Patent Publication Number: US-11391217-B2

Title: Stiffening member for epicyclical gear system housing assembly

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application is related to concurrently filed and co-pending U.S. patent application Ser. No. 16/592,492 entitled “EPICYCLICAL GEAR SYSTEM HOUSING ASSEMBLY,” U.S. patent application Ser. No. 16/592,494 entitled “BEARING SPRING FOR EPICYCLICAL GEAR SYSTEM HOUSING ASSEMBLY,” and U.S. patent application Ser. No. 16/592,499 entitled “STATIC CURVIC JOINT FOR EPICYCLICAL GEAR SYSTEM HOUSING ASSEMBLY,” the entirety of each of which are herein incorporated by reference. 
     BACKGROUND 
     Epicyclical gear systems may be used in rotating machinery to transfer energy from one component, such as a rotatable shaft, to another. By altering certain variables such as the number, size, and teeth of the gears, an epicyclical gear system may be designed to transfer energy between components at a desired ratio and often convert a high-speed, low-torque input to a lower-speed, higher-torque output. 
     Epicyclical gear systems may be suitable for a wide range of applications, including the transfer of energy from a turbine shaft to a fan rotor in a geared turbofan engine. However, in such dynamic applications the epicyclical gear system must be designed to allow some degree of relative movement between parts of the system to avoid excessive wear and, in extreme conditions, system failure. 
     SUMMARY 
     According to some aspects of the present disclosure, a planet gear housing assembly is disclosed for an epicyclical gear system of a gas turbine engine having an engine casing. The planet gear housing assembly comprises an aft planet carrier assembly, a forward planet carrier assembly, and a plurality of planet gears. The aft planet carrier assembly comprises an aft flange defining a central aperture and a plurality of gear shaft pockets positioned about the circumference and radially outward of the central aperture. Each pocket has a cylindrical wall, and the aft flange is coupled to the engine casing to define an aft torsional stiffness. The forward planet carrier assembly comprises a forward flange defining a central aperture and a plurality of gear shaft pockets positioned about the circumference and radially outward of the central aperture. Each pocket has a cylindrical wall, and the forward flange is coupled to the aft planet carrier assembly to define a forward torsional stiffness. The plurality of planet gears each comprise a cylindrical shaft having a forward end portion disposed in a gear shaft pocket of the forward planet carrier assembly coaxially with the cylindrical wall of the pocket, an aft end portion disposed within a gear shaft pocket of the aft planet carrier assembly coaxially with the cylindrical wall of the pocket, and one or more gears carried by the shaft between the forward and aft end portions. The forward torsional stiffness is between 60% and 80% of the aft torsional stiffness. 
     In some embodiments the forward torsional stiffness is between 65% and 75% of the aft torsional stiffness. In some embodiments the aft planet carrier assembly further comprises an annular mounting flange extending from the aft flange, the annular mounting flange positioned forward of and coaxial with the central aperture, the mounting flange forming a forward facing mounting surface. In some embodiments the forward planet carrier assembly further comprises an annular mounting flange extending from the forward flange, the annular mounting flange positioned aft of and coaxial with the central aperture, the mounting flange forming an aft facing mounting surface. In some embodiments the mounting surfaces are positioned relative to each other to thereby couple the aft planet carrier assembly and the forward planet carrier assembly. 
     In some embodiments the shaft and gears form a compound star gear in an epicyclical gear system. In some embodiments the forward planet carrier assembly further comprises a stiffening member positioned between the aft planet carrier assembly and the forward flange, the stiffening member comprising: an annular body defining a central aperture; and a plurality of radial flanges extending radially outward from the annular body, each of the plurality of radial flanges partly defining a gear-facing surface. In some embodiments the stiffening member abuts both the forward flange and the aft planet carrier assembly. 
     According to further aspects of the present disclosure, a planet gear housing assembly in an epicyclical gear assembly comprises an aft planet carrier assembly, a forward planet carrier assembly, a plurality of planet gears, and a stiffening member. The aft planet carrier assembly comprises an aft flange defining a central aperture and a plurality of gear shaft pockets positioned about the circumference and radially outward of the central aperture, each pocket having a cylindrical wall. The forward planet carrier assembly comprises a forward flange defining a central aperture and a plurality of gear shaft pockets positioned about the circumference and radially outward of the central aperture, each pocket having a cylindrical wall. The plurality of planet gears each comprise a cylindrical shaft having a forward end portion disposed in a gear shaft pocket of the forward planet carrier assembly coaxially with the cylindrical wall of the pocket, an aft end portion disposed within a gear shaft pocket of the aft planet carrier assembly coaxially with the cylindrical wall of the pocket, and one or more gears carried by the shaft between the forward and aft end portions. The stiffening member is positioned between the aft planet carrier assembly and the forward planet carrier assembly. The stiffening member comprises an annular body defining a central aperture and a plurality of radial flanges extending radially outward from the annular body, each of the plurality of radial flanges partly defining a gear-facing surface. 
