GAS TURBINE ENGINE MOUNT

A gas turbine engine includes an aft frame and a forward frame disposed upstream from the aft frame. The forward frame includes an outer ring. The outer ring includes an inner surface that is radially spaced from an outer surface. A first indention is defined in the outer surface, and a first engine mount flange protrudes radially outwardly from the first indention.

FIELD

The present disclosure relates to gas turbine engine mounts.

BACKGROUND

Aircraft may be powered by one or more gas turbine engines. The engine(s) may be mounted to the aircraft via one or more engine frames configured to interlock or couple to a pylon or other mounting feature of the aircraft structure.

DETAILED DESCRIPTION

The terms “forward” and “aft” refer to relative positions within a gas turbine engine or aircraft and refer to the normal operational attitude of the gas turbine engine or aircraft. For example, with regards to a gas turbine engine, forward refers to a position closer to an engine inlet and aft refers to a position closer to an engine nozzle or exhaust.

A “third stream” as used herein means a non-primary air stream capable of increasing fluid energy to produce a minority of total propulsion system thrust. A pressure ratio of the third stream may be higher than that of the primary propulsion stream (e.g., a bypass or propeller driven propulsion stream). The thrust may be produced through a dedicated nozzle or through mixing of an airflow through the third stream with a primary propulsion stream or a core air stream, e.g., into a common nozzle.

Generally, exhaust flowing from an open fan prop or unducted primary fan of a turbofan engine is sonic or even supersonic at climb and cruise conditions. This high-speed exhaust flow scrubs across or by the nacelle and pylon fairing. For sonic and supersonic flow around the nacelle and pylon fairing, shocks and shock losses are expected and can result in a considerable penalty to aircraft-level performance such as fuel burn rate and drag. Engine mounts are a key constraint on the nacelle flow path and depressing the engine mounts into an engine's frame, particularly the forward frame, enables a smoother, lower profile, more continuous curvature flow path around the engine nacelle and any local fairings around the engine mounts. The design disclosed herein reduces total drag, peak Mach numbers occurring on the nacelle and/or pylon, and therefore the risk of wave drag.

Referring now to the drawings,FIG.1is a perspective view of an exemplary aircraft10that may incorporate at least one exemplary embodiment of the present disclosure. As shown inFIG.1, the aircraft10has a fuselage12, wings14attached to the fuselage12, and an empennage16. The aircraft10further includes a propulsion system18that produces a propulsive thrust to propel the aircraft10in flight, during taxiing operations, etc. Although the propulsion system18is shown attached to the wing(s)14, in other embodiments it may additionally or alternatively include one or more aspects coupled to other parts of the aircraft10, such as, for example, the empennage16, the fuselage12, etc. The propulsion system18includes at least one engine. In the exemplary embodiment shown, the aircraft10includes a pair of gas turbine engines20. In particular embodiments, each gas turbine engine20is mounted to the aircraft10in an under-wing configuration via a respective pylon22. Each gas turbine engine20is capable of selectively generating a propulsive thrust for the aircraft10. It is to be appreciated that the gas turbine engines20may also be mounted in other locations of the aircraft such as but not limited to the empennage16. In addition, the gas turbine engines20may be configured to burn various forms of fuel including, but not limited to unless otherwise provided, jet fuel/aviation turbine fuel, and hydrogen fuel.

FIG.2is a schematic cross-sectional view of a gas turbine engine100according to another example embodiment of the present disclosure. Particularly,FIG.2provides a turbofan engine having a rotor assembly with a single stage of unducted rotor blades. In such a manner, the rotor assembly may be referred to herein as an “unducted fan,” or the entire engine100may be referred to as an “unducted turbofan engine.” In addition, the engine100ofFIG.2includes a third stream extending from the compressor section to a rotor assembly flowpath over the turbomachine, as will be explained in more detail below.

For reference, the engine100defines an axial direction A, a radial direction R, and a circumferential direction C. Moreover, the engine100defines an axial centerline or longitudinal axis102that extends along the axial direction A. In general, the axial direction A extends parallel to the longitudinal axis102, the radial direction R extends outward from and inward to the longitudinal axis102in a direction orthogonal to the axial direction A, and the circumferential direction extends three hundred sixty degrees (360°) around the longitudinal axis102. The engine100extends between a forward end104and an aft end106, e.g., along the axial direction A.

