METHOD AND SYSTEM FOR MOUNTING AN AIRCRAFT ENGINE

A system for mounting an engine to an aircraft includes a rigid structure coupled to a wing and including a forward mount interface and an aft mount interface. The system includes a frame including a first support connection and a second support connection spaced apart from the first support connection. A linkage structure couples the frame to the rigid structure and includes a first linkage pair extending between the forward mount interface and the first support connection at a first angle with respect to a rotational axis, and a second linkage pair extending between the aft mount interface and the second support connection at a second angle with respect to the rotational axis.

BACKGROUND

The present invention generally relates to systems and methods for mounting an aircraft engine to an aircraft. More particularly, this invention relates to a mounting system and method adapted to reduce backbone deflection that can occur in an aircraft engine as a result of aerodynamic, gravitational, inertial, and thrust loads during aircraft operation.

At least some known gas turbine engines, such as turbofans, include a fan, a core engine, and a power turbine. The core engine includes at least one compressor, a combustor, and a high-pressure turbine coupled together in a serial flow relationship. More specifically, the compressor and high-pressure turbine are coupled through a shaft to form a high-pressure rotor assembly. Air entering the core engine is mixed with fuel and ignited to form a high energy gas stream. The high energy gas stream flows through the high-pressure turbine to rotatably drive the high-pressure turbine such that the shaft rotatably drives the compressor. The gas stream expands as it flows through a power or low-pressure turbine positioned aft of the high-pressure turbine. The low-pressure turbine includes a rotor assembly having a fan coupled to a drive shaft. The low-pressure turbine rotatably drives the fan through the drive shaft. Turbine engine performance is enhanced when the low-pressure turbine operates at a relatively high rotational speed and when the fan operates at a relatively low rotational speed and with a low pressure ratio.

As engine bypass ratios are increased, the larger fan and increased airflow result in higher loads at take-off rotation. A large lift load is created on the engine inlet as internal and some external airflow is turned to align with the engine axis. This load represents a major contribution to the backbone bending moment. The engine thrust also creates a pitching moment depending on whether the focal point of the engine's mounting system is on, above or below the engine center-line. The smaller core diameters associated with increased bypass ratio engines, together with increased pressure ratios and smaller blade heights, make the core engine more sensitive to backbone bending. At least some engines include more open tip clearances, to accommodate backbone bending. However, such open tip clearances may result in a reduction in fuel economy.

BRIEF DESCRIPTION

In one aspect, a system for mounting an engine to an engine support structure of an aircraft is provided. The system includes a rigid structure coupled to a wing of the aircraft and including a forward mount interface and an aft mount interface. Each of the forward mount interface and the aft mount interface receives a thrust component of load. The system also includes a frame surrounding a rotational axis of the engine. The frame includes a first support connection and a second support connection spaced apart from the first support connection along an upper portion of the frame. A linkage structure couples the frame to the rigid structure. The linkage structure includes a first linkage pair and a second linkage pair. The first links extend between the forward mount interface and the first support connection at a first angle with respect to the rotational axis. The second links extend between the aft mount interface and the second support connection at a second angle with respect to the rotational axis. As viewed from the side, a projection of the load vector of the first linkage pair and the second linkage pair intersect proximate the rotational axis of the engine between a forward end of an engine propulsive fan assembly and the front of a high pressure compressor of the engine.

In another aspect, a method of coupling an engine to an aircraft wing is provided. The method includes coupling a rigid structure to the aircraft wing. The rigid structure includes a forward mount interface and an aft mount interface. Each of the forward mount interface and an aft mount interface receives a thrust component of load. The method also includes coupling a frame about the engine such the frame surrounds a rotational axis of the engine. The frame includes a first support connection and a second support connection spaced apart along an upper portion of the frame. The method further includes coupling the first links of a linkage structure to the forward mount interface and the first support connection at a first angle with respect to the rotational axis, and coupling the second links of the linkage structure to the aft mount interface and the second support connection at a second angle with respect to the rotational axis. As viewed from the side, a projection of the load vector of the first links and the second links intersect proximate the rotational axis of the engine and a plane of a fan rotor.