     In some embodiments the aft planet carrier assembly is coupled to the forward planet carrier assembly. In some embodiments the aft planet carrier assembly further comprises an annular mounting flange extending from the aft flange, the annular mounting flange positioned forward of and coaxial with the central aperture, the mounting flange forming a forward facing mounting surface. In some embodiments the forward planet carrier assembly further comprises an annular mounting flange extending from the forward flange, the annular mounting flange positioned aft of and coaxial with the central aperture, the mounting flange forming an aft facing mounting surface. In some embodiments the mounting surfaces are positioned relative to each other to thereby couple the aft planet carrier assembly and the forward planet carrier assembly. 
     In some embodiments the shaft and gears form a compound star gear in an epicyclical gear system. In some embodiments the epicyclical gear assembly is a portion of a gas turbine engine having an engine casing, and wherein the aft flange is coupled to the engine casing to define an aft torsional stiffness and the forward flange is coupled to the aft planet carrier assembly to define a forward torsional stiffness, and wherein the forward torsional stiffness is between 50% and 90% of the aft torsional stiffness. In some embodiments the forward torsional stiffness is between 60% and 80% of the aft torsional stiffness. In some embodiments the forward torsional stiffness is between 65% and 75% of the aft torsional stiffness. In some embodiments the stiffening member comprises seven gear-facing surfaces. In some embodiments the aft flange comprises a radially outer mounting surface for mounting the planet gear housing assembly to the engine casing. 
     According to still further aspects of the present disclosure, a gas turbine engine for an aircraft comprises an engine core, a fan, and a gearbox. The engine core comprises a turbine, a compressor, and a core shaft connecting the turbine to the compressor. The fan is located upstream of the engine core and comprises a plurality of fan blades. The gearbox receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The gearbox has a planet gear housing assembly comprising an aft planet carrier assembly, a forward planet carrier assembly, a plurality of planet gears, and a stiffening member. The aft planet carrier assembly comprises an aft flange defining a central aperture and a plurality of gear shaft pockets positioned about the circumference and radially outward of the central aperture, each pocket having a cylindrical wall. The forward planet carrier assembly comprises a forward flange defining a central aperture and a plurality of gear shaft pockets positioned about the circumference and radially outward of the central aperture, each pocket having a cylindrical wall. The plurality of planet gears each comprise a cylindrical shaft having a forward end portion disposed in a gear shaft pocket of the forward planet carrier assembly coaxially with the cylindrical wall of the pocket, an aft end portion disposed within a gear shaft pocket of the aft planet carrier assembly coaxially with the cylindrical wall of the pocket, and one or more gears carried by the shaft between the forward and aft end portions. The stiffening member is positioned between and abutting the aft planet carrier assembly and the forward planet carrier assembly. The stiffening member comprises an annular body defining a central aperture and a plurality of radial flanges extending radially outward from the annular body, each of the plurality of radial flanges partly defining a gear-facing surface. 
     In some embodiments each of the plurality of planet gears comprise a sun gear engaging gear and a ring gear engaging gear carried by the shaft between the forward and aft end portions. In some embodiments the shaft and gears form a compound star gear in an epicyclical gear system. In some embodiments the gearbox further comprises a roller element bearing disposed over at least a portion of one or both of the forward end portion and the aft end portion of the gear shaft. In some embodiments the aft planet carrier assembly defines an aft torsional stiffness and the forward planet carrier assembly defines a forward torsional stiffness, and wherein the forward torsional stiffness is between 65% and 75% of the aft torsional stiffness. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The following will be apparent from elements of the figures, which are provided for illustrative purposes. 
         FIG. 1  is a sectional side view of a gas turbine engine. 
         FIG. 2  is a close up sectional side view of an upstream portion of a gas turbine engine. 
         FIG. 3  is a partially cut-away view of a gearbox for a gas turbine engine. 
         FIG. 4  is a schematic and cross sectional view of an epicyclical gear system in accordance with some embodiments of the present disclosure. 
         FIG. 5  is a detailed schematic and cross sectional view of a planet gear disposed in an epicyclical gear system in accordance with some embodiments of the present disclosure. 
         FIG. 6  is an isometric view of a forward planet carrier assembly, a stiffening member, and an aft carrier housing assembly of a planet gear housing assembly in accordance with some embodiments of the present disclosure. 
         FIG. 7  is a partial cross sectional view of a stiffening member coupled between a forward planet carrier assembly and an aft planet carrier assembly of a planet gear housing assembly in accordance with some embodiments. 
         FIG. 8A  is an isometric view of a stiffening member in accordance with some embodiments. 
         FIG. 8B  is an isometric view of a stiffening member in accordance with some embodiments. 
         FIG. 9  is a flow diagram of a method in accordance with some embodiments of the present disclosure. 
