Patent Description:
Gas turbine engines are known, and typically include a fan delivering air into a bypass duct as propulsion air, and into a core engine housing. The core engine housing houses a compressor section which also receives air from the fan. The air is compressed and delivered into a combustor where it is mixed with fuel and ignited. Products of this combustion pass downstream over turbine rotors, driving them to rotate. The turbine rotors in turn rotate the fan and compressor rotors.

Historically the fan rotor was fixed to rotate at the same speed as a fan drive turbine rotor, which may also drive a low pressure compressor rotor. More recently a gear reduction has been incorporated between the fan drive turbine and the fan rotor, allowing the fan rotor to rotate at slower speeds than the fan drive turbine.

In modern gas turbine engines with such a gear reduction the fan case has been fixed to the core housing through a plurality of fan exit guide vanes, which provide the structural support between the fan case and the inner core housing.

<CIT> discloses a prior art gas turbine engine including a plurality of outlet guide vanes extending between an engine core housing and an outlet guide vane support region of a fan case.

<CIT> discloses prior art strut rods for structural guide vanes.

<CIT> discloses a prior art mounting arrangement for mounting a gas turbine engine to a vehicle.

<CIT> discloses a prior art ducted fan gas turbine engine.

<CIT> discloses a prior art bleed duct assembly for a gas turbine engine.

According to a first aspect of the present invention, there is provided a gas turbine engine as set forth in claim <NUM>.

In another embodiment according to the previous embodiment, a moment stiffness is defined about an axis extending perpendicular to a rotational axis of the gas turbine engine. A moment stiffness ratio of the moment stiffness of the plurality of fan exit guide vanes to a moment stiffness of the combination of the plurality of A-frames, the compressor wall, and the fan intermediate case being greater than or equal to <NUM> and less than or equal to <NUM>.

In another embodiment according to any of the previous embodiments, the moment stiffness ratio is less than or equal to <NUM>.

In another embodiment according to any of the previous embodiments, a first gap is defined between a leading edge of the fan exit guide vanes at a radially innermost point radially outward of the splitter wall and a leading edge of the plurality of A-frames at a radially inner location radially inward of the splitter wall. A second gap is defined between a trailing edge of the fan blades and a leading edge of the plurality of fan exit guide vanes at the radially inward point radially outward of the splitter wall each of the gaps being defined by extending a line from the respective points in a radially outward direction. Then a distance is measured with a line extending parallel to the rotational axis of the engine, and a first gap ratio of the second gap to the first gap is between <NUM> and <NUM>.

In another embodiment according to any of the previous embodiments, a third gap is defined between the trailing edge of the fan blades and a leading edge of the fan exit guide vane at a radially outermost location radially inward of the fan case. The third gap is also defined by extending a line from the respective points in a radially outward direction. Then a distance is measured with a line extending parallel to the rotational axis of the engine. A second gap ratio of the first gap to the third gap is greater than or equal to <NUM> and less than or equal to <NUM>.

In another embodiment according to any of the previous embodiments, a third gap ratio of the second gap to the third gap is greater than or equal to <NUM> and less than or equal to <NUM>.

In another embodiment according to any of the previous embodiments, the lateral stiffness ratio is less than or equal to <NUM>.

In another embodiment according to any of the previous embodiments, a ratio of a quantity of the plurality of fan exit guide vanes to a quantity of the legs of the plurality of A-frames being greater than or equal to <NUM>.

In another embodiment according to any of the previous embodiments, the low pressure compressor has between four to six stages.

According to a further aspect of the present invention, there is provided a gas turbine engine as set forth in claim <NUM>.

In another embodiment according to any of the previous embodiments, the moment ratio is less than or equal to <NUM>.

In another embodiment according to any of the previous embodiments, a ratio of a quantity of the plurality of fan exit guide vanes to a quantity of the legs of the plurality of A-frames is greater than or equal to <NUM>.

