Patent Publication Number: US-2023151741-A1

Title: Triangular-frame connection between fan case and core housing in a gas turbine engine

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Application No. 63/278,833 filed Nov. 12, 2021. 
    
    
     BACKGROUND OF THE INVENTION 
     This application relates to incorporating frame connections and fan exit guide vanes connections between a fan case and a core housing in a gas turbine engine. 
     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. 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. 
     SUMMARY OF THE INVENTION 
     In a featured embodiment, a gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. The fan rotor delivers air into a bypass duct defined between the fan case and the core engine. A rigid connection is between the fan case and the core engine includes three triangular-frame connecting members rigidly connected to the fan case at a fan case connection point, and to the core engine at a core engine connection point. The triangular-frame connecting members each are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other. A plurality of non-structural fan exit guide vanes and the non-structural fan exit guide vanes are provided with an acoustic feature to reduce noise. The non-structural fan exit guide vanes are rigidly mounted to at least one of the fan case and the core engine. 
     In another embodiment according to the previous embodiment, the acoustic feature includes the non-structural fan exit guide vanes being formed with chambers and a covering perforated face sheet. 
     In another embodiment according to the previous embodiment, there are also structural fan exit guide vanes rigidly secured to both the fan case and the core engine. The structural fan exit guide vanes include greater than 0 and less than 55% of a total number of fan exit guide vanes including the non-structural fan exit guide vane. 
     In another embodiment according to the previous embodiment, the structural fan exit guide vanes include 15-30% of the total fan exit guide vanes. 
     In another embodiment according to the previous embodiment, the low pressure compressor has four to six stages. 
     In another embodiment according to the previous embodiment, there is an outer diameter to the core engine at the location of the connection points to the rigid legs of the three triangular frame connecting members. There is a distance between the legs in each triangular frame connecting member at the connection points where the legs are attached to the core engine. The distance is greater than the outer diameter. 
     In another embodiment according to the previous embodiment, the core engine includes a fan intermediate case forward of the low pressure compressor. The structural fan guide vanes are rigidly connected to the fan intermediate case. 
     In another embodiment according to the previous embodiment, the core engine includes a compressor intermediate case intermediate the low pressure compressor and a high pressure compressor. The connection points between the triangular frame connecting members are to the compressor intermediate case. 
     In another embodiment according to the previous embodiment, the core engine includes a compressor intermediate case intermediate the low pressure compressor and a high pressure compressor. The connection points between the triangular frame connecting members are to the compressor intermediate case. 
     In another embodiment according to the previous embodiment, the connection point between one of the three triangular frame connected members and the fan case is at a top dead center position. The connection point to the fan case for the other two of the triangular frame connecting member are on circumferentially spaced sides of a bottom dead center position. 
     In another embodiment according to the previous embodiment, the connection point between one of the three triangular frame connecting members and the fan case is at a bottom dead center position. The connection points for the other two of the three triangular frame connecting members are on circumferentially spaced side of a top dead center position. 
     In another embodiment according to the previous embodiment, two rigid links connect one of the legs from two of the three triangular frame connecting members spaced circumferentially about the top dead center position to the fan case, with the two rigid links on opposed sides of the top dead center position. 
     In another featured embodiment, a gas turbine engine includes a fan rotor driven by a fan drive turbine about an axis through a gear reduction to reduce a speed of the fan rotor relative to a speed of the fan drive turbine. A fan case surrounds the fan rotor, and a core engine with a compressor section, including a low pressure compressor. Three triangular frame connecting members, the triangular-frame connecting members are defined by two rigid legs which extend between the fan case and to the core engine, along directions each have a component extending radially inwardly and a component in opposed circumferential directions to each other the two rigid legs being fixed to both the fan case and the core engine at connection points. A plurality of fan exit guide vanes are rigidly connected to the fan case, and connected to the core engine. The two rigid legs in each the triangular frame connecting members extend from the fan case to the core engine extend in a direction with a radially inward component and an axially aft direction. 
     In another embodiment according to the previous embodiment, the low pressure compressor has four to six stages. 
     In another embodiment according to the previous embodiment, the core engine includes a fan intermediate case forward of the low pressure compressor. The structural fan guide vanes are rigidly connected to the fan intermediate case. 