     
    
    
     The present application discloses illustrative (i.e., example) embodiments. The claimed inventions are not limited to the illustrative embodiments. Therefore, many implementations of the claims will be different than the illustrative embodiments. Various modifications can be made to the claimed inventions without departing from the spirit and scope of the disclosure. The claims are intended to cover implementations with such modifications. 
     DETAILED DESCRIPTION 
     For the purposes of promoting an understanding of the principles of the disclosure, reference will now be made to a number of illustrative embodiments in the drawings and specific language will be used to describe the same. 
     As noted elsewhere herein, the present disclosure may relate to a gas turbine engine. Such a gas turbine engine may comprise an engine core comprising a turbine, a combustor, a compressor, and a core shaft connecting the turbine to the compressor. Such a gas turbine engine may comprise a fan (having fan blades) located upstream of the engine core. 
     Arrangements of the present disclosure may be particularly, although not exclusively, beneficial for fans that are driven via a gearbox. Accordingly, the gas turbine engine may comprise a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft. The input to the gearbox may be directly from the core shaft, or indirectly from the core shaft, for example via a spur shaft and/or gear. The core shaft may rigidly connect the turbine and the compressor, such that the turbine and compressor rotate at the same speed (with the fan rotating at a lower speed). 
     The gas turbine engine as described and/or claimed herein may have any suitable general architecture. For example, the gas turbine engine may have any desired number of shafts that connect turbines and compressors, for example one, two or three shafts. Purely by way of example, the turbine connected to the core shaft may be a first turbine, the compressor connected to the core shaft may be a first compressor, and the core shaft may be a first core shaft. The engine core may further comprise a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor. The second turbine, second compressor, and second core shaft may be arranged to rotate at a higher rotational speed than the first core shaft. 
     In such an arrangement, the second compressor may be positioned axially downstream of the first compressor. The second compressor may be arranged to receive (for example directly receive, for example via a generally annular duct) flow from the first compressor. 
     The gearbox may be arranged to be driven by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example the first core shaft in the example above). For example, the gearbox may be arranged to be driven only by the core shaft that is configured to rotate (for example in use) at the lowest rotational speed (for example only be the first core shaft, and not the second core shaft, in the example above). Alternatively, the gearbox may be arranged to be driven by any one or more shafts, for example the first and/or second shafts in the example above. 
     The gearbox may be a reduction gearbox (in that the output to the fan is a lower rotational rate than the input from the core shaft). Any type of gearbox may be used. For example, the gearbox may be a “planetary” or “star” gearbox, as described in more detail elsewhere herein. The gearbox may have any desired reduction ratio (defined as the rotational speed of the input shaft divided by the rotational speed of the output shaft), for example greater than 2.5, for example in the range of from 3 to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, for example, between any two of the values in the previous sentence. Purely by way of example, the gearbox may be a “star” gearbox having a ratio in the range of from 3.1 or 3.2 to 3.8. In some arrangements, the gear ratio may be outside these ranges. 
     In any gas turbine engine as described and/or claimed herein, a combustor may be provided axially downstream of the fan and compressor(s). For example, the combustor may be directly downstream of (for example at the exit of) the second compressor, where a second compressor is provided. By way of further example, the flow at the exit to the combustor may be provided to the inlet of the second turbine, where a second turbine is provided. The combustor may be provided upstream of the turbine(s). 
     The or each compressor (for example the first compressor and second compressor as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes, which may be variable stator vanes (in that their angle of incidence may be variable). The row of rotor blades and the row of stator vanes may be axially offset from each other. 
     The or each turbine (for example the first turbine and second turbine as described above) may comprise any number of stages, for example multiple stages. Each stage may comprise a row of rotor blades and a row of stator vanes. The row of rotor blades and the row of stator vanes may be axially offset from each other. 
     Each fan blade may be defined as having a radial span extending from a root (or hub) at a radially inner gas-washed location, or 0% span position, to a tip at a 100% span position. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36, 0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. The ratio of the radius of the fan blade at the hub to the radius of the fan blade at the tip may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.32. These ratios may commonly be referred to as the hub-to-tip ratio. The radius at the hub and the radius at the tip may both be measured at the leading edge (or axially forwardmost) part of the blade. The hub-to-tip ratio refers, of course, to the gas-washed portion of the fan blade, i.e. the portion radially outside any platform. 
     The radius of the fan may be measured between the engine centreline and the tip of a fan blade at its leading edge. The fan diameter (which may simply be twice the radius of the fan) may be greater than (or on the order of) any of: 220 cm, 230 cm, 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches), 380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm (around 165 inches). The fan diameter may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 240 cm to 280 cm or 330 cm to 380 cm. 
     The rotational speed of the fan may vary in use. Generally, the rotational speed is lower for fans with a higher diameter. Purely by way of non-limitative example, the rotational speed of the fan at cruise conditions may be less than 2500 rpm, for example less than 2300 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cm to 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for example in the range of from 1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100 rpm. Purely by way of further non-limitative example, the rotational speed of the fan at cruise conditions for an engine having a fan diameter in the range of from 330 cm to 380 cm may be in the range of from 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpm to 1800 rpm. 