As shown inFIG.2the engine100includes a turbomachine108having a fan section136that is positioned upstream thereof. Generally, the turbomachine108includes, in serial flow order, a compressor section, a combustion section, a turbine section, and an exhaust section. Particularly, as shown inFIG.2, the turbomachine108includes a core cowl110that defines an annular core inlet112. The core cowl110further encloses at least in part a low-pressure system and a high-pressure system. For example, the core cowl110depicted encloses and supports at least in part a booster or low-pressure compressor114for pressurizing the air that enters the turbomachine108through core inlet112. A high-pressure, multi-stage, axial-flow compressor116receives pressurized air from the low-pressure compressor114and further increases the pressure of the air. The pressurized air stream flows downstream to a combustor118of the combustion section where fuel is injected into the pressurized air stream and ignited to raise the temperature and energy level of the pressurized air.

It will be appreciated that as used herein, the terms “high/low speed” and “high/low-pressure” are used with respect to the high-pressure/high speed system and low-pressure/low speed system interchangeably. Further, it will be appreciated that the terms “high” and “low” are used in this same context to distinguish the two systems and are not meant to imply any absolute speed and/or pressure values.

The high energy combustion products flow from the combustor118downstream to a high-pressure turbine120. The high-pressure turbine120drives the high-pressure compressor116through a high-pressure shaft124. In this regard, the high-pressure turbine120is drivingly coupled with the high-pressure compressor116. The high energy combustion products then flow to a low-pressure turbine122. The low-pressure turbine122drives the low-pressure compressor114and components of the fan section136through a low-pressure shaft126. In this regard, the low-pressure turbine122is drivingly coupled with the low-pressure compressor114and components of the fan section136. The low-pressure shaft126is coaxial with the high-pressure shaft124in this example embodiment. After driving each of the high-pressure turbine120and the low-pressure turbine122, the combustion products exit the turbomachine108through a turbomachine exhaust nozzle128. A core engine134of the gas turbine engine100is defined as the part of the gas turbine engine100that extends from the fan blades140of the fan section136to the turbomachine exhaust nozzle128.

Accordingly, the turbomachine108defines a working gas flowpath or core duct130that extends between the core inlet112and the turbomachine exhaust nozzle128. The core duct130is an annular duct positioned generally inward of the core cowl110along the radial direction R. The core duct130(e.g., the working gas flowpath through the turbomachine108) may be referred to as a second stream. The fan section136includes a fan138, which is the primary fan in this example embodiment. For the depicted embodiment ofFIG.2, the fan138is an open rotor or unducted fan138. In such a manner, the engine100may be referred to as an open rotor engine.

As depicted, the fan138includes a plurality or an array of fan blades140(only one shown inFIG.2). The fan blades140are rotatable, e.g., about the longitudinal axis102. As noted above, the fan138is drivingly coupled with the low-pressure turbine122via the low-pressure shaft126. For the embodiments shown inFIG.2, the fan138is coupled with the low-pressure shaft126via a speed reduction gearbox142, e.g., in an indirect-drive or geared-drive configuration.

Moreover, the array of fan blades140can be arranged in equal spacing around the longitudinal axis102. Each fan blade140has a root and a tip and a span defined therebetween. Each fan blade140defines a central blade axis144. For this embodiment, each fan blade140of the fan138is rotatable about its central blade axis144, e.g., in unison with one another. One or more actuators146are provided to facilitate such rotation and therefore may be used to change a pitch of the fan blades140about their respective central blades' axes144.

The fan section136further includes a fan guide vane array148that includes fan guide vanes150(only one shown inFIG.2) disposed around the longitudinal axis102. For this embodiment, the fan guide vanes150are not rotatable about the longitudinal axis102. Each fan guide vane150has a root and a tip and a span defined therebetween. The fan guide vanes150may be unshrouded as shown inFIG.2or, alternatively, may be shrouded, e.g., by an annular shroud spaced outward from the tips of the fan guide vanes150along the radial direction R or attached to the fan guide vanes150.

Each fan guide vane150defines a central blade axis152. For this embodiment, each fan guide vane150of the fan guide vane array148is rotatable about its respective central blade axis152, e.g., in unison with one another. One or more actuators154are provided to facilitate such rotation and therefore may be used to change a pitch of the fan guide vane150about its respective central blade axis152. However, in other embodiments, each fan guide vane150may be fixed or unable to be pitched about its central blade axis152. The fan guide vanes150are mounted to a fan cowl156.