In this specification the following non-SI units are used, which may be converted to the respective SI or metric unit according to the following conversion: Fahrenheit (°F) to Kelvin (K): T(K) = (T(°F) + <NUM> °F) × <NUM>/<NUM>; Rankin (°R) to Kelvin: T(K) = T(°R) × <NUM>/<NUM>; inch (in) to meter (m): m = in × <NUM>,<NUM>.

The fan section <NUM> may include a single-stage fan <NUM> having a plurality of fan blades <NUM>. The fan blades <NUM> may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan <NUM> drives air along a bypass flow path B in a bypass duct <NUM> defined within a housing <NUM> such as a fan case or nacelle, and also drives air along a core flow path C for compression and communication into the combustor section <NUM> then expansion through the turbine section <NUM>. A splitter <NUM> aft of the fan <NUM> divides the air between the bypass flow path B and the core flow path C. The housing <NUM> may surround the fan <NUM> to establish an outer diameter of the bypass duct <NUM>. The splitter <NUM> may establish an inner diameter of the bypass duct <NUM>. Although depicted as a two-spool turbofan gas turbine engine in the disclosed nonlimiting embodiment, it should be understood that the concepts described herein are not limited to use with two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. The engine <NUM> may incorporate a variable area nozzle for varying an exit area of the bypass flow path B and/or a thrust reverser for generating reverse thrust.

The inner shaft <NUM> is connected to the fan <NUM> through a speed change mechanism, which in the exemplary gas turbine engine <NUM> is illustrated as a geared architecture <NUM> to drive the fan <NUM> at a lower speed than the low speed spool <NUM>. The inner shaft <NUM> may interconnect the low pressure compressor <NUM> and low pressure turbine <NUM> such that the low pressure compressor <NUM> and low pressure turbine <NUM> are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine <NUM> drives both the fan <NUM> and low pressure compressor <NUM> through the geared architecture <NUM> such that the fan <NUM> and low pressure compressor <NUM> are rotatable at a common speed. Although this application discloses geared architecture <NUM>, its teaching may benefit direct drive engines having no geared architecture.

The fan <NUM> may have at least <NUM> fan blades <NUM> but no more than <NUM> or <NUM> fan blades <NUM>. In examples, the fan <NUM> may have between <NUM> and <NUM> fan blades <NUM>, such as <NUM> fan blades <NUM>. An exemplary fan size measurement is a maximum radius between the tips of the fan blades <NUM> and the engine central longitudinal axis A. The maximum radius of the fan blades <NUM> can be at least <NUM> inches, or more narrowly no more than <NUM> inches. For example, the maximum radius of the fan blades <NUM> can be between <NUM> inches and <NUM> inches, such as between <NUM> inches and <NUM> inches. Another exemplary fan size measurement is a hub radius, which is defined as distance between a hub of the fan <NUM> at a location of the leading edges of the fan blades <NUM> and the engine central longitudinal axis A. The fan blades <NUM> may establish a fan hub-to-tip ratio, which is defined as a ratio of the hub radius divided by the maximum radius of the fan <NUM>. The fan hub-to-tip ratio can be less than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, such as between <NUM> and <NUM>. The combination of fan blade counts and fan hub-to-tip ratios disclosed herein can provide the engine <NUM> with a relatively compact fan arrangement.

The low pressure compressor <NUM>, high pressure compressor <NUM>, high pressure turbine <NUM> and low pressure turbine <NUM> each include one or more stages having a row of rotatable airfoils. Each stage may include a row of vanes adjacent the rotatable airfoils. The rotatable airfoils are schematically indicated at <NUM>, and the vanes are schematically indicated at <NUM>.

