Abstract: A shackle for fastening an aircraft engine including a ball joint with a pin including a ring including a first bore and a sleeve assembled in the first bore and having a second bore configured to receive the pin and the diameter of which progressively increases to have a flared inner surface at each of its ends, the fastening shackle being arranged such that a spherical outer surface of the ring bears an anti-friction coating produced on a region extending beyond the contact region with the body of the shackle, this first region being bordered by rims forming a projection around the coating, and/or the second bore bears an anti-friction coating on a region extending beyond its contact region with the pin, this region being bordered by rims forming a projection around this anti-friction coating, at the ends of the flared inner surface.

Abstract: A support structure for attaching an engine to an aircraft pylon at front, mid and rear attachment positions thereof, including a front mount joined to the engine and configured to attach to the pylon at the front attachment position and a rear mount joined to a core casing to attach to the pylon at the rear attachment position, each of the front and rear mounts configured to transfer lateral and vertical loads from the engine to the pylon, and the rear mount being spaced from the front mount such that yaw and pitch torques are transferred from the engine to the pylon through the front and rear mounts. The support structure also includes an axial load transfer formation to transfer axial loads from the engine to the pylon and a roll-torque transfer formation to transfer roll torque from the core casing to the pylon.

Abstract: The present invention discloses a system for providing mobile power, in which the required equipment for the power supply system at fracturing fields as well as connection cables and connection hoses are integrated properly, assigned onto three transport vehicles for movement and effectively connected. Intake components and a turbine generation system are combined on a first transport vehicle and installed together, then transported to customer sites directly, thus saving the installation time at the user sites. The two different designs on the locations of an exhaust stack and an exhaust silencer not only meet the requirements of road transportation, but also meet the requirements of exhaust gas emission during operations.

Abstract: A system for mounting an engine on an aircraft. The system includes a lower forward element, lateral linkages, an upper forward element, and an aft element. The lower forward element couples a forward portion of an engine core with a lower forward portion of a pylon. The linkages extend laterally between the lower forward element and the forward portion of the engine core. The upper forward element couples an outer portion of the fan case with an upper portion of the pylon. The aft element is spaced aftwardly apart from the lower and upper forward elements, and couples an aft portion of the engine core to an aft portion of the pylon. The system reacts all major moments acting on the engine, and the system does not extend beyond an outer periphery of the fan case, which eliminates the need for or minimizes the height of an aerodynamic fairing.

Abstract: A combustor dome system includes an annular combustor dome defining a main axis therethrough. The combustor dome includes opposed upstream and downstream faces, wherein the upstream face is configured to face upstream toward a compressor discharge space, wherein the downstream face is configured to face downstream toward a combustor space. The downstream face has a curved cross-sectional profile. A plurality of nozzles extends at least partially through the combustor dome from the upstream face to the downstream face for injection of fuel into the combustor space. A fuel manifold is in fluid communication with the plurality of nozzles.

Abstract: Dual function links for gas turbine engine mounts are disclosed. An example apparatus for mounting a casing of a gas turbine engine to a pylon, the apparatus includes a first pin and a linkage including a first linkage portion including the first pin, a second linkage portion, the second linkage portion forming a first load path between the gas turbine engine and the pylon, the second linkage portion arranged with respect to the first linkage portion such that a gap prevents the first linkage portion from forming a second load path between the gas turbine engine and the pylon.

Abstract: A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine including: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a working gas flowpath and a fan duct flowpath; a primary fan driven by the turbomachine defining a primary fan tip radius R1 and a primary fan hub radius R2; a secondary fan located downstream of the primary fan and driven by the turbomachine, at least a portion of an airflow from the primary fan configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide a fan duct airflow through the fan duct flowpath during operation to generate a fan duct thrust, wherein the fan duct thrust is equal to % Fn3S of a total engine thrust during operation of the gas turbine engine at a rated speed during standard day operating conditions; wherein a ra

Abstract: An engine mounting pylon for suspending a turbofan beneath an aircraft wing, including a primary structure, a fastening device for fastening to the turbofan, and a fastening device for fastening to the wing. The engine mounting pylon furthermore comprises multiple rigid braces that mechanically connect the primary structure to the wing of the aircraft, a first brace-attachment device that mechanically connects a first end of each brace to the wing, and a second brace-attachment device that mechanically connects a second end of each brace to the primary structure. This braced-pylon architecture makes it possible to reduce the dimensions of the primary structure and the various elements that constitute the first and second attachment devices.