In yet another aspect, an aircraft is provided. The aircraft includes a wing, an engine, a pylon coupled between the wing and the engine. The pylon includes a forward mount interface and an aft mount interface. Each of the forward mount interface and the aft mount interface receives a thrust component of load. The aircraft also includes a frame surrounding a rotational axis of the engine. The frame includes a first support connection and a second support connection spaced apart from the first support connection along an upper portion of the frame. A linkage structure couples the frame to the pylon. The linkage structure includes a first linkage pair and a second linkage pair. The first linkage pair extends between the forward mount interface and the first support connection at a first angle with respect to the rotational axis. The second linkage pair extends between the aft mount interface and the second support connection at a second angle with respect to the rotational axis. As viewed from the side, the load vector of the first links and the second links intersect proximate the rotational axis of the engine and a plane of a fan rotor.

DETAILED DESCRIPTION

Embodiments of the present disclosure relate to mounting systems for mounting turbine engine assemblies to an aircraft wing. More specifically, the mounting systems described herein are designed to reduce or eliminate backbone bending of the engine within the engine nacelle during certain engine operating conditions. In one embodiment, an aft leaning linkage structure is coupled between a pylon of the aircraft wing and the inner frame of the engine. As viewed from the side, the linkage structure includes at least first and second links that each includes load vectors extending therefrom and intersecting at a focal point proximate the rotational axis of the engine and a fan rotor near the engine inlet. The selection of the position of the point of intersection (as viewed from the side) of the load vectors of the first and second links proximate the engine axis facilitates reducing or eliminating backbone bending of the engine during various engine operational modes.

Locating the focal point of the mounting system at or near a location relative to the inlet loading and engine centerline reduces backbone bending to negligible levels, even in large turbofan engines that generate high thrust levels. Additionally, the mounting system described below couples the aircraft wing to the frame of the engine such that the mounting system is not coupled to the nacelle or the core cowl. Furthermore, the mounting system is capable of achieving this benefit while avoiding a substantial penalty in cost or weight typically associated with prior efforts to reduce backbone bending.

As used herein, the terms “axial” and “axially” refer to directions and orientations that extend substantially parallel to a centerline of the turbine engine. Moreover, the terms “radial” and “radially” refer to directions and orientations that extend substantially perpendicular to the centerline of the turbine engine.

FIG. 1is a perspective view of an aircraft1in accordance with an example embodiment of the present disclosure. In the example embodiment, aircraft1includes a fuselage2and a pair of laterally extending wings3. Each wing includes an engine4fixedly coupled to wing3through a pylon5.FIG. A1schematically represents a high-bypass turbofan engine10that may be used with aircraft1(shown inFIG. 1). Engine10is schematically represented as including a nacelle12and a core engine (module)14. A fan assembly16located in front of core engine14includes a spinner nose20projecting forwardly from an array of fan blades18. Core engine14is schematically represented as including a high-pressure compressor (HPC)22, a combustor24, a high-pressure turbine (HPT)26and a low-pressure turbine (LPT)28. A large portion of the air that enters fan assembly16is bypassed to the rear of engine10to generate additional engine thrust. The bypassed air passes through an annular-shaped bypass duct30between nacelle12and an inner core cowl36, and exits duct30through a fan exit nozzle32. Core cowl36defines the radially inward boundary of bypass duct30, and provides an aft core cowl transition surface to a primary exhaust nozzle38that extends aftward from core engine14. Nacelle12defines the radially outward boundary of bypass duct30, and the bypassed fan air flows between bypass duct flow surfaces defined by nacelle12and core cowl36before being exhausted through fan exit nozzle32.

Nacelle12is typically composed of three primary elements that define the external boundaries of nacelle12: an inlet assembly12A located upstream of the fan assembly16, a fan cowl12B interfacing with an engine fan case42that surrounds fan blades18, and a thrust reverser assembly12C located aft of fan cowl12B. Furthermore, core cowl36is a component of nacelle12and provides a shell around core engine12.

When installed on an aircraft, engine10is supported by a rigid aircraft structure, for example, a pylon (not shown inFIG. 1) that extends outward from the aircraft. In the case of an engine mounted to a wing, the pylon typically extends downwardly beneath the wing. Structural components of the pylon are connected to a frame115of core engine12that supports the rotating components of HPC22and turbines26and28. In the exemplary embodiment, frame115includes an engine frame or a fan frame. Frame115typically includes a forward frame adjacent HPC22, an aft frame adjacent turbines26and28, and an engine casing that connects the forward and aft frames. The engine casing is often referred to as the backbone of engine10. Aircraft engines of the type represented inFIG. 1are typically mounted and secured to an aircraft in two planes normal to engine centerline40. One mount is typically connected to the forward frame often just rearward of fan assembly16, and a second mount is typically connected to the aft frame near the turbine section.