The low pressure compressor <NUM> and low pressure turbine <NUM> can include an equal number of stages. For example, the engine <NUM> can include a three-stage low pressure compressor <NUM>, an eight-stage high pressure compressor <NUM>, a two-stage high pressure turbine <NUM>, and a three-stage low pressure turbine <NUM> to provide a total of sixteen stages. In other examples, the low pressure compressor <NUM> includes a different (e.g., greater) number of stages than the low pressure turbine <NUM>. For example, the engine <NUM> can include a five-stage low pressure compressor <NUM>, a nine-stage high pressure compressor <NUM>, a two-stage high pressure turbine <NUM>, and a four-stage low pressure turbine <NUM> to provide a total of twenty stages. In other embodiments, the engine <NUM> includes a four-stage low pressure compressor <NUM>, a nine-stage high pressure compressor <NUM>, a two-stage high pressure turbine <NUM>, and a three-stage low pressure turbine <NUM> to provide a total of eighteen stages. It should be understood that the engine <NUM> can incorporate other compressor and turbine stage counts, including any combination of stages disclosed herein.

The engine <NUM> may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to <NUM> and less than or equal to about <NUM>, or more narrowly can be less than or equal to <NUM>. The geared architecture <NUM> may be an epicyclic gear train, such as a planetary gear system or a star gear system. The epicyclic gear train may include a sun gear, a ring gear, a plurality of intermediate gears meshing with the sun gear and ring gear, and a carrier that supports the intermediate gears. The sun gear may provide an input to the gear train. The ring gear (e.g., star gear system) or carrier (e.g., planetary gear system) may provide an output of the gear train to drive the fan <NUM>. A gear reduction ratio may be greater than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, and in some embodiments the gear reduction ratio is greater than or equal to <NUM>. The fan diameter is significantly larger than that of the low pressure compressor <NUM>. The low pressure turbine <NUM> can have a pressure ratio that is greater than or equal to <NUM> and in some embodiments is greater than or equal to <NUM>. All of these parameters are measured at the cruise condition described below.

The engine parameters described above, and those in the next paragraph are measured at this condition unless otherwise specified.

"Fan pressure ratio" is the pressure ratio across the fan blade <NUM> alone, without a Fan Exit Guide Vane ("FEGV") system. A distance is established in a radial direction between the inner and outer diameters of the bypass duct <NUM> at an axial position corresponding to a leading edge of the splitter <NUM> relative to the engine central longitudinal axis A. The fan pressure ratio is a spanwise average of the pressure ratios measured across the fan blade <NUM> alone over radial positions corresponding to the distance. The fan pressure ratio can be less than or equal to <NUM>, or more narrowly greater than or equal to <NUM>, such as between <NUM> and <NUM>. "Corrected fan tip speed" is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R) / (<NUM> °R)]<NUM>. The corrected fan tip speed can be less than or equal to <NUM> ft / second (<NUM> meters/second), and can be greater than or equal to <NUM> ft / second (<NUM> meters/second).

The fan <NUM>, low pressure compressor <NUM> and high pressure compressor <NUM> can provide different amounts of compression of the incoming airflow that is delivered downstream to the turbine section <NUM> and cooperate to establish an overall pressure ratio (OPR). The OPR is a product of the fan pressure ratio across a root (i.e., <NUM>% span) of the fan blade <NUM> alone, a pressure ratio across the low pressure compressor <NUM> and a pressure ratio across the high pressure compressor <NUM>. The pressure ratio of the low pressure compressor <NUM> is measured as the pressure at the exit of the low pressure compressor <NUM> divided by the pressure at the inlet of the low pressure compressor <NUM>. In examples, a product of the pressure ratio of the low pressure compressor <NUM> and the fan pressure ratio is between <NUM> and <NUM>, or more narrowly is between <NUM> and <NUM>. The pressure ratio of the high pressure compressor ratio <NUM> is measured as the pressure at the exit of the high pressure compressor <NUM> divided by the pressure at the inlet of the high pressure compressor <NUM>. In examples, the pressure ratio of the high pressure compressor <NUM> is between <NUM> and <NUM>, or more narrowly is between <NUM> and <NUM>. The OPR can be equal to or greater than <NUM>, and can be less than or equal to <NUM>, such as between <NUM> and <NUM>. The overall and compressor pressure ratios disclosed herein are measured at the cruise condition described above, and can be utilized in two-spool architectures such as the engine <NUM> as well as three-spool engine architectures.