Abstract: A gas turbine engine includes, among other things, a propulsor section including a rotor, a gear train, a low spool and a high spool. A static structure includes a first case and a second case. A mount system includes a forward mount and an aft mount arranged in a plane containing an engine axis of rotation. The forward mount is secured to the first case. The aft mount is secured to the second case.

Abstract: A turbine vane assembly adapted for use in a gas turbine engine includes a plurality of turbine vanes, an outer vane support, and an inner vane support. The plurality of turbine vanes comprise ceramic matrix composite material and are adapted to interact with hot gases flowing through a gas path of the gas turbine engine during use of the turbine vane assembly. A spar system couples the inner vane support with the outer vane support.

Abstract: A system for providing mobile electric power includes a first transport including an inlet plenum and a generator, a second transport including a gas turbine, and a turbine exhaust unit, and a third transport including a generator exhaust air handling system, a combustion air handling system, and a turbine enclosure air handling system. The first transport, the second transport, and the third transport are separate transports that are independently movable in a transportation mode. In an operation mode, the first transport and the second transport are connected to each other such that a first end side of the first transport faces and is connected to a first end side of the second transport, and the inlet plenum mounted on the first transport is connected to an intake of the gas turbine mounted on the second transport.

Abstract: A linking device connecting an aircraft engine and a primary structure of an aircraft pylon, comprising a support connected to the primary structure, a spreader having a central part connected to the support and a first and a second end that are connected to thrust rods, at least one limiting system for limiting the travel of the spreader having an upper and a lower stop between which is positioned the first or second end of the spreader, the upper and the lower stop being connected to one another so as to form a single part connected to a fitting secured to the support.

Abstract: An integrated propulsion system according to an example of the present disclosure includes, among other things, a fan section, a gas turbine engine, a geared architecture, a nacelle assembly and a mounting assembly. The nacelle assembly includes a fan nacelle and an aft nacelle, the fan nacelle arranged at least partially about a fan and the engine, and the fan nacelle arranged at least partially about a core cowling to define a bypass flow path.

Abstract: A method for installing a gas turbine assembly of a first type at a position of an existing power plant where previously a gas turbine assembly of a second type was installed on a foundation specially designed for said second type. The gas turbine assembly includes at least one housing, a compressor, a combustion chamber, a gas turbine, and a plurality of venting and removal lines guided along the exterior of the housing. Modifications to the venting and/or removal lines of the gas turbine assembly of the first type are carried out in a first step, and the modified gas turbine assembly is installed on the existing foundation in a second step.

Abstract: A junction of a pylon with an aircraft wing, the pylon including a primary structure extending from front to rear along a longitudinal axis and in the form of a box with a rear face and an upper spar forming an upper face of the box, a longitudinal median plane separating the primary structure into two parts, left and right, the junction comprising a rear attachment system for attaching the pylon to the wing, this system being arranged at the rear of the primary structure, the rear attachment system including a shoe attached beneath the wing, the shoe being connected, via at least one articulated connecting rod, to a fitting attached to the rear face of the pylon by an articulation whose articulation pin, termed the horizontal articulation pin, is perpendicular to the longitudinal median plane, and the shoe being connected to the upper spar via a force reacting system.

Abstract: A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine including: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a working gas flowpath and a fan duct flowpath; a primary fan driven by the turbomachine defining a primary fan tip radius R1 and a primary fan hub radius R2; a secondary fan located downstream of the primary fan and driven by the turbomachine, at least a portion of an airflow from the primary fan configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R3 and a secondary fan hub radius R4, wherein the secondary fan is configured to provide a fan duct airflow through the fan duct flowpath during operation to generate a fan duct thrust, wherein the fan duct thrust is equal to % Fn3S of a total engine thrust during operation of the gas turbine engine at a rated speed during standard day operating conditions; wherein a ra

Abstract: A vane arc segment includes an airfoil fairing that has first and second fairing platforms and a hollow airfoil section extending there between. There is a spar having a spar platform adjacent the first fairing platform and a spar leg that extends from the spar platform and through the hollow airfoil section. The spar leg has an end portion with a clevis mount that that protrudes from the second fairing platform. A support platform adjacent the second fairing platform has first and second through-holes. The end portion of the spar leg extends through the first through-hole and a pin extends though the clevis mount to lock the support platform to the spar leg. A stabilizer rod extends through the second through-hole of the support platform and limits rotation of the support platform about the pin.