As shown inFIG.2, in addition to the fan138, which is unducted, a ducted fan170is included aft of the fan138, such that the engine100includes both a ducted and an unducted fan which both serve to generate thrust through the movement of air without passage through at least a portion of the turbomachine108(e.g., without passage through the high-pressure compressor116and combustion section for the embodiment depicted). The ducted fan170is rotatable about the same axis (e.g., the longitudinal axis102) as the fan blade140. The ducted fan170is, for the embodiment depicted, driven by the low-pressure turbine122(e.g., coupled to the low-pressure shaft126). In the embodiment depicted, as noted above, the fan138may be referred to as the primary fan, and the ducted fan170may be referred to as a secondary fan. It will be appreciated that these terms “primary” and “secondary” are terms of convenience, and do not imply any particular importance, power, or the like.

The ducted fan170includes a plurality of fan blades (not separately labeled inFIG.2) arranged in a single stage, such that the ducted fan170may be referred to as a single stage fan. The fan blades of the ducted fan170can be arranged in equal circumferential spacing around the longitudinal axis102. Each blade of the ducted fan170has a root and a tip and a span defined therebetween.

The fan cowl156annularly encases at least a portion of the core cowl110and is generally positioned outward of at least a portion of the core cowl110along the radial direction R. Particularly, a downstream section of the fan cowl156extends over a forward portion of the core cowl110to define a fan duct flowpath, or simply a fan duct158. According to this embodiment, the fan flowpath or fan duct158may be understood as forming at least a portion of the third stream of the engine100.

Incoming air may enter through the fan duct158through a fan duct inlet162and may exit through a fan exhaust nozzle164to produce propulsive thrust. The fan duct158is an annular duct positioned generally outward of the core duct130along the radial direction R. The fan cowl156and the core cowl110are connected together and supported by a plurality of substantially radially extending and circumferentially spaced stationary struts160(only one shown inFIG.2).

The stationary struts160may each be aerodynamically contoured to direct air flowing thereby. Other struts in addition to the stationary struts160may be used to connect and support the fan cowl156and/or core cowl110. In many embodiments, the fan duct158and the core duct130may at least partially co-extend (generally axially) on opposite sides (e.g., opposite radial sides) of the core cowl110. For example, the fan duct158and the core duct130may each extend directly from a leading edge132of the core cowl110and may partially co-extend generally axially on opposite radial sides of the core cowl110.

The exemplary engine100shown inFIG.2also defines or includes an inlet duct166. The inlet duct166extends between an engine inlet168and the core inlet112and fan duct inlet162. The engine inlet168is defined generally at the forward end of the fan cowl156and is positioned between the fan138and the fan guide vane array148along the axial direction A. The inlet duct166is an annular duct that is positioned inward of the fan cowl156along the radial direction R. Air flowing downstream along the inlet duct166is split, not necessarily evenly, into the core duct130and the fan duct158by a fan duct splitter or the leading edge132of the core cowl110. In the embodiment depicted, the inlet duct166is wider than the core duct130along the radial direction R. The inlet duct166is also wider than the fan duct158along the radial direction R.

Notably, for the embodiment depicted, the engine100includes one or more features to increase an efficiency of a third-stream thrust, Fn3S (e.g., a thrust generated by an airflow through the fan duct158exiting through the fan exhaust nozzle164, generated at least in part by the ducted fan170). In particular, the engine100further includes an array of inlet guide vanes172positioned in the inlet duct166upstream of the ducted fan170and downstream of the engine inlet168. The array of inlet guide vanes172are arranged around the longitudinal axis102. For this embodiment, the inlet guide vanes172are not rotatable about the longitudinal axis102.

Each inlet guide vane172defines a central blade axis (not labeled for clarity), and is rotatable about its respective central blade axis, e.g., in unison with one another. In such a manner, the inlet guide vanes172may be considered a variable geometry component. One or more actuators174are provided to facilitate such rotation and therefore may be used to change a pitch of the inlet guide vanes172about their respective central blade axes. However, in other embodiments, each inlet guide vanes172may be fixed or unable to be pitched about its central blade axis.