     In use of the gas turbine engine, the fan (with associated fan blades) rotates about a rotational axis. This rotation results in the tip of the fan blade moving with a velocity Utip. The work done by the fan blades  13  on the flow results in an enthalpy rise dH of the flow. A fan tip loading may be defined as dH/Utip2, where dH is the enthalpy rise (for example the 1-D average enthalpy rise) across the fan and Utip is the (translational) velocity of the fan tip, for example at the leading edge of the tip (which may be defined as fan tip radius at leading edge multiplied by angular speed). The fan tip loading at cruise conditions may be greater than (or on the order of) any of: 0.28, 0.29, 0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in this paragraph being Jkg−1K−1/(ms−1)2). The fan tip loading may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 0.28 to 0.31, or 0.29 to 0.3. 
     Gas turbine engines in accordance with the present disclosure may have any desired bypass ratio, where the bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core at cruise conditions. In some arrangements the bypass ratio may be greater than (or on the order of) any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5, 15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of form 12 to 16, 13 to 15, or 13 to 14. The bypass duct may be substantially annular. The bypass duct may be radially outside the core engine. The radially outer surface of the bypass duct may be defined by a nacelle and/or a fan case. 
     The overall pressure ratio of a gas turbine engine as described and/or claimed herein may be defined as the ratio of the stagnation pressure upstream of the fan to the stagnation pressure at the exit of the highest pressure compressor (before entry into the combustor). By way of non-limitative example, the overall pressure ratio of a gas turbine engine as described and/or claimed herein at cruise may be greater than (or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65, 70, 75. The overall pressure ratio may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 50 to 70. 
     Specific thrust of an engine may be defined as the net thrust of the engine divided by the total mass flow through the engine. At cruise conditions, the specific thrust of an engine described and/or claimed herein may be less than (or on the order of) any of the following: 110 Nkg−1s, 105 Nkg−1s, 100 Nkg−1s, 95 Nkg−1s, 90 Nkg−1s, 85 Nkg−1s or 80 Nkg−1s. The specific thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 80 Nkg−1s to 100 Nkg−1 s, or 85 Nkg−1 s to 95 Nkg−1 s. Such engines may be particularly efficient in comparison with conventional gas turbine engines. 
     A gas turbine engine as described and/or claimed herein may have any desired maximum thrust. Purely by way of non-limitative example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust of at least (or on the order of) any of the following: 160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN, 450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). Purely by way of example, a gas turbine as described and/or claimed herein may be capable of producing a maximum thrust in the range of from 330 kN to 420 kN, for example 350 kN to 400 kN. The thrust referred to above may be the maximum net thrust at standard atmospheric conditions at sea level plus 15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.), with the engine static. 
     In use, the temperature of the flow at the entry to the high pressure turbine may be particularly high. This temperature, which may be referred to as TET, may be measured at the exit to the combustor, for example immediately upstream of the first turbine vane, which itself may be referred to as a nozzle guide vane. At cruise, the TET may be at least (or on the order of) any of the following: 1400K, 1450K, 1500K, 1550K, 1600K or 1650K. The TET at cruise may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds). The maximum TET in use of the engine may be, for example, at least (or on the order of) any of the following: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. The maximum TET may be in an inclusive range bounded by any two of the values in the previous sentence (i.e. the values may form upper or lower bounds), for example in the range of from 1800K to 1950K. The maximum TET may occur, for example, at a high thrust condition, for example at a maximum take-off (MTO) condition. 
     A fan blade and/or aerofoil portion of a fan blade described and/or claimed herein may be manufactured from any suitable material or combination of materials. For example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a composite, for example a metal matrix composite and/or an organic matrix composite, such as carbon fibre. By way of further example at least a part of the fan blade and/or aerofoil may be manufactured at least in part from a metal, such as a titanium based metal or an aluminium based material (such as an aluminium-lithium alloy) or a steel based material. The fan blade may comprise at least two regions manufactured using different materials. For example, the fan blade may have a protective leading edge, which may be manufactured using a material that is better able to resist impact (for example from birds, ice or other material) than the rest of the blade. Such a leading edge may, for example, be manufactured using titanium or a titanium-based alloy. Thus, purely by way of example, the fan blade may have a carbon-fibre or aluminium based body (such as an aluminium lithium alloy) with a titanium leading edge. 
     A fan as described and/or claimed herein may comprise a central portion, from which the fan blades may extend, for example in a radial direction. The fan blades may be attached to the central portion in any desired manner. For example, each fan blade may comprise a fixture which may engage a corresponding slot in the hub (or disc). Purely by way of example, such a fixture may be in the form of a dovetail that may slot into and/or engage a corresponding slot in the hub/disc in order to fix the fan blade to the hub/disc. By way of further example, the fan blades maybe formed integrally with a central portion. Such an arrangement may be referred to as a bladed disc or a bladed ring. Any suitable method may be used to manufacture such a bladed disc or bladed ring. For example, at least a part of the fan blades may be machined from a block and/or at least part of the fan blades may be attached to the hub/disc by welding, such as linear friction welding. 