The engine <NUM> establishes a turbine entry temperature (TET). The TET is defined as a maximum temperature of combustion products communicated to an inlet of the turbine section <NUM> at a maximum takeoff (MTO) condition. The inlet is established at the leading edges of the axially forwardmost row of airfoils of the turbine section <NUM>, and MTO is measured at maximum thrust of the engine <NUM> at static sea-level and <NUM> degrees Fahrenheit (°F). The TET may be greater than or equal to <NUM> °F, or more narrowly less than or equal to <NUM> °F, such as between <NUM> °F and <NUM> °F. The relatively high TET can be utilized in combination with the other techniques disclosed herein to provide a compact turbine arrangement.

The engine <NUM> establishes an exhaust gas temperature (EGT). The EGT is defined as a maximum temperature of combustion products in the core flow path C communicated to at the trailing edges of the axially aftmost row of airfoils of the turbine section <NUM> at the MTO condition. The EGT may be less than or equal to <NUM> °F, or more narrowly greater than or equal to <NUM> °F, such as between <NUM> °F and <NUM> °F. The relatively low EGT can be utilized in combination with the other techniques disclosed herein to reduce fuel consumption.

<FIG> shows an engine <NUM>, which may operate similar to the engine <NUM> of <FIG>. A shaft <NUM> is driven by a fan drive turbine to drive a fan rotor <NUM> through a gear reduction <NUM>. The drive connection here may be generally as described above with regard to <FIG>. A fan case <NUM> surrounds the fan rotor <NUM>, and an core engine <NUM> may include a splitter wall <NUM> that surrounds compressor housing wall <NUM> which houses a low pressure compressor <NUM>, and a high pressure compressor <NUM>, and combustor and turbine sections (not shown in this Figure). The core engine <NUM> must be rigidly connected to the fan case <NUM>, to address torque and other loads.

Applicant has previously developed a geared gas turbine engine. In this first generation engine the fan case was connected to the core engine through a plurality of fan exit guide vanes. Each of these fan exit guide vanes were structural elements that provided a load path between the fan case and the core engine.

In engine <NUM>, as will be described below, there are fewer structural fan exit guide vanes <NUM>. A-frames <NUM> have been added to provide additional rigidity.

As shown, the low pressure compressor <NUM> has five rotating stages. In embodiments the low pressure compressor may have four to six stages, which is longer than the first generation gas turbine engine manufactured by Applicant mentioned above. With such a long low pressure compressor <NUM>, mounting challenges are raised.

As can been seen from <FIG>, the fan exit guide vanes <NUM>/<NUM> extend from an inner point <NUM> at an angle to a radially outer point <NUM>, with the angle having a component in a radially outer direction, and another component in an axially aft direction.

Conversely, the A-frames <NUM> extend from a radially inner point <NUM> attached to the core engine <NUM>, radially outwardly at an angle to an outer point <NUM> connected to the fan case <NUM>. The angle of the A-frame <NUM> has a component in a radially outer direction and another component in an axially forward direction.

As shown in <FIG>, the core engine <NUM> includes a fan intermediate case <NUM> including a plurality of struts, and having a mount bracket <NUM>. A non-structural guide vane <NUM> is attached to the bracket <NUM> through pins <NUM>. In this manner, the non-structural guide vanes <NUM> can "float" or adjust radially relative to the bracket <NUM>, but are prevented from moving circumferentially. Alternatively, the non-structural guide vanes may be fixed to bracket <NUM>.

<FIG> shows a detail of the mount of a structural guide vane <NUM> to the fan intermediate case <NUM> and to the bracket <NUM> through a first pin <NUM> preventing circumferential movement, and a second pin <NUM> preventing radial movement.

Returning to <FIG>, it can be seen that the A-frames <NUM> are attached at inner ends <NUM> at a bracket <NUM> which is fixed with a compressor intermediate case <NUM> having a plurality of struts <NUM>. The compressor intermediate cases <NUM> is intermediate the low pressure compressor <NUM> and a high pressure compressor <NUM>, and includes struts. This view shows one such strut.

