Patent Description:
Gas turbine engines include a compressor section to pressurize a supply of air, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases. The compressor section discharges air into a pre-diffuser upstream of the combustion section. The pre-diffuser converts a portion of dynamic pressure to static pressure. A diffuser receives the air from the pre-diffuser and supplies the compressed core flow around an aerodynamically-shaped cowl of the combustion chamber. The core flow is typically separating into three branches. One branch is the cowl passage to supply air to fuel nozzles and for dome cooling. The other branches are annular outer plenum and inner plenums where air is introduced into the combustor for cooling and to complete the combustion process. A further portion of the air may be utilized for turbine cooling.

The pre-diffuser is exposed to large thermal gradients and requires various features for anti-rotation, axial retention, and centrality with respect to the central engine axis. These features may result in local discontinuities which may generate stress risers and consequently reduced operational life.

<CIT> discloses a prior art pre-diffuser according to the preamble of claim <NUM>.

<CIT> discloses another prior art pre-diffuser.

In a first aspect, a pre-diffuser is provided as set forth in claim <NUM>.

A further aspect of the present disclosure includes that each of the multiple of hollow struts include a cavity.

A further aspect of the present disclosure includes a passage in communication with the cavity.

A further aspect of the present disclosure includes that the passage is in communication with a combustor section of the gas turbine engine.

A further aspect of the present disclosure includes that the inlet to each of the multiple of diffusion passages are smaller than an exit from the diffusion passage.

A further aspect of the present disclosure includes that each of the multiple of hollow struts align with one of a respective multiple of exit guide vanes of an exit guide vane ring.

A further aspect of the present disclosure includes that the ring-strut-ring structure is cast.

A further aspect of the present disclosure includes that the static structure is subjected to a lower temperature than the ring-strut-ring structure.

These features and elements as well as the operation of the invention will become more apparent in light of the following description and the accompanying drawings.

The fan section <NUM> drives air along a bypass flowpath while the compressor section <NUM> drives air along a core flowpath for compression and communication into the combustor section <NUM>, then expansion through the turbine section <NUM>. Although depicted as a turbofan gas turbine engine in the disclosed non-limiting embodiment, it should be understood that the concepts described herein are not limited to use with turbofans as the teachings may be applied to other types of turbine engines.

The engine <NUM> generally includes a low spool <NUM> and a high spool <NUM> mounted for rotation about an engine central longitudinal axis A relative to an engine case structure <NUM> via several bearing structures <NUM>. The low spool <NUM> generally includes an inner shaft <NUM> that interconnects a fan <NUM>, a low pressure compressor (LPC) <NUM> and a low pressure turbine (LPT) <NUM>. The inner shaft <NUM> drives the fan <NUM> directly or through a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low spool <NUM>. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.

The high spool <NUM> includes an outer shaft <NUM> that interconnects a high pressure compressor (HPC) <NUM> and high pressure turbine (HPT) <NUM>. A combustor <NUM> is arranged between the HPC <NUM> and the HPT <NUM>. The inner shaft <NUM> and the outer shaft <NUM> are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes. Core airflow is compressed by the low pressure compressor <NUM>, then the high pressure compressor <NUM>, mixed with the fuel and burned in the combustor <NUM>, then expanded over the HPT <NUM> and LPT <NUM>. The HPT <NUM> and LPT <NUM> rotationally drive the respective high spool <NUM> and low spool <NUM> in response to the expansion.

With reference to <FIG>, the combustor <NUM> generally includes an outer liner <NUM>, an inner liner <NUM> and a diffuser case module <NUM>. The outer liner <NUM> and the inner liner <NUM> are spaced apart such that a combustion chamber <NUM> is defined therebetween. The combustion chamber <NUM> is generally annular in shape. The outer liner <NUM> and the inner liner <NUM> are spaced radially inward of the outer diffuser case <NUM> to define an annular outer plenum <NUM> and an annular inner plenum <NUM>. It should be understood that although a particular combustor is illustrated, other combustor types with various combustor liner arrangements will also benefit herefrom. It should be further understood that the disclosed cooling flow paths are but an illustrated embodiment and should not be limited only thereto.

The liners <NUM>, <NUM> contain the combustion products for direction toward the turbine section <NUM>. Each liner <NUM>, <NUM> generally includes a respective support shell <NUM>, <NUM> which supports one or more heat shields <NUM>, <NUM> that are attached thereto with fasteners <NUM>.

The combustor <NUM> also includes a forward assembly <NUM> downstream of the compressor section <NUM> to receive compressed airflow through a pre-diffuser <NUM> into the combustor section <NUM>. The pre-diffuser <NUM> includes a hot fairing structure <NUM> and an exit guide vane ring <NUM>. The exit guide vane ring <NUM> includes a row of Exit Guide Vanes (EGVs) <NUM> downstream of the HPC <NUM>. The EGVs <NUM> are static engine components which direct core airflow from the HPC <NUM> between outboard and inboard walls <NUM> and <NUM>.