Further, located downstream of the ducted fan170and upstream of the fan duct inlet162, the engine100includes an array of outlet guide vanes176. As with the array of inlet guide vanes172, the array of outlet guide vanes176are not rotatable about the longitudinal axis102. However, for the embodiment depicted, unlike the array of inlet guide vanes172, the array of outlet guide vanes176are configured as fixed-pitch outlet guide vanes.

Further, it will be appreciated that for the embodiment depicted, the fan exhaust nozzle164of the fan duct158is further configured as a variable geometry exhaust nozzle. In such a manner, the engine100includes one or more actuators178for modulating the variable geometry exhaust nozzle. For example, the variable geometry exhaust nozzle may be configured to vary a total cross-sectional area (e.g., an area of the nozzle in a plane perpendicular to the longitudinal axis102) to modulate an amount of thrust generated based on one or more engine operating conditions (e.g., temperature, pressure, mass flowrate, etc. of an airflow through the fan duct158). A fixed geometry exhaust nozzle may also be adopted.

The combination of the array of inlet guide vanes172located upstream of the ducted fan170, the array of outlet guide vanes176located downstream of the ducted fan170, and the fan exhaust nozzle164may result in a more efficient generation of third-stream thrust, Fn3S, during one or more engine operating conditions. Further, by introducing a variability in the geometry of the inlet guide vanes172and the fan exhaust nozzle164, the engine100may be capable of generating more efficient third-stream thrust, Fn3S, across a relatively wide array of engine operating conditions, including takeoff and climb (where a maximum total engine thrust FnTotal, is generally needed) as well as cruise (where a lesser amount of total engine thrust, FnTotal, is generally needed).

Moreover, referring still toFIG.2, in exemplary embodiments, air passing through the fan duct158may be relatively cooler (e.g., lower temperature) than one or more fluids utilized in the turbomachine108. In this way, one or more heat exchangers180may be positioned in thermal communication with the fan duct158. For example, one or more heat exchangers180may be disposed within the fan duct158and utilized to cool one or more fluids from the core engine134with the air passing through the fan duct158, as a resource for removing heat from a fluid, e.g., compressor bleed air, oil or fuel.

Although not depicted in detail, the heat exchanger180may be an annular heat exchanger extending substantially 360 degrees in the fan duct158(e.g., at least 300 degrees, such as at least 330 degrees). In such a manner, the heat exchanger180may effectively utilize the air passing through the fan duct158to cool one or more systems of the engine100(e.g., lubrication oil systems, compressor bleed air, electrical components, etc.). The heat exchanger180uses the air passing through duct158as a heat sink and correspondingly increases the temperature of the air downstream of the heat exchanger180and exiting the fan exhaust nozzle164.

Although depicted above as an unshrouded or open rotor engine in the embodiments depicted above, it should be appreciated that aspects of the disclosure provided herein may be applied to shrouded or ducted engines, partially ducted engines, aft-fan engines, or other gas turbine engine configurations, including those for marine, industrial, or aero-propulsion systems. Certain aspects of the disclosure may be applicable to turbofan, turboprop, or turboshaft engines.

FIG.3is a schematic cross-sectional view of the gas turbine engine100as shown inFIG.2according to exemplary embodiments of the present disclosure. As shown inFIG.3, the engine100further includes an aft frame182and a forward frame184which is disposed upstream of the aft frame182with respect to fluid flow through the engine100. The aft frame182and the forward frame184at least partially define the fluid flow path through the gas turbine engine100. In addition, the aft frame182and the forward frame184are used to couple or mount the gas turbine engine100to the pylon22via respective engine mounts186and188. The term “forward frame” as used herein may include any frame that is forward of the aft frame. For example, in particular embodiments, the forward frame may be an inlet frame, a mid-frame or a turbine vane frame.

In certain configurations, the engine100may further include a thrust link or linkage190that transmits axial, engine thrust loads between the engine100and the pylon22and/or the airframe of the aircraft10shown inFIG.1.

FIG.4provides a front perspective view of an exemplary forward frame200according to various embodiments of the present disclosure.FIG.5provides a back perspective view of the exemplary forward frame200as shown inFIG.4, according to various embodiments of the present disclosure.FIG.6is a top partial view of the exemplary forward frame as shown inFIGS.4and5, according to various embodiments of the present disclosure.

