A turbine engine rotor wheel including a disk including at least one slot formed in its outer periphery for mounting roots of blades, a presser being mounted between each blade root and a bottom of the slot. The presser is bistable in position and is capable of occupying a first stable position for assembly and disassembly in which it does not exert a force on the blade root, and a second stable position in which it exerts a radial force on the blade root to hold the blade stationary and to stabilize the blade in a final position.

The invention proposes a fan, in particular for a turbomachine of small size such as a jet engine, having a hub ratio which corresponds to the ratio of the diameter of the inner limit of the incoming air stream (26) at the radially inner ends of the leading edges of the fan blades (10), divided by the diameter of the circle around which the outer ends of the fan blades pass, having a value of between 0.20 and 0.265.

A securement system for securing an attachment portion of at least one rotor blade or rotor blade segment in a mounting site of a rotor base. The securement system has at least one holder element and a second holder element, each of which has a holding portion and a connecting portion. The connecting portions can be introduced into the mounting site so as to establish a clip connection and can be locked together in the mounting site. The holding portions are configured so as to fasten the attachment portion of the at least one rotor blade or of the rotor blade segment in the mounting site in the locked state of the connecting portions. Further disclosed is a method for securing an attachment portion of at least one rotor blade or rotor blade segment in a mounting site of a rotor base and a rotor having a securement system.

A compressor section includes a compressor disk having a slot receiving a blade with an airfoil extending from a platform and the compressor disk having an upstream end and a downstream end. A cover plate at one of the upstream end and the downstream end covers an area between the blade and the compressor disk across a circumference of the compressor disk. The cover plate covers an area between an inner peripheral surface of the platform to a radially innermost end of the slot in the disk. The compressor disk has hooks, and there is a retention ring mounted in a cavity between the hooks and the compressor disk. The cover plate is mounted between the retention ring and the compressor disk. The retention ring retains the cover plate against the compressor disk. A gas turbine engine is also disclosed.

An aircraft gas turbine with a fan disc at which fan blades, distributed around the circumference and forming an intermediate space in between each other, are attached, with a sealing disc that is arranged on the back side of the fan disc, and with an inlet cone that is mounted at the front side of the fan disc, as well as with filling elements that are arranged in the intermediate spaces, wherein the inlet cone includes strip-shaped lugs which are configured in one piece with the same and which can be inserted into the intermediate spaces, with their free ends being inserted into a ring groove that is formed at the sealing disc, and in that a filling element is arranged on each strip-shaped lug.

A variable pitch fan for a gas turbine engine includes a fan blade defining a pitch axis and attached at a radially inner end to a trunnion mechanism. The variable pitch fan also includes a disk having a disk segment with the trunnion mechanism extending at least partially through the disk segment. A key is positioned at least partially in a key slot defined in a base of the trunnion mechanism, and further positioned adjacent to a support member of the disk segment. The key defines a first contact line between the key and the key slot and a second contact line between the key and the support member. The first and second contact lines respectively define a first contact reference line in a second contact reference line. The first and second contact reference lines define an angle with the pitch axis of the fan blade greater than zero degrees and less than ninety degrees.

A fan assembly for an aviation turbine engine, the assembly including a fan disk having at least one tooth and at least one platform mounted on the tooth of the fan disk, the platform including a box of composite material made from fiber reinforcement densified by a matrix, the box having a flow passage wall, a bottom wall, and two side walls extending radially between the bottom wall and the flow passage wall. The box of the platform includes an upstream opening at an upstream end of the platform and a downstream opening, and the assembly also includes a locking key housed in the box and passing through the upstream and downstream openings of the box, the locking key being blocked at each of its ends by a blocking element.

A turbine engine having a disk, a composite airfoil assembly and a pitch snubber. The disk having a slot. The composite airfoil assembly having an airfoil portion and a dovetail portion. The dovetail portion having a radially inner surface spaced from the slot to define a gap therebetween. The pitch snubber being provided on one of the slot or the radially inner surface of the dovetail portion.

A radial turbine rotor incorporating ceramic matrix composite turbine blades is disclosed. The radial turbine rotor can include a dovetail shape retention features for coupling the ceramic matrix composite turbine blades to a central hub.

One aspect of present application provides an intermediate structure in a gas turbine engine. The intermediate structure is positioned between a first component and another component. The first component may be a composite component. The components may be interlocking. The intermediate structure may be load bearing. Also disclosed is a method using the intermediate structure.