During climb and certain aircraft operating modes, centerline40of engine10is pitched relative to the direction of approaching airflow, with the result that nacelle12can be subjected to upward aerodynamic loading. This aerodynamically-induced load, often referred to as the inlet load and represented by the vector Fi inFIG. 1A, is in addition to the thrust load, represented by the vector Ft inFIG. 1. These loads induce bending moments in the engine casing (backbone); with the result that the backbone is deflected (bends) from its concentric position about engine centerline40. Maintaining concentricity of the engine backbone about centerline40is important from the standpoint of minimizing blade tip clearances within HPC22and turbine sections26and28of engine10, which has the beneficial effect of improving engine specific fuel consumption (SFC) and fuel burn. In addition, reduced backbone bending reduces the incidence of blade tip rub encounters with the surrounding engine structures (including the fan case42), which promotes in-service performance retention. Engines with a longer interval for time on-wing to removal for service provide reduced service contract costs to their operators.

FIG. 2is a side view of an exemplary high-bypass turbofan engine10equipped with a mounting system100for mounting engine10to a wing52of an aircraft50.FIG. 2Ais a plan view taken of area2A (shown inFIG. 2).FIG. 2represents engine10in a nonoperational mode with the weight W of engine100being the only load acting on mounting system100.FIG. 3represents engine10in a take-off or cruise loading condition where thrust from engine10propels aircraft50at a high rate of speed down the runway prior to takeoff or through the air.FIG. 4represents engine10in a climbing or lift-off loading condition. Mounting system100can be installed in a high-bypass gas turbofan engine of the type represented inFIG. 1and therefore, as a matter of convenience, the same numbers used inFIG. 1to identify engine10and its components will be used inFIGS. 2 through 4to identify the same or functionally equivalent components. To facilitate the description of the system100provided below, the terms “vertical,” “horizontal,” “lateral,” “forward,” “aft,” “upper,” “lower,” “above,” “below,” etc., may be used in reference to the perspective of the installation and orientation of engine10on aircraft50, and therefore are relative terms that indicate the construction, installation and use of the invention and help to define the scope of the invention. However, it is within the scope of the invention that system100could be installed on an engine that markedly differs from engine10shown in the drawings, or installed at other points of aircraft50, for example, the fuselage. Finally, it is foreseeable that system100could find uses in applications other than aircraft engines.

As shown inFIG. 2, mounting system100includes a rigid structure, or pylon,102coupled to the engine support structure of aircraft wing52through, for example, a clevis and lug type connector with a uni-ball or a spherical bearing. Pylon102is also coupled to engine10and includes a forward mount interface104and an aft mount interface106. In the exemplary embodiment, each of forward mount interface104and aft mount interface106receives at least a portion of a thrust component of a load in certain engine operating modes.

Mounting system100also includes frame115surrounding rotational axis40of engine10. In the exemplary embodiment, frame115includes a first support connection108and a second support connection110spaced apart from first support connection108.

In the exemplary embodiment, mounting system100also includes a linkage structure114coupled between frame115and pylon104and configured to secure frame115to pylon102. Together, pylon102, frame115, and linkage structure114form a statically determinate structure. Linkage structure114includes at least a pair of first links116, a pair of second links118, and a pair of aft links119. As viewed in the side view ofFIG. 2A, both links of first links116form the same angle with respect to rotational axis40and both links of second links118form the same angle with respect to rotational axis40. As illustrated inFIG. 2A, however, first links116are angled inward with respect to rotational axis40, while second links118are approximately parallel to rotational axis40.

In the exemplary embodiment, first pair of symmetric links116(only one shown inFIG. 2) extend between a respective forward mount interface104on pylon102to first support connection108on frame115. More specifically, first linkage pair116is pivotally coupled at one end to pylon forward mount interface104and also pivotally coupled at the opposite end thereof to frame115first support connection108. In the exemplary embodiment, first linkage pair116is coupled between pylon102and frame115such that first linkage pair116defines a first angle α with respect to rotational axis40.