In exemplary embodiments, as shown inFIGS.4,5and6collectively, the forward frame200includes an outer ring202. The outer ring202includes an inner surface204that is radially spaced from an outer surface206with respect to radial direction R. As shown inFIGS.5and6, the outer ring202further includes a first depression or indention208defined in the outer surface206. A first engine mount flange210protrudes radially outwardly with respect to radial direction R from the first indention208.

In exemplary embodiments, as shown inFIGS.4and5, the forward frame200includes an inner ring212coaxially aligned with the outer ring202with respect to axial centerline or longitudinal axis102of the gas turbine engine100(FIG.2) and/or axial centerline214of the forward frame200. The inner ring212includes an outer surface216that is circumferentially surrounded by the outer ring202. A flow channel218is defined between the outer surface216of the inner ring212and the inner surface204of the outer ring202.

As shown inFIGS.4and5collectively, the forward frame200includes an upstream end220axially spaced with respect to axial direction A from a downstream end222. In various embodiments, the flow channel218converges along longitudinal axis102or axial centerline214between the upstream end220and the downstream end222. In particular embodiments, the forward frame200further comprises a plurality of struts.224that extend radially from the outer surface216of the inner ring212to the inner surface204of the outer ring202within the flow channel218.

In exemplary embodiments, as shown inFIG.4, a portion or portions of the inner surface204of the outer ring202at the first indention208as indicated with arrows204(a) and204(b) protrudes or extend(s) radially inward with respect to radial direction R and into the flow channel218. In particular embodiments, as shown inFIG.4, the inner surface204of the outer ring202at the first indention208extends radially inward into the flow channel218at circumferentially opposing sides226,228of a first strut224(a) of the plurality of struts224.

In exemplary embodiments, as shown inFIGS.4,5and6collectively, the first engine mount flange210is configured to mount to the pylon22(FIGS.1and3). For example, the first engine mount flange210may be connected to a mounting device230such as but not limited to a coupler, pin, tab, block or the like which is shaped or formed to interlock with a complementary mounting feature (not shown) of the pylon22.

FIG.7provides a back or rear side perspective view of the exemplary forward frame200according to particular embodiments of the present disclosure.FIG.8provides a partial front view of the exemplary forward frame200as shown inFIG.7.

In particular embodiments, as shown inFIGS.7and8collectively, the outer ring202further includes a second indention232defined along the outer surface206of the outer ring202and a second engine mount flange234(FIG.7) protruding radially outwardly from the second indention232with respect to radial direction R. The second indention232and the second engine mount flange234are circumferentially spaced with respect to circumferential direction C from the first indention208and the first engine mount flange210. In particular embodiments, as shown inFIG.8, the inner surface204of the outer ring202at the first indention208and at the second indention232extends radially inward with respect to radial direction R into the flow channel218.

In particular embodiments, as shown inFIG.8, a portion or portions of the inner surface204of the outer ring202at the second indention232as indicated with arrows204(c) and204(d) protrudes or extend(s) radially inward with respect to radial direction R and into the flow channel218. In particular embodiments, as shown inFIG.8, a portion or portions of the inner surface204of the outer ring202as indicated by204(c) and204(d) at the second indention232extends radially inward into the flow channel218with respect to the radial direction R on circumferentially opposing sides236,238of a second strut224(b) of the plurality of struts224(FIG.5) with respect to circumferential direction C. The degree of the indention(s)208,232into the forward frame200, particularly into the flow channel218can be parameterized in different ways. For example, the degree or intrusion of the indention into the flow channel218may be parameterized as a % of flow blockage through the flow channel218at maximum depression bowl size of 0% to 50%. In addition or in the alternative, the degree or intrusion of the indention into the flow channel218may be parameterized as a radial extent of blockage, vs. flow-channel span, as viewed cross-sectionally from 0% to 50%.

It is to be appreciated that although it is not illustrated in the figures, the forward frame200may include more than two indentions each with respective engine mount flanges depending on the mounting configuration/requirements of a particular engine design. Although not shown, it is also to be appreciated that the gas turbine engine100may include two or more forward frames which includes one or more indentions as described above and as shown inFIGS.4-8. It is also to be appreciated that the number of indentions of each forward frame can be different where multiple forward frames are used.