Abstract: A rear engine attachment having a beam fixed to a pylon, a spreader pivotally mounted on the beam, the spreader comprising two ends situated on either side of a median plane, two rods located one on either side of the median plane, each rod having a first end fixed to one end of the spreader by a ball joint and a second end fixed to the engine, each ball joint comprising a clevis with two arms at the first end of the rod, a swivel bearing arranged on the end of the spreader, and a hinge pin passing through the two arms, the swivel bearing arranged between the two arms and fitted on the hinge pin. For each ball joint, an interface part is removably interposed between the end of the spreader and the first end of the rod and forms a contact surface between the spreader and the rod.

Abstract: A gas turbine engine for an aircraft is provided. The gas turbine engine comprises an engine core comprising a compressor, a combustor, a turbine, and a core shaft connecting the turbine to the compressor. The gas turbine engine further comprises a fan located upstream of the engine core, the fan comprising a plurality of fan blades, the fan generating a core airflow which enters the engine core and a bypass airflow which flows through a bypass duct surrounding the engine core. The gas turbine engine further comprises a circumferential row of outer guide vanes located in the bypass duct rearwards of the fan, the outer guide vanes extending radially outwardly from an inner ring which defines a radially inner surface of the bypass duct.

Abstract: An assembly for an aircraft, which includes a pylon with a bearing face and a contact face, an engine mount with a bearing face and a contact face bearing against the contact face of the pylon, at least two bolts comprising a threaded stem with a head and a nut bearing against the bearing face of the engine mount, for each head, a bearing washer with a first bearing face and a second bearing face, where the second bearing face of the bearing washer bears against the bearing face of the pylon, and where the head bears against the first bearing face of the bearing washer, a shoe, and, for each bearing washer, a clamp that is secured to the shoe and comprises a first branch and a second branch between which the bearing washer is clamped and is free to rotate.

Abstract: A connection assembly for mounting an engine to a mounting structure. The engine is rotatable about an axis of rotation and defines an axial direction extending along the axis of rotation from a forward end to an aft end. The engine has a center of gravity. The connection assembly includes: an engine coupling piece coupled to the engine; and a connection piece connecting the engine coupling piece and the mounting structure, the connection piece being inclined toward the axial direction. A mounting system including the connection assembly is also described.

Abstract: A gas turbine engine turbine section has tie rod assemblies interconnecting an inner diameter structure and an outer casing. Each tie rod has: an inner diameter end; an outer diameter end; and an eyelet of an outer diameter spherical bearing formed at the outer diameter end. A first clevis carries a spherical ball of the bearing, a shank of said clevis extending to an outer diameter (OD) end. A tensioning bolt is mated to a threaded opening in said clevis OD end whereby tightening said bolt applies a tension to said rod. A radial span between a center of the outer diameter spherical bearing and an inner diameter surface of the outer casing is at least 50% greater than that between an outer diameter (OD) surface of the outer ring and the center.

Abstract: An accessory system for a gas turbine engine having a driveshaft with an axis of rotation is provided. The system includes a towershaft coupled to the driveshaft and rotatable about a towershaft axis of rotation. The towershaft includes a towershaft bevel gear. The system includes a primary shaft including a first bevel gear and a second bevel gear that each revolve about a primary shaft axis of rotation. The first bevel gear is coupled to the towershaft bevel gear. The system includes a secondary shaft including a third bevel gear and a fourth bevel gear that each revolve about a secondary shaft axis of rotation. The third bevel gear is coupled to the second bevel gear. The system includes a tertiary shaft including a fifth bevel gear that revolves about a tertiary shaft axis of rotation. The fifth bevel gear is coupled to the fourth bevel gear.

Abstract: A compound engine assembly with an engine core including at least one internal combustion engine, a turbine section including a turbine shaft in driving engagement with the engine shaft, and a compressor, and a firewall. The compressor is located on one side of the firewall, and the turbine section and the engine core are located on the other side. The assembly may include a gearbox module with the turbine section and the engine core located on a same side of the gearbox module casing and the compressor located on the opposite side of the gearbox module casing, and with the firewall extending from the gearbox module casing. One or more rotatable accessory may be located on a same side of the firewall as the compressor. A method of reducing fire hazard in a compound engine assembly is also discussed.