     The gas turbine engines described and/or claimed herein may or may not be provided with a variable area nozzle (VAN). Such a variable area nozzle may allow the exit area of the bypass duct to be varied in use. The general principles of the present disclosure may apply to engines with or without a VAN. 
     The fan of a gas turbine as described and/or claimed herein may have any desired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26 fan blades. 
     As used herein, cruise conditions have the conventional meaning and would be readily understood by the skilled person. Thus, for a given gas turbine engine for an aircraft, the skilled person would immediately recognise cruise conditions to mean the operating point of the engine at mid-cruise of a given mission (which may be referred to in the industry as the “economic mission”) of an aircraft to which the gas turbine engine is designed to be attached. In this regard, mid-cruise is the point in an aircraft flight cycle at which 50% of the total fuel that is burned between top of climb and start of descent has been burned (which may be approximated by the midpoint—in terms of time and/or distance—between top of climb and start of descent. Cruise conditions thus define an operating point of the gas turbine engine that provides a thrust that would ensure steady state operation (i.e. maintaining a constant altitude and constant Mach Number) at mid-cruise of an aircraft to which it is designed to be attached, taking into account the number of engines provided to that aircraft. For example where an engine is designed to be attached to an aircraft that has two engines of the same type, at cruise conditions the engine provides half of the total thrust that would be required for steady state operation of that aircraft at mid-cruise. 
     In other words, for a given gas turbine engine for an aircraft, cruise conditions are defined as the operating point of the engine that provides a specified thrust (required to provide—in combination with any other engines on the aircraft—steady state operation of the aircraft to which it is designed to be attached at a given mid-cruise Mach Number) at the mid-cruise atmospheric conditions (defined by the International Standard Atmosphere according to ISO 2533 at the mid-cruise altitude). For any given gas turbine engine for an aircraft, the mid-cruise thrust, atmospheric conditions and Mach Number are known, and thus the operating point of the engine at cruise conditions is clearly defined. 
     Purely by way of example, the forward speed at the cruise condition may be any point in the range of from Mach 0.7 to 0.9, for example 0.75 to 0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example 0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach 0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Any single speed within these ranges may be part of the cruise condition. For some aircraft, the cruise conditions may be outside these ranges, for example below Mach 0.7 or above Mach 0.9. 
     Purely by way of example, the cruise conditions may correspond to standard atmospheric conditions (according to the International Standard Atmosphere, ISA) at an altitude that is in the range of from 10000 m to 15000 m, for example in the range of from 10000 m to 12000 m, for example in the range of from 10400 m to 11600 m (around 38000 ft), for example in the range of from 10500 m to 11500 m, for example in the range of from 10600 m to 11400 m, for example in the range of from 10700 m (around 35000 ft) to 11300 m, for example in the range of from 10800 m to 11200 m, for example in the range of from 10900 m to 11100 m, for example on the order of 11000 m. The cruise conditions may correspond to standard atmospheric conditions at any given altitude in these ranges. 
     Purely by way of example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 30 kN to 35 kN) at a forward Mach number of 0.8 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 38000 ft (11582 m). Purely by way of further example, the cruise conditions may correspond to an operating point of the engine that provides a known required thrust level (for example a value in the range of from 50 kN to 65 kN) at a forward Mach number of 0.85 and standard atmospheric conditions (according to the International Standard Atmosphere) at an altitude of 35000 ft (10668 m). 
     In use, a gas turbine engine described and/or claimed herein may operate at the cruise conditions defined elsewhere herein. Such cruise conditions may be determined by the cruise conditions (for example the mid-cruise conditions) of an aircraft to which at least one (for example 2 or 4) gas turbine engine may be mounted in order to provide propulsive thrust. 
     According to an aspect, there is provided an aircraft comprising a gas turbine engine as described and/or claimed herein. The aircraft according to this aspect is the aircraft for which the gas turbine engine has been designed to be attached. Accordingly, the cruise conditions according to this aspect correspond to the mid-cruise of the aircraft, as defined elsewhere herein. 
     According to an aspect, there is provided a method of operating a gas turbine engine as described and/or claimed herein. The operation may be at the cruise conditions as defined elsewhere herein (for example in terms of the thrust, atmospheric conditions and Mach Number). 
     According to an aspect, there is provided a method of operating an aircraft comprising a gas turbine engine as described and/or claimed herein. The operation according to this aspect may include (or may be) operation at the mid-cruise of the aircraft, as defined elsewhere herein. 