As previously mentioned, engine mounts are a key constraint on the nacelle flow path and depressing the engine mounts into an engine's frame in the manner described and claimed herein, particularly the forward frame, enables a smoother, lower profile, more continuous curvature flow path around the engine nacelle and any local fairings around the engine mounts. More particularly, the design disclosed herein reduces total drag, peak Mach numbers occurring on the nacelle and/or pylon, and therefore the risk of wave drag.

A gas turbine engine comprising: an aft frame, and a forward frame disposed upstream from the aft frame. The forward frame including an outer ring, wherein the outer ring includes an inner surface radially spaced from an outer surface, a first indention defined in the outer surface, and a first engine mount flange protruding radially outwardly from the first indention.

The gas turbine engine of the preceding clause, wherein the gas turbine engine has an unducted primary fan.

The gas turbine engine of any preceding clause, wherein gas turbine engine is a three-stream gas turbine engine.

The gas turbine engine of any preceding clause, wherein the forward frame further comprises an inner ring including an outer surface circumferentially surrounded by the outer ring, wherein a flow channel is defined between the outer surface of the inner ring and the inner surface of the outer ring.

The gas turbine engine of any preceding clause, wherein the forward frame includes an upstream end axially spaced from a downstream end, wherein the flow channel converges between the upstream end and the downstream end.

The gas turbine engine of any preceding clause, wherein the inner surface of the outer ring at the first indention extends radially inward into the flow channel.

The gas turbine engine of any preceding clause, wherein the forward frame further comprises a plurality of struts that extend radially from the outer surface of the inner ring to the inner surface of the outer ring within the flow channel.

The gas turbine engine of any preceding clause, wherein the inner surface of the outer ring at the first indention extends radially inward into the flow channel at opposing sides of a strut of the plurality of struts.

The gas turbine engine of any preceding clause, wherein the outer ring further includes a second indention defined along the outer surface of the outer ring and a second engine mount flange protruding radially outwardly from the second indention.

The gas turbine engine of any preceding clause, wherein the second indention and the second engine mount are circumferentially spaced from the first indention and the first engine mount flange.

The gas turbine engine of any preceding clause, wherein the forward frame further comprises an inner ring including an outer surface circumferentially surrounded by the outer ring, wherein a flow channel is defined between the outer surface of the inner ring and the inner surface of the outer ring, and wherein the inner surface of the outer ring at the first indention and at the second indention extends radially inward into the flow channel.

The gas turbine engine of any preceding clause, wherein the forward frame further comprises a plurality of struts that extend radially from the outer surface of the inner ring to the inner surface of the outer ring within the flow channel, wherein the inner surface of the outer ring at the first indention extends radially inward into the flow channel at opposing sides of a first strut of the plurality of struts, and the second indention extends radially inward into the flow channel at opposing sides of a second strut of the plurality of struts.

An aircraft, comprising: a wing including a mounting pylon; and a gas turbine engine. The gas turbine engine comprising: a forward frame including an outer ring, wherein the outer ring includes an inner surface that is radially spaced from an outer surface, a first indention defined in the outer surface, and a first engine mount flange protruding radially outwardly from the first indention, wherein the first engine mount flange is coupled to the pylon.

The aircraft of the preceding clause, wherein the gas turbine engine has an unducted primary fan.

The aircraft of any preceding clause, wherein gas turbine engine is a three-stream gas turbine engine.

The aircraft of any preceding clause, wherein the forward frame further comprises an inner ring including an outer surface circumferentially surrounded by the outer ring, wherein a flow channel is defined between the outer surface of the inner ring and the inner surface of the outer ring.

The aircraft of any preceding clause, wherein the forward frame includes an upstream end axially spaced from a downstream end, wherein the flow channel converges between the upstream end and the downstream end.

The aircraft of any preceding clause, wherein the inner surface of the outer ring at the first indention extends radially inward into the flow channel.

The aircraft of any preceding clause, wherein the forward frame further comprises a plurality of struts that extend radially from the outer surface of the inner ring to the inner surface of the outer ring within the flow channel, wherein the inner surface of the outer ring at the first indention extends radially inward into the flow channel at opposing sides of a strut of the plurality of struts.

The aircraft of any preceding clause, wherein the outer ring further includes a second indention defined along the outer surface of the outer ring and a second engine mount flange protruding radially outwardly from the second indention, wherein the second indention and the second engine mount flange are circumferentially spaced from the first indention and the first engine mount flange, wherein the inner surface of the outer ring at the first indention and at the second indention extends radially inward into the flow channel.