     In another embodiment according to the previous embodiment, the connection point between one of the three triangular frame connected members and the fan case is at a top dead center position. The connection point to the fan case for the other two of the triangular frame connecting member are on circumferentially spaced sides of a bottom dead center position. 
     In another embodiment according to the previous embodiment, the connection point between one of the three triangular frame connecting members and the fan case is at a bottom dead center position. The connection points for the other two of the three triangular frame connecting members are on circumferentially spaced side of a top dead center position. Two rigid links connect one of the rigid legs from two of the three triangular frame connecting members spaced circumferentially about the top dead center position to the fan case, with the two rigid links being on opposed sides of the top dead center position. 
     In another embodiment according to the previous embodiment, the core engine includes a compressor intermediate case intermediate the low pressure compressor and a high pressure compressor. The connection points between the triangular frame connecting members to the core engine are to the compressor intermediate case. 
     In another embodiment according to the previous embodiment, there is an outer diameter to the core engine at the location of the connection points to the rigid legs of the three triangular frame connecting members. There is a distance between the legs in each triangular frame connecting member at the connection points where the legs are attached to the core engine. The distance is greater than the outer diameter. 
     In another embodiment according to the previous embodiment, the plurality of fan exit guide vanes include structural and non-structural fan exit guide vanes. 
     The present disclosure may include any one or more of the individual features disclosed above and/or below alone or in any combination thereof. 
     These and other features of the present invention can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    schematically shows a gas turbine engine. 
         FIG.  2 A  schematically shows details of the connection between a fan case and an inner core housing. 
         FIG.  2 B  shows another connection detail. 
         FIG.  3 A  shows details of a first embodiment frame connection. 
         FIG.  3 B  shows details of an alternative frame connection. 
         FIG.  4    shows a detail of a plurality of fan exit guide vanes. 
         FIG.  5    shows details of one fan exit guide vane. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    schematically illustrates a gas turbine engine  20 . The gas turbine engine  20  is disclosed herein as a two-spool turbofan that generally incorporates a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . The fan section  22  may include a single-stage fan  42  having a plurality of fan blades  43 . The fan blades  43  may have a fixed stagger angle or may have a variable pitch to direct incoming airflow from an engine inlet. The fan  42  drives air along a bypass flow path B in a bypass duct  13  defined within a housing  15  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  26  then expansion through the turbine section  28 . A splitter  29  aft of the fan  42  divides the air between the bypass flow path B and the core flow path C. The housing  15  may surround the fan  42  to establish an outer diameter of the bypass duct  13 . The splitter  29  may establish an inner diameter of the bypass duct  13 . Although depicted as a two-spool 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 two-spool turbofans as the teachings may be applied to other types of turbine engines including three-spool architectures. The engine  20  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 exemplary engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided, and the location of bearing systems  38  may be varied as appropriate to the application. 
     The low speed spool  30  generally includes an inner shaft  40  that interconnects, a first (or low) pressure compressor  44  and a first (or low) pressure turbine  46 . The inner shaft  40  is connected to the fan  42  through a speed change mechanism, which in the exemplary gas turbine engine  20  is illustrated as a geared architecture  48  to drive the fan  42  at a lower speed than the low speed spool  30 . The inner shaft  40  may interconnect the low pressure compressor  44  and low pressure turbine  46  such that the low pressure compressor  44  and low pressure turbine  46  are rotatable at a common speed and in a common direction. In other embodiments, the low pressure turbine  46  drives both the fan  42  and low pressure compressor  44  through the geared architecture  48  such that the fan  42  and low pressure compressor  44  are rotatable at a common speed. Although this application discloses geared architecture  48 , its teaching may benefit direct drive engines having no geared architecture. The high speed spool  32  includes an outer shaft  50  that interconnects a second (or high) pressure compressor  52  and a second (or high) pressure turbine  54 . A combustor  56  is arranged in the exemplary gas turbine  20  between the high pressure compressor  52  and the high pressure turbine  54 . A mid-turbine frame  57  of the engine static structure  36  may be arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  57  further supports bearing systems  38  in the turbine section  28 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via bearing systems  38  about the engine central longitudinal axis A which is collinear with their longitudinal axes. 