       FIG. 1  illustrates a gas turbine engine  10  having a principal rotational axis  9 . The engine  10  comprises an air intake  12  and a propulsive fan  23  that generates two airflows: a core airflow A and a bypass airflow B. The gas turbine engine  10  comprises a core  11  that receives the core airflow A. The engine core  11  comprises, in axial flow series, a low pressure compressor  14 , a high-pressure compressor  15 , combustion equipment  16 , a high-pressure turbine  17 , a low pressure turbine  19  and a core exhaust nozzle  20 . A nacelle  21  surrounds the gas turbine engine  10  and defines a bypass duct  22  and a bypass exhaust nozzle  18 . The bypass airflow B flows through the bypass duct  22 . The fan  23  is attached to and driven by the low pressure turbine  19  via a shaft  26  and an epicyclic gearbox  30 . 
     In use, the core airflow A is accelerated and compressed by the low pressure compressor  14  and directed into the high pressure compressor  15  where further compression takes place. The compressed air exhausted from the high pressure compressor  15  is directed into the combustion equipment  16  where it is mixed with fuel and the mixture is combusted. The resultant hot combustion products then expand through, and thereby drive, the high pressure and low pressure turbines  17 ,  19  before being exhausted through the nozzle  20  to provide some propulsive thrust. The high pressure turbine  17  drives the high pressure compressor  15  by a suitable interconnecting shaft  27 . The fan  23  generally provides the majority of the propulsive thrust. The epicyclic gearbox  30  is a reduction gearbox. 
     An exemplary arrangement for a geared fan gas turbine engine  10  is shown in  FIG. 2 . The low pressure turbine  19  (see  FIG. 1 ) drives the shaft  26 , which is coupled to a sun wheel, or sun gear,  28  of the epicyclic gear arrangement  30 . Radially outwardly of the sun gear  28  and intermeshing therewith is a plurality of planet gears  32  that are coupled together by a planet carrier  34 . The planet carrier  34  constrains the planet gears  32  to precess around the sun gear  28  in synchronicity whilst enabling each planet gear  32  to rotate about its own axis. The planet carrier  34  is coupled via linkages  36  to the fan  23  in order to drive its rotation about the engine axis  9 . Radially outwardly of the planet gears  32  and intermeshing therewith is an annulus or ring gear  38  that is coupled, via linkages  40 , to a stationary supporting structure  24 . 
     Note that the terms “low pressure turbine” and “low pressure compressor” as used herein may be taken to mean the lowest pressure turbine stages and lowest pressure compressor stages (i.e. not including the fan  23 ) respectively and/or the turbine and compressor stages that are connected together by the interconnecting shaft  26  with the lowest rotational speed in the engine (i.e. not including the gearbox output shaft that drives the fan  23 ). In some literature, the “low pressure turbine” and “low pressure compressor” referred to herein may alternatively be known as the “intermediate pressure turbine” and “intermediate pressure compressor”. Where such alternative nomenclature is used, the fan  23  may be referred to as a first, or lowest pressure, compression stage. 
     The epicyclic gearbox  30  is shown by way of example in greater detail in  FIG. 3 . Each of the sun gear  28 , planet gears  32  and ring gear  38  comprise teeth about their periphery to intermesh with the other gears. However, for clarity only exemplary portions of the teeth are illustrated in  FIG. 3 . There are four planet gears  32  illustrated, although it will be apparent to the skilled reader that more or fewer planet gears  32  may be provided within the scope of the claimed invention. Practical applications of a planetary epicyclic gearbox  30  generally comprise at least three planet gears  32 . 
     The epicyclic gearbox  30  illustrated by way of example in  FIGS. 2 and 3  is of the planetary type, in that the planet carrier  34  is coupled to an output shaft via linkages  36 , with the ring gear  38  fixed. However, any other suitable type of epicyclic gearbox  30  may be used. By way of further example, the epicyclic gearbox  30  may be a star arrangement, in which the planet carrier  34  is held fixed, with the ring (or annulus) gear  38  allowed to rotate. In such an arrangement the fan  23  is driven by the ring gear  38 . By way of further alternative example, the gearbox  30  may be a differential gearbox in which the ring gear  38  and the planet carrier  34  are both allowed to rotate. 
     It will be appreciated that the arrangement shown in  FIGS. 2 and 3  is by way of example only, and various alternatives are within the scope of the present disclosure. Purely by way of example, any suitable arrangement may be used for locating the gearbox  30  in the engine  10  and/or for connecting the gearbox  30  to the engine  10 . By way of further example, the connections (such as the linkages  36 ,  40  in the  FIG. 2  example) between the gearbox  30  and other parts of the engine  10  (such as the input shaft  26 , the output shaft and the fixed structure  24 ) may have any desired degree of stiffness or flexibility. By way of further example, any suitable arrangement of the bearings between rotating and stationary parts of the engine (for example between the input and output shafts from the gearbox and the fixed structures, such as the gearbox casing) may be used, and the disclosure is not limited to the exemplary arrangement of  FIG. 2 . For example, where the gearbox  30  has a star arrangement (described above), the skilled person would readily understand that the arrangement of output and support linkages and bearing locations would typically be different to that shown by way of example in  FIG. 2 . 