Although specific mount locations are shown, other connections between the fan exit guide vanes <NUM>/<NUM> and A-frames <NUM> to the core engine <NUM> may be utilized. For purposes of this application, the core engine <NUM> is defined to include at least the compressor housing wall <NUM>, the fan intermediate case <NUM>, the compressor intermediate case <NUM>, the low pressure compressor <NUM>, the high pressure compressor <NUM> and a combustor end turbine section, not shown, but which may be as disclosed with regard to <FIG>.

As shown in <FIG>, the A-frames <NUM> comprise two rigid members <NUM> and <NUM> which extend from a connection point <NUM> at the fan case <NUM> inwardly to connection points <NUM> with the compressor housing <NUM>. As can be seen, the members <NUM> extend away from each other moving away from the connection point <NUM>, and in opposed directions/angles relative to a plane drawn through connection point <NUM>, parallel to a center axis X. In <FIG> there are sixteen structural guide vanes <NUM> illustrated. Other numbers can be used. They are rigidly connected at <NUM> to the fan case <NUM>. It should be understood that <FIG> is an over simplification. In fact, as mentioned above, the fan exit guide vanes <NUM> and <NUM> and A-frames <NUM> may be attached as shown in <FIG>. The view of <FIG> is shown simply to illustrate some general relationships. In embodiments, the engine <NUM> can have more or less than eight of the A-frame legs <NUM> and <NUM>.

As shown in <FIG>, there are forty-eight fan exit guide vanes total, with sixteen of the fan exit guide vanes being structural vanes <NUM>. Intermediate each structural fan exit guide vane <NUM> are two non-structural guide vanes <NUM>. The non-structural guide vanes <NUM> are rigidly connected to the fan case <NUM> at <NUM>, but as mentioned may float relative to the wall <NUM> at a radially inner point <NUM>. Alternatively, inner point <NUM> may be fixed.

Since the guide vanes <NUM> are not structural, they can provide an acoustic function. As shown in <FIG>, one of the non-structural fan exit guide vanes <NUM> is illustrated. The guide vane <NUM> has a pressure wall <NUM> and a suction wall <NUM>. An outer skin <NUM> is perforated at <NUM> and provided over a plurality of chambers <NUM>. The chambers <NUM> can have any number of shapes including honeycomb, or other cross-sections. The perforations <NUM> could be circular, but can be other shapes, including elongated slots.

In embodiments the structural guide vanes <NUM> may include <NUM>% to <NUM>% of the total fan exit guide vanes. In other embodiments the structural fan exit guide vanes <NUM> may include <NUM>% to <NUM>% of the total fan exit guide vanes. In yet another embodiment, the structural fan exit guide vanes provide <NUM>% to <NUM>% of the total fan exit guide vanes.

The details of the fan exit guide vanes and A-frame structure as disclosed above are disclosed and claimed in co-pending <CIT> and owned by the Applicant of this application.

Stiffness relationships can be defined between the combined stiffness of the fan exit guide vanes <NUM> and <NUM>, compared to the combined stiffness of A-frames <NUM>, intermediate case <NUM> and compressor wall <NUM>. Note, if the A-frames <NUM> are attached to the compressor intermediate case <NUM>, it is also considered as part of the compressor wall <NUM>, including its struts <NUM>. A lateral stiffness can be defined in a radial direction. A ratio of lateral stiffness of the combination of structural and non-structural fan exit guide vane to a lateral stiffness of a combination of the A-frames, the compressor wall <NUM>, and the fan intermediate case <NUM> is greater than or equal to <NUM>, and less than or equal to <NUM>. In embodiments, it is less than or equal to <NUM>, and in further embodiments it is less than or equal to <NUM>. A moment stiffness may be defined as a moment about an axis perpendicular to a rotational axis of the engine. A ratio of the moment stiffness of the combination of the structural and non-structural fan exit guide vanes to the moment stiffness of the combination of the A-frame, compressor wall <NUM>, and case <NUM> is greater than or equal to <NUM> and less than or equal to <NUM>. In embodiments, it is less than or equal to <NUM>, and in further embodiments less than or equal to <NUM>.