     Airflow in the core flow path C is compressed by the low pressure compressor  44  then the high pressure compressor  52 , mixed and burned with fuel in the combustor  56 , then expanded through the high pressure turbine  54  and low pressure turbine  46 . The mid-turbine frame  57  includes airfoils  59  which are in the core flow path C. The turbines  46 ,  54  rotationally drive the respective low speed spool  30  and high speed spool  32  in response to the expansion. It will be appreciated that each of the positions of the fan section  22 , compressor section  24 , combustor section  26 , turbine section  28 , and fan drive gear system  48  may be varied. For example, gear system  48  may be located aft of the low pressure compressor, or aft of the combustor section  26  or even aft of turbine section  28 , and fan  42  may be positioned forward or aft of the location of gear system  48 . 
     The fan  42  may have at least  10  fan blades  43  but no more than 20 or 24 fan blades  43 . In examples, the fan  42  may have between 12 and 18 fan blades  43 , such as 14 fan blades  43 . An exemplary fan size measurement is a maximum radius between the tips of the fan blades  43  and the engine central longitudinal axis A. The maximum radius of the fan blades  43  can be at least 38 inches, or more narrowly no more than 75 inches. For example, the maximum radius of the fan blades  43  can be between 45 inches and 60 inches, such as between 50 inches and 55 inches. Another exemplary fan size measurement is a hub radius, which is defined as distance between a hub of the fan  42  at a location of the leading edges of the fan blades  43  and the engine central longitudinal axis A. The fan blades  43  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  42 . The fan hub-to-tip ratio can be less than or equal to 0.35, or more narrowly greater than or equal to 0.20, such as between 0.25 and 0.30. The combination of fan blade counts and fan hub-to-tip ratios disclosed herein can provide the engine  20  with a relatively compact fan arrangement. 
     The low pressure compressor  44 , high pressure compressor  52 , high pressure turbine  54  and low pressure turbine  46  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  47 , and the vanes are schematically indicated at  49 . 
     The low pressure compressor  44  and low pressure turbine  46  can include an equal number of stages. For example, the engine  20  can include a three-stage low pressure compressor  44 , an eight-stage high pressure compressor  52 , a two-stage high pressure turbine  54 , and a three-stage low pressure turbine  46  to provide a total of sixteen stages. In other examples, the low pressure compressor  44  includes a different (e.g., greater) number of stages than the low pressure turbine  46 . For example, the engine  20  can include a five-stage low pressure compressor  44 , a nine-stage high pressure compressor  52 , a two-stage high pressure turbine  54 , and a four-stage low pressure turbine  46  to provide a total of twenty stages. In other embodiments, the engine  20  includes a four-stage low pressure compressor  44 , a nine-stage high pressure compressor  52 , a two-stage high pressure turbine  54 , and a three-stage low pressure turbine  46  to provide a total of eighteen stages. It should be understood that the engine  20  can incorporate other compressor and turbine stage counts, including any combination of stages disclosed herein. 
     The engine  20  may be a high-bypass geared aircraft engine. The bypass ratio can be greater than or equal to 10.0 and less than or equal to about 18.0, or more narrowly can be less than or equal to 16.0. The geared architecture  48  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  42 . A gear reduction ratio may be greater than or equal to 2.3, or more narrowly greater than or equal to 3.0, and in some embodiments the gear reduction ratio is greater than or equal to 3.4. The gear reduction ratio may be less than or equal to 4.0. The fan diameter is significantly larger than that of the low pressure compressor  44 . The low pressure turbine  46  can have a pressure ratio that is greater than or equal to 8.0 and in some embodiments is greater than or equal to 10.0. The low pressure turbine pressure ratio can be less than or equal to 13.0, or more narrowly less than or equal to 12.0. Low pressure turbine  46  pressure ratio is pressure measured prior to an inlet of low pressure turbine  46  as related to the pressure at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. It should be understood, however, that the above parameters are only exemplary of one embodiment of a geared architecture engine and that the present invention is applicable to other gas turbine engines including direct drive turbofans. All of these parameters are measured at the cruise condition described below. 
     A significant amount of thrust is provided by the bypass flow B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and 35,000 ft (10,668 meters), with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (&#39;TSFC&#39;)”—is the industry standard parameter of lbm of fuel being burned divided by lbf of thrust the engine produces at that minimum point. 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  43  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  13  at an axial position corresponding to a leading edge of the splitter  29  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  43  alone over radial positions corresponding to the distance. The fan pressure ratio can be less than or equal to 1.45, or more narrowly greater than or equal to 1.25, such as between 1.30 and 1.40. “Corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram ° R)/(518.7° R)] 0.5 . The corrected fan tip speed can be less than or equal to 1150.0 ft/second (350.5 meters/second), and can be greater than or equal to 1000.0 ft/second (304.8 meters/second). 