Abstract: A system and methods for reducing flow field velocity into a propeller is disclosed. An aerodynamic structure is enhanced to provide a reduced velocity flow field region, and a propeller is positioned in the reduced velocity flow field region.

Abstract: The present invention provides an electrical raft comprising a rigid material having multiple electrical conductors embedded therein. The electrical raft comprises an electrical connector having a first set of electrical contacts connected to at least one of the electrical conductors, a second set of electrical contacts for electrical connection to another component, and a housing having a first end and a second end. The first set of electrical contacts is at the first end and the second set of electrical contacts is at the second end. The first end of the housing is at least partly embedded in the rigid material, and the second end of the housing is accessible from outside the electrical raft, allowing the electrical raft to be electrically connected to said other component. The electrical connector has a back-shell which sealingly encloses the first end of the housing to prevent ingress of contaminants into the connector.

Abstract: An electrical raft (200) is provided that has electrical conductors (252) embedded in a rigid material (220). The electrical raft is provided with an electrical connector (700). The electrical connector (700) is mounted in the electrical raft (200) at a mounting angle (730). The mounting angle is set such that the electrical conductors (252) can be connected to the electrical connector without having to turn through an angle or a radius of curvature that would subject them to excessive bending stress. Similarly, the mounting angle means that any conductors (766) that may be connected to the electrical connector (700) do not have to turn through an angle or a radius of curvature that would subject them to excessive bending stress. Furthermore, mounting the electrical connector (700) at the mounting angle (730) may allow the assembled electrical raft (200) to be more compact.

Abstract: A method of deploying a modular gas compression plant is disclosed. In an embodiment, the method comprises constructing a building of four substructures including a first lower housing, a first upper housing, a second lower housing, and a second upper housing. The method further includes assembling a turbomachinery equipment into the first lower housing. The method further includes assembling an air inlet duct and an exhaust duct to the first upper housing. The method further includes disassembling the building into the four substructures.

Abstract: Gas turbine engine frames are disclosed. An example gas turbine engine frame may include a generally annular outer casing disposed coaxially about a hub; a plurality of circumferentially spaced apart struts joined to the hub and the outer casing, individual struts extending radially outwardly from the hub to the outer casing; and a stiffening rail monolithically formed with the outer casing circumferentially between two of the struts. The stiffening rail may extend radially inward beyond the inner surface of the outer casing between the struts.

Abstract: An assembly for a gas turbine engine includes a frame, a mount, and a fairing. The mount is attached to the frame and the fairing is connected to the mount. The fairing and mount have mating anti-deflection features that engage to prevent circumferential movement of the fairing relative to the frame.

Abstract: Undue flexing and cracking of an internal exhaust shield of a turbine is eliminated by isolating thermally-induced movement in the shield from that of strut apparatus passing therethrough to a surrounding outer exhaust casing. Methods and apparatus are disclosed.

Abstract: An internal wall of an air intake and a housing of a fan form a single tubular piece made from a fiber/resin composite material. A rear end has the single tubular component connected by a flange and bolts to an external cowl of the fan duct, such that the rear end of the tubular component is uniform and the flange is attached and fixed to the rear end.

Abstract: An assembly and method for attaching an engine pylon to an aircraft wing. The method may include attaching a first forward wing-mounted fitting to the pylon at a first side of the pylon between upper and lower spars of the pylon and attaching a second forward wing-mounted fitting to the pylon at a second side of the pylon between the upper and lower spars of the pylon. The second side of the pylon is located opposite of the first side of the pylon. Mechanical attachment devices may be inserted through holes on the sides of the pylon aligned with holes formed through the forward wing-mounted fittings. The mechanical attachment devices may be made of steel and the forward wing-mounted fittings may be made of composite or aluminum material.

Abstract: A forward engine mount assembly for a gas turbine engine includes a mount beam having a main body with a fore end and an aft end and a forward shackle assembly supported by the fore end of the mount beam. The forward shackle assembly comprises a first link configured to be connected to a first engine case structure and a second link configured to be connected to a second engine case structure. The first and second links are pivotally attached to each other.