Similarly, second linkage pair118extends between aft mount interface106on pylon102and second support connection110on frame115. More specifically, second linkage pair118is pivotally coupled at one end to pylon aft mount interface106and is also pivotally coupled at the opposite end thereof to frame115second support connection110. In the exemplary embodiment, second linkage pair118is coupled between pylon102and frame115such that second linkage pair118defines a second angle β with respect to rotational axis40. As shown inFIG. 2, second angle β is less than first angle α.

Linkage structure114provides a connection between engine10and pylon102of aircraft wing52(or other suitable support structure) that significantly reduces backbone bending/deflection within core engine14that would otherwise result from thrust and inlet loads of the type previously described in reference toFIG. 1. Additionally, linkage structure114is coupled only between frame115and pylon102, and, therefore, is not coupled to any portion of nacelle12or fan casing42. In some embodiments, backbone bending/deflection may potentially be reduced to negligible levels or even zero.

As shown inFIG. 2, first linkage pair116includes a load vector120of a force (or forces within the same plane) transmitted through first linkage pair116. Similarly, second linkage pair118includes a load vector122of a force (or forces within the same plane) transmitted through second linkage pair118. In the exemplary embodiment, a projection of resultant load vectors120and122onto a vertical plane that extends through rotational axis40intersect at a point that lies within a planar area124. Planar area124, in a first embodiment, extends from an approximate axial midpoint of HPC22forward to approximately a tip of spinner nose20. In a second embodiment, planar area124extends from an approximate forward end of HPC22forward to a forward end of fan assembly16. In another embodiment, planar area124extends axially between an HPC22inlet and a fan assembly plane17. In other embodiments, a planar area124is angled approximately 25° to approximately 30° counterclockwise as viewed inFIG. 2. More specifically, a projection of load vectors120and122onto planar area124intersect with each other proximate rotational axis40at a focal point Pf along fan assembly plane17. Because load vectors120are not parallel with respect to each other and because load vectors122are approximately parallel to rotational axis40, they may not intersect rotational axis40at any point. However, the projections of load vectors120and122onto planar area124can intersect each other within the boundaries of planar area124proximate rotational axis40.

The capability of mounting system100to potentially reduce backbone bending/deflection to low values or zero can be further understood from reference toFIGS. 2-4.FIG. 2diagrammatically represents relative force vectors that exist solely as a result of the weight, W, of engine10(no engine operation), and indicates that the engine weight is shared between first and second links116and118and aft mount vertical reaction Ra at aft links119. More specifically,FIG. 2represents a first mode of engine10operation where engine10is either shut down, or all other forces acting upon engine10are balanced.

FIG. 3diagrammatically represents conditions that exist during a second mode of engine10operation, such as during take-off while aircraft50is traveling down the runway or during level flight when additional thrust is required. As shown inFIG. 3first and second links116and118of linkage structure114are subjected to additional forces resulting from engine thrust, Ft. Specifically, as can be seen inFIG. 3, because the focal point Pf is shown above the centerline, engine thrust Ft causes a reduced forward vertical reaction Rf at focal point Pf acting in an opposite direction to the weight force reaction W. Similarly, engine thrust Ft causes an aft reaction Ra at aft links119acting in an upward direction adding to the weight force reaction W. In the loading condition shown inFIG. 3, forward reaction Rf is reduced by an equal and opposite amount, together offsetting the thrust Ft moment. Additionally, the closer focal point Pf is to centerline40, the smaller the increase in aft reaction Ra from the balanced static load condition ofFIG. 2to the additional thrust condition inFIG. 3.

FIG. 4diagrammatically represents conditions that exist during a third mode of engine10operation, such as during lift-off or when aircraft50is climbing. As shown inFIG. 4, links116,118, and119of linkage structure114are subjected to additional forces resulting from engine thrust, Ft, but are also subjected to inlet load Fi caused by the increased angle of attack of aircraft50during lift-off and climbing load conditions. Specifically, as can be seen inFIG. 4, engine thrust Ft and inlet load Fi cause forward reaction Rf at focal point Pf acting in a downward direction opposite the weight force W reaction. Similarly, engine thrust Ft and inlet load Fi cause an increase in aft reaction Ra at aft links119acting in an upward direction adding to the weight force W reaction. In the loading condition shown inFIG. 4, forward reaction Rf is reduced or may become negative and the aft reaction Ra is increased as compared to the loading conditions shown inFIG. 3. In the exemplary embodiment, the more forward the focal point Pf, the smaller the increase in backbone bending due to inlet load Fi. Additionally, in the loading condition shown inFIG. 4, aft reaction Ra is larger than aft reaction Ra during the loading conditions shown inFIG. 3. As aircraft50reaches cruise and levels out, inlet load Fi decreases and forward and aft reactions Rf and Ra return to canceling the weight reaction W and thrust force Ft only, as shown inFIG. 3. In the exemplary embodiment, as described herein, the location of focal point Pf, forward of HPC22, frame115, and a leading edge of core cowl36facilitates reduced backbone bending of engine10during the lift-off and climbing load conditions.