     The fan  42 , low pressure compressor  44  and high pressure compressor  52  can provide different amounts of compression of the incoming airflow that is delivered downstream to the turbine section  28  and cooperate to establish an overall pressure ratio (OPR). The OPR is a product of the fan pressure ratio across a root (i.e., 0% span) of the fan blade  43  alone, a pressure ratio across the low pressure compressor  44  and a pressure ratio across the high pressure compressor  52 . The pressure ratio of the low pressure compressor  44  is measured as the pressure at the exit of the low pressure compressor  44  divided by the pressure at the inlet of the low pressure compressor  44 . In examples, a product of the pressure ratio of the low pressure compressor  44  and the fan pressure ratio is between 3.0 and 6.0, or more narrowly is between 4.0 and 5.5. The pressure ratio of the high pressure compressor ratio  52  is measured as the pressure at the exit of the high pressure compressor  52  divided by the pressure at the inlet of the high pressure compressor  52 . In examples, the pressure ratio of the high pressure compressor  52  is between 7.0 and 12.0, or more narrowly is between 10.0 and 11.5. The OPR can be equal to or greater than 44.0, and can be less than or equal to 70.0, such as between 50.0 and 60.0. 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  20  as well as three-spool engine architectures. 
     The engine  20  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  28  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  28 , and MTO is measured at maximum thrust of the engine  20  at static sea-level and 86 degrees Fahrenheit (° F.). The TET may be greater than or equal to 2700.0° F., or more narrowly less than or equal to 3500.0° F., such as between 2750.0° F. and 3350.0° F. The relatively high TET can be utilized in combination with the other techniques disclosed herein to provide a compact turbine arrangement. 
     The engine  20  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  28  at the MTO condition. The EGT may be less than or equal to 1000.0° F., or more narrowly greater than or equal to 800.0° F., such as between 900.0° F. and 975.0° F. The relatively low EGT can be utilized in combination with the other techniques disclosed herein to reduce fuel consumption. 
       FIG.  2 A  shows an engine  100 , which may be similar to the engine  20  of  FIG.  1   . A shaft  102  is driven by a fan drive turbine to drive a fan rotor  106  through a gear reduction  104 . The drive connection here may be generally as described above with regard to  FIG.  1   . A fan case  108  surrounds the fan rotor  106 , and an core engine  400  may include a splitter wall  110  that surrounds compressor housing wall  111  which houses a low pressure compressor  112 , and a high pressure compressor  230 , and combustor and turbine sections (not shown in this Figure). The core engine  400  must be rigidly connected to the fan case  108 , 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  100 , as will be described below, there are fewer structural fan exit guide vanes  114 . Applicant has developed a new type frame connection it calls a triangular-frame. Triangular-frames  116  have been added to engine  100  to provide additional rigidity. 
     As shown, the low pressure compressor  112  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  112 , mounting challenges are raised. 
     As can been seen from  FIGS.  2 A and  2 B , the fan exit guide vanes  114 / 129  extend from an inner point  200  attached at an angle to a radially outer point  201 , with the angle having a component in a radially outer direction, and another component in an axially aft direction. 
     Conversely, the triangular-frames  116  extend from a radially inner point  202  attached to the core engine  400 , radially outwardly at an angle to an outer point  203  connected to the fan case  108 . The angle of the triangular-frame  116  has a component in a radially outer direction and another component in an axially forward direction. 
     As shown in  FIG.  2 A , the core engine  400  includes a fan intermediate case  210  including a plurality of struts, and having a mount bracket  220 . A non-structural guide vane  129  is attached to the bracket  220  through pins  222 . In this manner, the non-structural guide vanes  129  can “float” or adjust radially relative to the bracket  220 , but are prevented from moving circumferentially. In other embodiments, the floating connection may be between vanes  129  and fan case  108 , with vanes  129  rigidly connected to the core engine. Pins  600  are shown to support that embodiment. Alternatively, the non-structural guide vanes may be fixed to bracket  220  ado to the fan case  108 . 
       FIG.  2 B  shows a detail of the mount of a structural guide vane  114  to the fan intermediate case  210  and to the bracket  220  through a first pin  300  preventing circumferential movement, and a second pin  302  preventing radial movement. 