     Accordingly, the present disclosure extends to a gas turbine engine having any arrangement of gearbox styles (for example star or planetary), support structures, input and output shaft arrangement, and bearing locations. 
     Optionally, the gearbox may drive additional and/or alternative components (e.g. the intermediate pressure compressor and/or a booster compressor). 
     Other gas turbine engines to which the present disclosure may be applied may have alternative configurations. For example, such engines may have an alternative number of compressors and/or turbines and/or an alternative number of interconnecting shafts. By way of further example, the gas turbine engine shown in  FIG. 1  has a split flow nozzle  18 ,  20  meaning that the flow through the bypass duct  22  has its own nozzle  18  that is separate to and radially outside the core engine nozzle  20 . However, this is not limiting, and any aspect of the present disclosure may also apply to engines in which the flow through the bypass duct  22  and the flow through the core  11  are mixed, or combined, before (or upstream of) a single nozzle, which may be referred to as a mixed flow nozzle. One or both nozzles (whether mixed or split flow) may have a fixed or variable area. 
     The geometry of the gas turbine engine  10 , and components thereof, is defined by a conventional axis system, comprising an axial direction (which is aligned with the rotational axis  9 ), a radial direction (in the bottom-to-top direction in  FIG. 1 ), and a circumferential direction (perpendicular to the page in the  FIG. 1  view). The axial, radial and circumferential directions are mutually perpendicular. 
       FIG. 4  provides a schematic view of an epicyclical gear system  100  in accordance with some embodiments of the present disclosure. The epicyclical gear system  100  may be a compound star gear system. A sun gear  101  is coupled to and driven by a first rotatable shaft  103 . The sun gear  101  is engaged with one or more planet gears  105 , such that rotation of the sun gear  101  causes rotation of the one or more planet gears  105 . The planet gears  105  may be star gears, such that the planet gears  105  rotate about an axis that is fixed relative to the axis of rotation of the sun gear  101 . 
     Each of the one or more planet gears  105  are engaged with a ring gear  107 . The ring gear  107  is coupled via a ring gear hub  108  to a second rotatable shaft  109 . Rotation of the first rotatable shaft  103  thus drives rotation of the second rotatable shaft  109  via rotation of the sun gear  101 , one or more planet gears  105 , and ring gear  107 . In some embodiments, the first rotatable shaft  103  may be a turbine shaft (i.e. high or low speed spool) of a turbine engine, and the second rotatable shaft  109  may be a fan shaft or fan rotor. 
       FIG. 5  provides a detailed and schematic view of a housing assembly  111  of an epicyclical gear system  100  in accordance with some embodiments of the present disclosure. Each of the one or more planet gears  105  may comprise a cylindrical gear shaft  123 , a sun gear engaging gear  125 , and a ring gear engaging gear  127 . The sun gear engaging gear  125  and ring gear engaging gear  127  may be carried by the cylindrical gear shaft  123 . Each of the one or more planet gears  105  is carried by the housing assembly  111 . The planet gears  105  may be a compound star gear of the epicyclical gear system  100 . 
     The housing assembly  111  may comprise a forward housing member  113  and an aft housing member  115 . In some embodiments, the housing assembly  111  further comprises an intermediate housing member  114 . One or more of the housing members  113 ,  114 ,  115  may be joined together. The forward housing member  113  and aft housing member  115  each define a plurality of gear shaft pockets  116  having a cylindrical wall  120 . The intermediate housing member  114  may define a plurality of bores  118 . 
     A cylindrical gear shaft  123  of a planet gear  105  may have a forward end portion  141  disposed within one of the plurality of gear shaft pockets  116  formed by the forward housing member  113 . The cylindrical gear shaft  123  of the same planet gear  105  may have an aft end portion  142  disposed within one of the plurality of gear shaft pockets  116  formed by the aft housing member  115 . The cylindrical gear shaft  123  may be disposed within each gear shaft pocket  116  coaxially with the cylindrical wall  120  of the gear shaft pocket  116 . The cylindrical gear shaft  123  may extend through the bore  118  defined by the intermediate housing member  114 . The sun gear engaging gear  125  and ring gear engaging gear  127  of planet gear  105  may be carried by the gear shaft  123  between the forward end portion  141  and the aft end portion  142 . 
     The housing assembly  111  may further comprise a bearing assembly  117 . The housing assembly  111  may comprise a forward bearing assembly and an aft bearing assembly. The bearing assembly  117  may comprise a bearing. For example, a forward bearing  119  may be disposed over at least a portion of the forward end portion  141  of the gear shaft  123  and an aft bearing  121  may be disposed over at least a portion of the aft end portion  142  of the gear shaft  123 . Each of the bearings  119 ,  121  may rotatably carry the planet gear  105 . Each of the bearings  119 ,  121  may be a roller element bearing. 