In the exemplary embodiment, inlet load Fi, is indicated as being additionally present as a result of the aircraft being in a climb, during which nacelle12is subjected to upward aerodynamic loading as a result of centerline40of engine10being pitched upward relative to the direction of approaching airflow. Notably,FIG. 4represents loads for engine thrust and inlet loads, Ft and Fi, as imposing load moments in the same or opposite directions around focal point Pf of links116and118depending on whether the focal point Pf is above or below the engine center line40. The result is that their moments can cancel out each other to some extent if their distances relative to focal point Pf are appropriate for their respective magnitudes of Ft and Fi. By reducing the bending moment induced in engine10by the inlet load Fi, and the thrust force Ft, the backbone of core engine14will be subjected to less bending or deflection.

The magnitude of the load in first support structure108and second support structure110under the conditions represented inFIG. 4will vary depending upon the actual values of weight W, Thrust Ft and Inlet load Fi experienced during a flight. Consequently, the location of focal point Pf being forward of frame115is a preferred aspect of the invention in order to significantly reduce bending and deflection of the backbone of core engine14. In addition, the focal point Pf is preferably located in close proximity to the intersection of the engine thrust and inlet load vectors, Ft and Fi, so as to be located within the fan inlet assembly12A of engine10. In practice, the distributed mass of the engine also contributes to engine backbone bending and the best position of the focal point will be found within the area claimed to minimize all sources of backbone bending throughout an aircraft flight. More specifically, moving focal point Pf forward as compared to known engine designs minimizes bending. Even more specifically, load vectors120and122intersect at focal point Pf proximate the intersection of rotational axis40of engine10and fan assembly plane17near the engine inlet. The selection of the position of the point of intersection (as viewed from the side) of the load vectors of the first and second links proximate the engine axis facilitates reducing or eliminating backbone bending of the engine during various engine operational modes.

It should be understood that the system described is statically determinate and that “fail safe” considerations would include additional “waiting fail-safe” features or additional links, making a non-statically determinate system of the same performance with respect to reducing backbone bending.

From the foregoing, it should be appreciated that the location of the focal point, Pf, can be achieved with combinations and configurations of links and mounting locations that differ from what is represented in the Figures, and such other combinations and configurations are within the scope of the invention. Suitable alternatives can be readily ascertained by utilizing applied mathematics vector analysis to derive moments.

A technical effect of the invention is the ability to locate the focal point of the mounting system at or near a location relative to the inlet loading and engine centerline that can potentially reduce backbone bending to negligible levels, even in large turbofan engines that generate high thrust levels. Additionally, the mounting system couples the aircraft wing to the frame of the engine such that the mounting system is not coupled to the nacelle or the fan case. Furthermore, the mounting system is capable of achieving this benefit while avoiding a substantial penalty in cost or weight typically associated with prior efforts to reduce backbone bending.

The above-described embodiments of a method and system of coupling on engine to an aircraft wing through a pylon provides a cost-effective and reliable means for reducing loads in an aft mount of the pylon and reducing load variations during different modes of operation. More specifically, the methods and systems described herein also facilitate improving build clearances for the high pressure compressor and the high pressure turbine and permitting a decreased profile of fairing of the pylon. As a result, the methods and systems described herein facilitate coupling the engine to the aircraft in a cost-effective and reliable manner.

Exemplary embodiments of mounting systems are described above in detail. The mounting systems, and methods of operating such systems and devices are not limited to the specific embodiments described herein, but rather, components of systems and/or steps of the methods may be utilized independently and separately from other components and/or steps described herein. For example, the methods may also be used in combination with other systems requiring mounting of components, and are not limited to practice with only the systems and methods as described herein.