     Returning to  FIG.  2 A , it can be seen that the triangular-frames  116  are attached at inner ends  202  at a bracket  234  which is fixed with a compressor intermediate case  232  having a plurality of struts. This view illustrates one strut. The compressor intermediate cases  232  is intermediate the low pressure compressor  112  and a high pressure compressor  230 . 
     Although specific mount locations are shown, other connections between the fan exit guide vanes  114 / 129  and triangular-frames  116  to the core engine  400  may be utilized. For purposes of this application, the core engine is defined to include at least the compressor housing wall  111 , the fan intermediate case  210 , the compressor intermediate case  232 , the low pressure compressor  112 , the high pressure compressor  230  and a combustor end turbine section, not shown, but which may be as disclosed with regard to  FIG.  1   . 
     Engine  100  is illustrated in  FIG.  3 A  having three sets of triangular frame connecting members  116 . Each triangular frame connecting member  116  includes a pair of legs  118  and  120 . The legs in each pair extend along directions such that each leg extends with a component in a radially inward direction, and in opposed circumferential directions. As can be seen, the three triangular frame connecting members together create an overall triangle. An angle A defined between each pair of legs  118  and  120  is between 55 and 65 degrees, and an embodiment 60 degrees. Each leg  120  and  118  is attached to a fan case  108  at connection  222 . One of the triangular frame connecting member  116  is attached to the fan case  108  at a top dead center  223  rigid connection  222 . The other two triangular frame connecting members  116  are attached to the fan case  108  at points  122 , on opposed circumferential sides of bottom dead center  224 . Each of the legs  118  and  120  are attached to the core engine  232  at brackets  234 . The brackets  234  are fixed to the compressor intermediate case  232  in one embodiment. As shown, there is an outer diameter d 1  to compressor intermediate case  232  at the axial location of the brackets  234 . There is also a lateral distance d 2  between the connection of the legs  118  and  120  in each triangular frame connecting member pair  116  at the location where the legs are attached to brackets  234 . Distance d 2  is greater than distance d 1  as illustrated. 
       FIG.  3 B  shows another alternative  150 . Again, there are three triangular frame connecting members  118  and  120 . However, in the embodiment of  FIG.  3    there is a connection  242  between one pair of legs  118  and  120  and a fan case  108  which is at bottom dead center  224 . The other two pairs are connected to the fan case at points  152  which are spaced circumferentially from top dead center  223 . In addition, in the embodiment  150 , there are two load links with one each extending from a leg  120  to be rigidly connected to one of the legs  120 , and extending radially outward to be rigidly connected to the fan case  108  at connection  222 . 
     In general, the triangular frame connecting members provide less obstruction to airflow than many other connecting member arrangements. 
     As shown in  FIG.  4   , there are forty-eight fan exit guide vanes total, with sixteen of the fan exit guide vanes being structural vanes  114 . Of course, other numbers may be used. Intermediate each structural fan exit guide vane  114  are two non-structural guide vanes  129 . The non-structural guide vanes  129  are connected to the fan case  108  at  130 , and to the wall  111  at a radially inner point  132 . Both connections may be rigidly connected or one may float as mentioned above. 
     Since the guide vanes  129  are not structural, they can provide an acoustic function. As shown in  FIG.  5   , one of the non-structural fan exit guide vanes  129  is illustrated. The guide vane has a pressure wall  134  and a suction wall  136 . An outer skin  137  is perforated at  138  and provided over a plurality of chambers  140 . The chambers  140  can have any number of shapes including honeycomb, or other cross-sections. The perforations  138  could be circular, but can be other shapes, including elongated slots. 
     In embodiments the structural guide vanes  114  may include 0% to 55% of the total fan exit guide vanes. In other embodiments the structural fan exit guide vanes  114  may be 15 to 55% of the total fan exit guide vanes. In other embodiments the structural fan exit guide vanes  114  may include 15 to 50% of the total fan exit guide vanes. In yet another embodiment, the structural fan exit guide vanes provide 15 to 30% of the total fan exit guide vanes. 
     The 0% (and other lower percentages) might be indicated when the engine is core mounted to the aircraft. The higher percentages may be indicated when the engine is fan mounted. 
     Although embodiments of this disclosure have been shown, a worker of ordinary skill in this art would recognize that modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the true scope and content of this disclosure.