Didier Hernandez

Bio: Didier Hernandez is an academic researcher from Snecma. The author has contributed to research in topics: Combustion chamber & Casing. The author has an hindex of 20, co-authored 206 publications receiving 1465 citations. Previous affiliations of Didier Hernandez include Office National d'Études et de Recherches Aérospatiales.

Mounting a turbine nozzle on a combustion chamber having CMC walls in a gas turbine

Abstract: The gas turbine comprises an annular combustion chamber having inner and outer walls made of ceramic matrix composite material, and a high pressure turbine nozzle secured to a downstream end of the combustion chamber and comprising a plurality of stationary airfoils extending between the inner and outer walls of an annular flow path through the nozzle for the gas stream coming from the combustion chamber. The turbine nozzle is made of ceramic matrix composite material and it is connected to the downstream end of the combustion chamber by brazing.

Fastening a CMC combustion chamber in a turbomachine using brazed tabs

Abstract: A turbomachine has inner and outer annular shells of metal material containing, in a gas flow direction F, a fuel injector assembly, an annular combustion chamber of composite material, and an annular nozzle of metal material forming the fixed-blade inlet stage of a high pressure turbine. Provision is made for the combustion chamber to be held in position between the inner and outer metal annular shells by a plurality of flexible metal tabs having first ends interconnected by a metal ring fixed securely to each of the annular shells by first fixing means, and second ends fixed by second fixing means on a ring of composite material fixed securely to said composite material combustion chamber, the flexibility of said metal fixing tongues allowing expansion to take place freely in a radial direction at high temperatures between said composite material combustion chamber and said metal annular shells.

Annular combustion chamber of a turbomachine

Abstract: Annular combustion chamber of a turbomachine is provided. The combustion chamber includes an inner wall, an outer wall, a chamber bottom disposed between the walls in the upstream region of the chamber, and two attachment flanges disposed downstream of the chamber bottom and respectively enabling the walls to be attached to other parts of the turbomachine. Each wall is divided into several adjacent sectors and each sector is attached to the chamber bottom and to one of the attachment flanges. Advantageously, the adjacent sectors overlap at their lateral edges and there exists a degree of radial play between two adjacent sectors. In addition, the lateral edges of the sectors are inclined circumferentially relative to the principal axis of the combustion chamber.

Turbojet engine combustion chamber with a double wall converging zone

Abstract: The combustion chamber according to the present invention utilizes a double-walled structure only in the converging portion of the combustion chamber. A hot wall is retained adjacent to, but spaced from a corresponding cold wall so as to provide a cooling space between them. An upstream portion of the hot wall is retained in a notch defined by the cold wall, while the downstream portions each have generally radially extending flanges. Studs pass through openings in the adjacent flanges to prevent any relative circumferential movement, while a clamp serves to attach the flanges to the inlet of a turbine. The clamp slidably retains the flanges therein so as to permit relative radial movement to accommodate any thermal expansion and contraction of the combustion chamber, solely by the appended claims.

Fastening a CMC combustion chamber in a turbomachine using the dilution holes

Abstract: In a turbomachine comprising inner and outer annular shells of metal material containing, in a gas flow direction F: a fuel injection assembly; an annular combustion chamber of composite material having a plurality of air feed orifices in a primary zone and/or in a dilution zone; and an annular nozzle of metal material forming the inlet stage with fixed blades for a high pressure turbine, provision is made for the combustion chamber to be held in position between the inner and outer annular metal shells by means of a plurality of flexible metal tongues having first ends interconnected by respective metal rings fixed securely to said inner and outer annular shells by first fixing means, and second ends fixed securely to the composite material combustion chamber at it air feed orifices by second fixing means.

Gas turbine combustor

Abstract: The present invention provides a gas turbine combustor that has a fuel injection nozzle using a fluid other than liquid fuel to atomize the liquid fuel and that can suppress a reduction in power generation efficiency resulting from a heat loss while promoting the atomization of the liquid fuel. The gas turbine combustor is adapted to mix liquid fuel with combustion air led from a compressor, burn the mixture, and supply combustion gas generated to a gas turbine. The gas turbine combustor includes a fuel injection nozzle that atomizes the liquid fuel into fine liquid droplets. The fuel injection nozzle includes a first system adapted to supply the liquid fuel and a second system adapted to supply a fluid for atomizing the liquid fuel. Low-boiling liquid fuel is supplied to the second system as the fluid. The second system is adapted to heat and supply the low-boiling liquid fuel.

Self-accommodating brush seal for gas turbine combustor

Abstract: The gas turbine includes a combustor 10 having a transition piece 18 with its exit end spaced from a first-stage nozzle 22. A brush seal 26 seals about the space between the transition piece and first-stage nozzle. The brush seal includes a sealing cap 28 supported by one of the transition piece and the first-stage nozzle and having a portion 38 overlying the other of the transition piece and first-stage nozzle. A seal ring 30 is secured to the other of the transition piece and the first-stage nozzle and carries bristles 32 projecting therefrom for engagement with the overlying portion 38 of the sealing cap 28 to seal the gap between the transition piece and the first-stage nozzle. In one form, the side face portions of the bristles of the bristle pack form the seal by engaging against a planar sealing surface.

Gas turbine combustor

Abstract: The present invention provides a gas turbine combustor which makes it possible to achieve both a high combustion efficiency and low NOx emissions characteristics over a wide output power range without using a device that can vary the flow rate of the air used for combustion, by burning a lean mixture using high-temperature burned gas. The gas turbine combustor 10 comprises pre-mixture injection tubes 16 which conduct a mixture of fuel and air into a combustion chamber 11 . The mixture from the pre-mixture injection tubes 16 is injected toward burned gas 19 a present on the downstream side of the flame 19 , which is injected from burners 15 that open into the combustion chamber 11 , and this mixture is mixed with the burned gas 19 a . Even in a mixture which is leaner than the lower limit of inflammability, the radicals in the burned gas 19 a are effective in initiating reactions, so that the combustion of the mixture can be started. Since the burned gas has a lower oxygen concentration than fresh mixture, and since the pre-mixture is dispersed in space, the generation of NOx can be suppressed.

Methods and apparatuses for vertebral body distraction and fusion employing a coaxial screw gear sleeve mechanism

Abstract: Improved methods and apparatuses for vertebral body distraction and fusion in accordance with various embodiments of the present invention employ one or more coaxial screw gear sleeve mechanisms. In various embodiments, coaxial screw gear sleeve mechanisms include a post with a threaded exterior surface and a corresponding sleeve configured to surround the post, the corresponding sleeve having a threaded interior surface configured to interface with the threaded exterior surface of the post and a geared exterior surface. A drive mechanism can be configured to interface with the geared exterior surface of the sleeve, causing the device to distract.

Sealing assembly for the aft end of a ceramic matrix composite liner in a gas turbine engine combustor

Abstract: An assembly for providing a seal at an aft end of a combustor liner for a gas turbine engine including a longitudinal centerline axis extending therethrough. The sealing assembly includes a substantially annular first sealing member positioned between an aft portion of a support member and the liner aft end so as to seat on a designated surface portion of the liner aft end and a substantially annular second sealing member positioned between the support member aft portion and a turbine nozzle located downstream of the liner aft end so as to seat on a designated surface portion of the support member aft portion. Accordingly, the first sealing member is maintained in its seated position as the support member aft portion moves radially with respect to the liner aft end and the second sealing member is maintained in its seated position as the support member aft portion moves axially with respect to the turbine nozzle. The first and second sealing members are also maintained in their respective seating positions as the support member aft portion moves axially with respect to the liner aft end and radially with respect to the turbine nozzle.