The pre-diffuser <NUM> is secured to a static structure <NUM> to at least partially form the diffuser module between the compressor section <NUM> and the combustor section <NUM>. The hot fairing structure <NUM> is exposed to large thermal gradients and directs the core airflow while forming a shell within the relatively colder static structure <NUM>. The static structure <NUM> is thereby segregated from the core airflow and generally operates at a relatively lower temperature than the hot fairing structure <NUM>. The hot fairing structure <NUM> and the exit guide vane ring <NUM> are full ring structures that are assembled in a manner that allows common thermal growth yet still remain centered with respect to the static structure <NUM> along the engine central longitudinal axis A.

With reference to <FIG>, the hot fairing structure <NUM> includes a ring-strut-ring structure <NUM> which forms a multiple of diffusion passages <NUM> that each communicate with one of a multiple of diffusion passage ducts <NUM> (<FIG>) that extend the diffusion passage of the ring-strut-ring structure <NUM> along each flow passage P. Each of the diffusion passages <NUM> in the ring-strut-ring structure <NUM> includes an inlet to the pre-diffuser <NUM> and a diffusion passage exit that mates with the diffusion passage duct <NUM>. Each of the diffusion passage ducts <NUM> include a diffusion duct inlet <NUM> (<FIG>) adjacent to the ring-strut-ring structure <NUM>. A diffusion duct exit <NUM> from each diffusion passage duct <NUM> provide the outlet from the pre-diffuser <NUM>. The diffusion duct exits <NUM> (<FIG>) are larger than the respective diffusion duct inlets <NUM> which are positioned between each of the EGVs <NUM>. In one example, the number of EGVs are <NUM>-<NUM> times more than the number of diffusion duct inlets <NUM>. In this embodiment, the diffusion passage ducts <NUM> expand primarily in the radial direction to the diffusion duct exits <NUM>.

The hot fairing structure <NUM> and the exit guide vane ring <NUM> include an anti-rotation interface <NUM> that positions the anti-rotation features <NUM>, <NUM> in a region of low stress inboard of the diffusion passages <NUM>. The hot fairing structure <NUM> includes a multiple of circumferentially located anti-rotation tabs <NUM> (<FIG>) that engage respective anti-rotation slots <NUM> (<FIG>) in the exit guide vane ring <NUM>. The inboard location of the anti-rotation features <NUM>, <NUM> allow the multiple, static, hot components to grow and interact together, with low stress, and simultaneously remain aligned with the rotating components to facilitate a longer service life and engine efficiency.

An axial extension <NUM> of the hot fairing structure <NUM> extends along an inner diameter flow surface of the flow passage P. The axial extension <NUM> at least partially overlaps a recessed area <NUM> of the exit guide vane ring <NUM>. That is, the axial extension <NUM> extends in a direction opposite that of the core flow in the flow passage P and overlaps the recessed area <NUM> (<FIG>) in the exit guide vane ring <NUM>.

A hot fairing radial flange <NUM> extends from the hot fairing structure <NUM> parallel to an exit guide vane radial flange <NUM> of the exit guide vane ring <NUM>. A static structure flange <NUM> extends radially outwardly from the static structure <NUM> with respect to the engine axis A to abut the hot fairing radial flange <NUM>. That is, the static structure flange <NUM> operates as a mount location for the hot fairing structure <NUM> and the exit guide vane ring <NUM>. The hot fairing radial flange <NUM> also includes a multiple of circumferentially located anti-rotation tabs <NUM> (<FIG>) opposite the anti-rotation tabs <NUM> that engage respective anti-rotation slots <NUM> (<FIG>) in the static structure flange <NUM> of the static structure <NUM>.

A clamp ring <NUM> abuts the exit guide vane radial flange <NUM> to sandwich a seal member <NUM> between the exit guide vane radial flange <NUM> and the hot fairing radial flange <NUM>. A seal member <NUM>, e.g., a torsional spring seal, dogbone, or diamond seal, that accommodates compression of the hot fairing structure <NUM> and the exit guide vane ring <NUM> in response to axial assembly of the static structure modules. A multiple of circumferentially arranged fasteners <NUM> fastens the clamp ring <NUM> to the static structure <NUM>.

An outer radial interface <NUM> between the hot fairing structure <NUM> and the exit guide vane ring <NUM> includes a radial interface <NUM> and an axial interface <NUM>. Since the outer radial interface <NUM> of the hot fairing structure <NUM> and the exit guide vane ring <NUM> are devoid of discontinuities and are uniform in cross-section around the circumference of the full hoop structures, service life is significantly increased. The anti-rotation interface <NUM> and the outer radial interface <NUM> are essentially hidden from the gas path and are located in low stress regions.