Abstract: A static structure of a gas turbine engine according to an exemplary aspect of the present disclosure includes a multiple of airfoil sections between an outer ring and an inner ring. A spring biased tie-rod assembly is mounted through at least one of the multiple of airfoil sections.

Abstract: A bypass duct of a gas turbine engine has an inner bypass duct wall having a front end portion which is rigidly mounted to an engine core case and a rear end portion which is flexibly mounted to the core case. The flexible mounting between the rear portion of the inner bypass duct wall and the core case allows hot engine areas of the core case to thermally grow and contract relative to the inner bypass duct wall without imposing additional loads on the inner bypass duct wall.

Abstract: A disclosed gear assembly support for a gas turbine engine includes a first portion configured for attachment to a case of the gas turbine engine and a second portion configured for supporting a gear assembly. The support further includes a snap portion defining a fit within the case. The snap portion includes a tunable feature for adjusting a fit within the case. A torque reacting portion of the support transfers torque from the second portion to the first portion separate from the snap portion and include separately tunable features for adjusting the snap fit independent of the torque transfer portions.

Abstract: A fuel control module for a gas turbine engine and method of installation. The fuel control module includes a first frame unit, a second frame unit, a fuel path, and a fuel controller components. The first frame unit and the second frame unit include features for installation in multiple installations and/or multiple configurations.

Abstract: A split fuel control module for a gas turbine engine and method of installation. The split fuel control module includes a first frame unit, a second frame unit, a segmented fuel path, and a distributed fuel controller. The first frame unit and the second frame unit are joined together at a frame unit interface. The segmented fuel path includes an upstream fuel interface fixed to the first frame unit and a downstream fuel interface fixed to the second frame unit and detachably coupled to the upstream fuel interface at the frame unit interface. A first portion of the distributed fuel controller is fixed to the first frame unit, and a second portion of the distributed fuel controller is fixed to the second frame unit.

Abstract: Provided is a tail cylinder attaching and detaching fixture that attaches and detaches a tail cylinder to and from a casing, the tail cylinder being included in a combustor inserted into the casing so that the front end is connected to an inlet portion of a combustion gas passageway, the tail cylinder attaching and detaching fixture including: a guide portion of which the front end is disposed inside the casing and the front end and the base end are respectively supported by the casing and which supports the tail cylinder so as to be movable in the axial direction of the combustor; and an advancing and retracting mechanism that advances and retracts the tail cylinder in the axial direction.

Abstract: A preassembled modular drop-in combustor having internal components in conformity with assembly and function specifications prior to and after insertion into an industrial gas turbine access port and internal combustor transition. The combustor assembly maintains conformity with those specifications after insertion into the combustor case if it does not inadvertently impact other turbine components during its installation. Inadvertent impact is avoided by having a combustor service zone proximal the combustor case, enabling slidable insertion of each combustor assembly into its corresponding access port and transition along its corresponding insertion path without contacting other turbine system components. A multi-axis motion combustor handling tool in the combustor service zone, preferably under automatic control, is coupled to each combustor and facilitates precise alignment along the insertion path.

Abstract: A mounting assembly for attaching a ducted fan gas turbine engine that includes an intake, a propulsive fan, a fan case surrounding the fan, and a core engine. The air intake is attached to the front of the fan case such that loads acting on the air intake are transmitted to the fan case. The mounting assembly includes: a support structure extending in an axial direction of the engine and having a rearward region adapted to attach to the aircraft; and a load distribution ring coaxial with and rearward of the fan case, adapted to join to the fan case, and joined to a forward region of the support structure. The support structure and the load distribution ring are adapted such that the primary load path for the loads transmitted to the fan case by the air intake is through the load distribution ring and the support structure.

Abstract: An aircraft engine assembly including an attachment mechanism of a turbojet onto a rigid structure of a mounting strut, including a first, second, and third forward engine attachment for taking up thrust loads brought onto the fan frame, and arranged such that the third attachment passes through a first diametrical plane of the turbojet, the first and second attachments being disposed on one side and another of the first plane. The first and second forward engine attachments are respectively brought to the fan frame at two points, situated beyond the second diametrical plane of the turbojet, orthogonal to the first diametrical plane, with respect to the third attachment.

