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A blade or vane arrangement for a turbomachine, comprising:a turbomachine blade or vane including a recess defined by at least one recess wall; and
at least one tuner guide housing with at least one cavity, in which at least one tuner that is provided for impact contact with the tuner guide housing is taken up, and a locking plate joined to the tuner guide housing,
wherein the tuner guide housing is fastened in the recess of the turbomachine blade or vane by the locking plate in a form-fitting or friction-fitting manner,
wherein the cavity of the tuner guide housing in which the tuner is taken up, is closed in a gas-tight manner, and
wherein the locking plate is in direct contact with the tuner guide housing and the at least one recess wall.

US Pat. No. 10,240,472

BRUSH SEAL FOR A TURBOMACHINE

MTU Aero Engines AG, Mun...

1. A brush seal for a turbomachine, comprising:a carrier that has a recess in which a spring element is axially braced that immobilizes a brush element, the spring element being axially latched against an undercut in the recess,
wherein the spring element comprises
a first leg having a first end and a second end,
a second leg having a first end and a second end, wherein the first end of the second leg is interconnected with said second end of said first leg,
a third leg having a first end and a second end, wherein the first end of the third leg is interconnected with said second end of said second leg, and
a fourth leg having a first end and a second end, wherein the first end of the fourth leg is interconnected with said second end of said third leg;
the brush element being clamped between the first and second legs, wherein the fourth leg braces against the undercut.

US Pat. No. 10,240,474

TURBOMACHINE HAVING A SEAL DEVICE

MTU Aero Engines AG, Mun...

1. A turbomachine having a seal device for sealing a gap between two stator components, having a plurality of sealing elements that are mounted on one of the stator components and bridge the gap, and having a plurality of prestressing elements, by which the sealing elements are prestressed against a respective sealing edge of the stator components, wherein the prestressing elements are disposed on the stator component on the mounting side, and the sealing edge of the stator component on the mounting side is next to the mounting and the sealing edge of the stator component without the mounting is far from the mounting, wherein the sealing edge next to the mounting is disposed between the mounting and the sealing edge far from the mounting, and wherein the prestressing elements engage, between the sealing edges, on the sealing elements near the sealing edge next to the mounting and at a distance from the sealing edge far from the mounting, wherein the prestressing elements are each supported at a front support point and a back support point of the stator component on the mounting side, which are distanced from one another in the axial direction and radial direction of the turbomachine, wherein the front support point is disposed radially outside and the back support point is disposed radially inside with respect to the mounting, wherein the front support point is upstream of the back support point along a direction of gas flow.

US Pat. No. 10,208,616

TURBOMACHINE WITH BLADES HAVING BLADE TIPS LOWERING TOWARDS THE TRAILING EDGE

MTU AERO ENGINES AG, Mun...

1. A turbomachine, comprising:a rotor, which is mounted rotatably around its longitudinal axis in a stator, and which has at least one row of rotating blades, which is formed by a plurality of rotating blades,
the stator having at least one abradable layer,
each rotating blade of the plurality of rotating blades having blade tips comprising a front blade tip region, at least one middle blade tip region, and a trailing blade tip region that are lowered radially inward in sections from a leading edge of the rotating blade to a trailing edge of the rotating blade,
wherein, during operation, the front blade tip region and the at least one middle blade tip region of each of the rotating blades runs into, and is in contact with, the at least one abradable layer, and the trailing blade tip region is radially distanced from the at least one abradable layer, forming a radial gap, whereby formation of a parasitic gap is prevented,
wherein the front blade tip region extends from the leading edge of the rotating blade to the at least one middle blade tip region in a direction towards the trailing edge of the rotating blade,
wherein the trailing blade tip region extends from the at least one middle blade tip region to the trailing edge of the rotating blade, and
wherein radial dimensions of the front blade region are greater than radial dimensions of the at least one middle blade tip region, and the radial dimensions of the at least one middle blade tip region are greater than radial dimensions of the trailing tip region.

US Pat. No. 10,201,845

METHOD FOR MANUFACTURING BRUSH SEALS WITH OBLIQUELY POSITIONED BRISTLES AND A CORRESPONDING DEVICE

MTU Aero Engines AG, Mun...

1. A method for manufacturing brush seals with obliquely positioned bristles, comprising the steps of:a.) winding a metal filament or wire comprising a material for bristles over two wire cores disposed at a distance from and running parallel to one another, to form a tightly packed metal filament or wire package;
b.) fastening the metal filament or wire package to at least one wire core;
c.) local heating of the region around at least one of the two wire cores by at least one induction coil disposed in at least one fastening jaw; and
d.) displacing at least one of the two wire cores so as to obtain a brush seal with obliquely positioned bristles made of the metal bristle material with the tightly packed metal filament or wire package.

US Pat. No. 10,494,954

CONNECTION SYSTEM FOR HOUSING ELEMENTS OF A TURBINE INTERMEDIATE CASING

MTU Aero Engines AG, Mun...

1. A connection system for a hot gas-conducting annular duct of a turbine intermediate casing of a gas turbine, the connection system comprising:a first housing element;
a second housing element situated next to the first housing element in a circumferential direction; and
a fastening unit configured for connecting the first housing element and the second housing element to one another at adjacent edges of the first housing element and of the second housing element adjacent to one another in the circumferential direction, the fastening unit including a clamp mounted on the second housing element with the aid of a connection, the clamp resting on a clamping surface provided on the first housing element in such a way that the first housing element is accommodated between the clamp and the second housing element;
a top side of the clamping surface facing the clamp has a clamping surface contour, relative to a longitudinal section extending along the adjacent edges, designed in such a way to thwart a rotation of the clamp about an axis of the connection.

US Pat. No. 10,436,052

LEAF SEAL FOR SEALING OFF A SHAFT ROTATING AROUND AN AXIS

MTU Aero Engines AG, Mun...

1. A leaf seal to seal off a shaft rotating around an axis, comprising:a plurality of leaves arranged spaced apart from one another; and
an outer shroud that supports the plurality of leaves; wherein
the plurality of leaves are formed together with the outer shroud as a unitary element by a generative production process; and
the plurality of leaves have recesses forming first end areas of reduced cross-sectional thicknesses defining notches adjacent the outer shroud, and second end areas of increased cross-sectional thicknesses, located opposite the first end areas, defining spacer elements adjacent ends of the plurality of leaves situated away from the outer shroud.