With reference to <FIG>, the ring-strut-ring structure <NUM> may be cast from nickel alloys to provide for structural attachment and efficient sealing between turbine engine components combined with independently manufactured thin-wall diffusion passage ducts <NUM>. The diffusion passage ducts <NUM> can be manufactured by several methods including cast, sheet-metal formed, additively manufactured, or combinations thereof. The wall thickness and local stiffness of the diffusion passage ducts <NUM> can be tailored to a specific requirement thereof without excessive weight as is typical of cast components. The joining of the diffusion passage ducts <NUM> to the ring-strut-ring structure <NUM> to form each complete diffusion passage may be by brazing, bonding, welding, mechanical, or others. Light weight diffusion passage ducts <NUM> reduce the overall weight of the design, simplify the ring-strut-ring structure <NUM> casting process, and increase the natural frequencies of the hot fairing structure <NUM> by minimizing the cantilevered mass of the diffusion passage ducts <NUM>.

With reference to <FIG>, the one-piece ring-strut-ring structure <NUM> of the hot fairing structure <NUM> includes a multiple of hollow struts <NUM> that align with the respective multiple of upstream EGVs <NUM> of the exit guide vane ring <NUM> and split the flow into two adjacent diffusion passage ducts <NUM> (<FIG>). Each of the multiple of hollow struts <NUM> are generally airfoil shaped. In this embodiment, the hollow struts <NUM> reduce thermal mass and thickness so that the transient thermal gradient within the strut is minimal. The hollow strut <NUM> includes a cavity <NUM> that may be manufactured with ceramic cores, and a core exit via a passage <NUM> may be located at a location that has the least impact on thermal stiffness. Alternatively, the struts <NUM> may be solid (<FIG>).

Each passage <NUM> is located along an axis D and is in communication with the cavity <NUM> in the hollow strut <NUM>. The passage <NUM> may be reinforced and permits diffusion air from the diffuser side of the pre-diffuser <NUM>, i.e., the air around the combustor <NUM>, to be received into the respective cavity <NUM>. The diffuser air facilitates thermal control of the ring-strut-ring structure <NUM> of the hot fairing structure <NUM> to reduce the mass of the ring-strut-ring structure <NUM>. The reduced mass of the ring-strut-ring structure <NUM> of the hot fairing structure <NUM> results in a more responsive thermal characteristic. The strut geometry maximizes the perimeter of the ring-strut-ring structure <NUM> that is engaged in torsional stiffness. That is, the mass close to the centroid <NUM> has little to no effect on stiffness. To resist multi-node sinusoidal waves travelling around the circumference of the hot fairing structure <NUM>, local torsional sectional properties of the ring-strut-ring structure <NUM> facilitate control of the natural frequencies of the hot fairing structure <NUM>.

The ring-strut-ring structure <NUM> with the hollow regions with the core breakout located close to the centroid <NUM> of the torsional section forms a pre-diffuser <NUM> that can have both high natural frequencies and more uniform transient thermal gradients which enables a lightweight, high performance low thermal stress design. The hot fairing structure <NUM> with a hollow leading edge region and the core opening on the aft side of the hollow strut <NUM>, is located about the mid-axis of the airfoil shape to connect outer diameter static structure, with minimal thermal mass, and an inner diameter static structure with distributed mass such that the transient thermal response is optimized to reduce thermal stress.

The ring-strut-ring structure <NUM> also allows coupled Exit Guide Vanes with the floating hot fairing to provide improved cyclic life. Light weight tubular flowpath extensions reduce the overall weight of the design, simplify the ring-strut-ring structure <NUM> casting process, and increase the natural frequencies of the hot fairing by minimizing the cantilevered mass of the tubes. Additionally, the torsionally stiff ring-strut-ring structure <NUM> ensures that the design can be incorporated with features on the inner diameter structure which facilitates attachment to other structures with the least amount of contact, yet have sufficient frequency margin with respect to engine operating vibration sources.

It should also be understood that although a particular component arrangement is disclosed in the illustrated embodiment, other arrangements will benefit herefrom.

Claim 1:
A pre-diffuser (<NUM>) for a gas turbine engine
comprising a hot fairing structure (<NUM>) and an exit guide vane ring (<NUM>), the hot fairing structure (<NUM>) comprising:
a ring-strut-ring structure (<NUM>) that comprises a multiple of hollow struts (<NUM>) and a multiple of inlets to a respective diffusion passage (<NUM>), one of the multiple of inlets formed between each one of the multiple of hollow struts (<NUM>) located between two diffusion passages (<NUM>), wherein each diffusion passage (<NUM>) communicates with one of a multiple of diffusion passage ducts (<NUM>), and wherein the hot fairing structure (<NUM>) is a full ring structure; and
a full ring hot fairing radial flange (<NUM>) that extends transverse to the multiple of diffusion passages (<NUM>), wherein the full ring hot fairing radial flange (<NUM>) extends from the hot fairing structure (<NUM>) parallel to an exit guide vane radial flange (<NUM>) of the exit guide vane ring (<NUM>);
characterised in that the hot fairing structure (<NUM>) further comprises:
a first anti-rotation feature (<NUM>) on one side of the full ring hot fairing radial flange (<NUM>) and a second anti-rotation feature (<NUM>) on an opposite side of the full ring hot fairing radial flange (<NUM>), wherein the first anti-rotation feature (<NUM>) is configured to engage the exit guide vane ring (<NUM>) and the second anti-rotation feature (<NUM>) is configured to engage a static structure (<NUM>).