Abstract: A system for packaging a gas turbine includes an enclosure peripherally surrounding and extending along a length of the gas turbine. The enclosure comprises first and second ends with a passage between the enclosure and the gas turbine. The system further includes means for connecting the first and second ends to an air filtration system and an exhaust system, respectively. A plurality of supports are connected to and extend radially inward from the enclosure with means for connecting the plurality of supports to the gas turbine. A method for packaging a gas turbine includes connecting the gas turbine to a plurality of supports and connecting the plurality of supports to an enclosure peripherally surrounding and extending along a length of the gas turbine. The method further includes connecting first and second ends to an air filtration system and an exhaust system, respectively.

Abstract: An engine strut for providing fan hub frame structural support and monitoring an air flow within an aircraft engine includes an airfoil coupled to the aircraft engine and has a first portion and a second portion. The first portion is positioned upstream of the second portion with respect to the air flow. A shield is coupled to the engine and positioned between the first portion and the second portion. The shield includes a first side spaced from the first portion and defining a first flow path with the first portion. The shield further includes a second side spaced from the second portion and defining a second flow path with the second portion. At least one sensor is coupled to the aircraft engine and positioned in flow communication with the second flow path.

Abstract: An assembly for mounting a turbine engine to a pylon includes a mounting beam, a plurality of fasteners, a first mounting linkage and a second mounting linkage. The mounting beam includes a mount beam fitting that extends axially between a first mount beam end and a second mount beam end, and a mount beam flange that extends radially out from the mount beam fitting at the first mount beam end. A first fastener aperture extends radially through the mount beam fitting, and a second fastener aperture extends axially through the mount beam flange. The fasteners connect the mounting beam to the pylon. A first of the fasteners is mated with the first fastener aperture, and a second of the fasteners is mated with the second fastener aperture. The first mounting linkage connects the first mount beam end to a first engine attachment. The second mounting linkage connects the second mount beam end to second engine attachments.

Abstract: The invention relates to a device (1) for suspending a casing (2) of a turbine engine (3) to the structure (12) of an aircraft, characterized in that it comprises a suspension ring (4), said suspension ring (4) comprising; a plurality of attachment surfaces (33) positioned so as to protrude on the circumference of the suspension ring (4) and having attachment holes (6), for attachment to the structure (12), a circumferential flange (5) comprising a plurality of holes (35) for attaching the suspension ring (4) to the casing (2); one or several hanging surfaces (10) for hanging at least one piece of equipment (15) , positioned so as to protrude on the circumference of the suspension ring (4), and having holes (8) for hanging the piece of equipment (15) to the ring (4). The invention also relates to a turbine engine and to a propulsion assembly comprising such a device.

Abstract: A fuel control module including a gas vent manifold is disclosed. The gas vent manifold includes a first gas vent portion including a first gas vent connection and a second gas vent portion including a second gas vent connection. The gas vent manifold also includes a condensation portion extending between the first gas vent portion and the second gas vent portion. The gas vent manifold further includes a condensation drain port fluidly connected to the condensation portion. The gas vent manifold also includes a plurality of connection ports. Each connection port is configured to fluidly couple the gas vent manifold to a vent line of one or more pneumatic valves.

Abstract: An adapting part is designed to be held on a casing of a turbo-engine, partially covering the casing. The adapting part includes, in the extension of one of its axial ends, a first connector to engage with a first complementary connector associated with the casing to form a sliding connection. The adapting part also includes, at its other axial end, a second connector to be secured to a second complementary connector associated with the casing to form a rigid connection.

Abstract: The invention relates to a fixed turbine engine receiver part comprising a fixed hollow shaft for carrying ancillary systems of a turbine engine, centered on the turbine engine axis, turbine engine ancillary systems, an assembly for holding ancillary systems in position situated inside the fixed hollow shaft. The assembly for holding ancillary systems in position comprises at least one first ancillary system support ring, having an axis substantially parallel with the turbine engine axis, a first distance sleeve for holding the first ancillary system support ring in position, having an axis substantially parallel with the turbine engine axis, the first distance sleeve bearing against the first support ring, such that the first support ring comprises a plurality of mutually separated through holes, each through hole defining a passage for at least one ancillary system, each through hole being traversed by at least one of the ancillary systems of the turbine engine.