A quantity of the combined structural guide vanes to a ratio of the quantity of the individual legs <NUM>/<NUM> in the A-frames <NUM> should be greater than or equal to <NUM>.

<FIG> shows geometric relationships incorporated into this disclosure.

A first gap A is defined between a leading edge <NUM> of the fan exit guide vanes at a radially innermost point outward of wall <NUM> and a leading edge <NUM> of the A-frames <NUM>. A second gap B is defined between the trailing edge <NUM> of the fan blade and a leading edge <NUM> of the fan exit guide vanes at the radially innermost point. A third gap C is defined between the trailing edge <NUM> of the fan blades and a leading edge <NUM> of the fan exit guide vanes at a radially outermost location radially inward of the fan case <NUM>. Note that the fan blade may having a twisting shape such that its trailing edge is defined at a rearmost point on the blades. Each of the gaps is defined by extending a line from the mentioned point in a radially outward direction, and then measuring the distance with a line extending parallel to the rotational axis X of the engine.

Low pressure compressor <NUM> is shown as is a gear reduction <NUM>. As is clear this view is quite schematic.

These gaps impact how the structure works together for fan blade clearance and noise, as well as overcoming stresses and other operational challenges. Three gap ratios may be defined as B/A, A/C and B/C. In one embodiment, the gap ratio A/C was greater than or equal to <NUM> and less than or equal to <NUM>. In embodiments the gap ratio B/C is greater than or equal to <NUM> and less than or equal to <NUM>. In embodiments the gap ratio B/A is greater than or equal to <NUM> and less than or equal to <NUM>.

The stiffness ratios and gap ratios may also be influenced by a diameter of the engine and a center line of the geared architecture. In one embodiment the center line of the gear reduction may be slightly aft utilizing a planetary gear relative to the center line of the first generation's star gear configuration in Applicant's first generation engine. Moving this center line forward may cause the fan to moved forward. Additional low pressure compressor stages also change the gap ratios.

Claim 1:
A gas turbine engine (<NUM>) comprising:
a fan rotor (<NUM>) comprising fan blades (<NUM>) driven by a fan drive turbine about an axis through a gear reduction (<NUM>) to reduce a speed of said fan rotor (<NUM>) relative to a speed of said fan drive turbine;
a fan case (<NUM>) surrounding said fan rotor (<NUM>), and an inner core housing surrounding a compressor section (<NUM>), including a low pressure compressor (<NUM>);
said fan rotor (<NUM>) delivering air into a bypass duct (<NUM>) defined between said fan case (<NUM>) and a splitter wall (<NUM>) of said inner core housing, and a rigid connection between said fan case (<NUM>) and said inner core housing including a plurality of A-frames (<NUM>) each including a pair of legs (<NUM>, <NUM>) rigidly connected at a connection point (<NUM>) to said fan case (<NUM>), and each leg (<NUM>, <NUM>) in said pair extending away from said connection point (<NUM>) in opposed circumferential directions to be connected to a compressor wall (<NUM>) of said inner core housing to form an A-shape; and
the rigid connection also including a plurality of fan exit guide vanes (<NUM>, <NUM>) rigidly connected to said fan case (<NUM>), the plurality of fan exit guide vanes (<NUM>, <NUM>) including a plurality of structural guide vanes (<NUM>) and a plurality of non-structural guide vanes (<NUM>), each of the plurality of non-structural guide vanes (<NUM>) comprising an outer skin (<NUM>) with perforations (<NUM>) and provided over a plurality of chambers (<NUM>); and
a lateral stiffness is defined in a radial direction, a lateral stiffness ratio of the lateral stiffness of the plurality of fan exit guide vanes (<NUM>, <NUM>) and a lateral stiffness of a combination of the plurality of A-frame, the compressor wall (<NUM>), and a fan intermediate case (<NUM>) which is forward of the low pressure compressor (<NUM>, <NUM>) being greater than or equal to <NUM> and less than or equal to <NUM>.