     During operation of the epicyclical gear system  100  the constituent pieces of the system  100  as described above and including additional carrier assemblies, housings, and bearings, may move relative to one another. Even small changes in the relative positioning of one component to another can have significant impacts on performance of the system  100 . For example, misalignment of enmeshed gear teeth and/or bearings can cause uneven gear and/or bearing loading, and degradation or damage to gear teeth. For example, misalignment of enmeshed gear teeth and/or bearings can cause uneven gear and/or bearing loading, degradation or damage to gear teeth, and reduction of bearing lives and bearing stability. 
     Of particular concern is the fore-to-aft alignment of the gear shaft  123  of each planet gear  105 . Since each planet gear  105  is carried by the gear shaft  123  partly disposed within gear shaft pockets  116  of a forward housing member  113  and an aft housing member  115 , relative movement or changes in relative positioning between the forward housing member  113  and aft housing member  115  can cause misalignment of the planet gear  105 . Similarly, relative movement or changes in relative positioning between the intermediate housing member  114  and either or both of the forward housing member  113  and aft housing member  115  can cause misalignment of the planet gear  105 . 
     This misalignment may in turn lead to uneven load sharing among the planet gears  105 , gear degradation, and shortened useful life of the planet gears  105  and/or the planet bearings  119 ,  121 . Factors that may contribute to planet gear misalignment include unaligned forces between the forward bearing  119  and aft bearing  121 , manufacturing inaccuracies for the positions of the gear shaft pockets  116  and bore  118 , ability to reassembly the forward and aft housings in the same position that they were machined, gear tolerance, and inflexibility and/or relative stiffness between each of the housing members  113 ,  114 ,  115 . 
     The present disclosure is therefore directed to systems and methods for improving and maintaining planet gear alignment in an epicyclical gear system. More specifically, the present disclosure is directed to a planet gear housing assembly for an epicyclical gear system having a forward planet carrier assembly, an aft planet carrier assembly, a stiffening member positioned between the forward and aft planet carrier assemblies, and a plurality of planet gears each carried by the forward and aft planet carrier assemblies. The stiffening member may achieve a desired stiffness ratio between the forward and aft planet carrier assemblies. 
     As shown in  FIGS. 4-8 , a planet gear housing assembly  111  may comprise an aft planet carrier assembly  131 , a forward planet carrier assembly  161 , a stiffening member  171 , and a plurality of planet gears  105  each carried by the aft planet carrier assembly  131  and forward carrier planet assembly.  FIG. 6  is an isometric view of a forward planet carrier assembly, a stiffening member, and an aft carrier housing assembly of a planet gear housing assembly in accordance with some embodiments of the present disclosure.  FIG. 7  is a partial cross sectional view of a stiffening member coupled between a forward planet carrier assembly and an aft planet carrier assembly of a planet gear housing assembly in accordance with some embodiments.  FIGS. 8A and 8B  provide isometric views of a stiffening member in accordance with some embodiments. 
     An aft planet carrier assembly  131  may comprise one or both of the intermediate housing member  114  and the aft housing member  115 . The aft planet carrier assembly  131  may comprise an aft flange  136 . The aft flange  136  may be the intermediate housing member  114 , the aft housing member  115 , or another flange member. The aft flange  136  may comprise more than one flange, as shown in  FIG. 6  with a first aft flange  136 A and a second aft flange  136 B. The aft flange  136  may define a central aperture  132  and a plurality of gear shaft pockets  137  positioned about the circumference of and radially outward of the central aperture  132 . Each of the gear shaft pockets  116  may have a cylindrical wall  138 . In some embodiments the aft flange  136  may further comprise a radially outer mounting surface  139  for coupling the aft planet carrier assembly  131  to an engine casing. 
     The aft planet carrier assembly  131  may further comprise an annular mounting flange  133 . The annular mounting flange  133  may be positioned forward of and coaxial with the central aperture  132 . The annular mounting flange  133  may extend substantially perpendicularly from the aft flange  136 . The annular mounting flange  133  may form a forward facing mounting surface  134  that may comprise a curvic structure. 
     A forward planet carrier assembly  161  may comprise a forward flange  162  and an annular mounting flange  143 . The forward planet carrier assembly  161  may be the forward housing member  113 . The forward flange  162  may define a central aperture  144  and a plurality of gear shaft pockets  145  positioned about the circumference and radially outward of the central aperture  144 . Each gear shaft pocket  145  may have a cylindrical wall  146 . The annular mounting flange  143  may be positioned aft of and coaxial with the central aperture  144 . The annular mounting flange  143  may form an aft facing mounting surface  147  that may comprise a curvic structure. The annular mounting flange  143  may extend substantially perpendicularly from the forward flange  162 . 
     A plurality of planet gears  105  may be carried by the forward planet carrier assembly  161  and the aft planet carrier assembly  131 . Each of the planet gears  105  may comprise a cylindrical shaft  123  having a forward end portion  141  disposed in a gear shaft pocket  145  of the forward flange  162 . The forward end portion  141  may be disposed in the gear shaft pocket  145  coaxial with the cylindrical wall  146  defining gear shaft pocket  145 . The cylindrical shaft may have an aft end portion  142  disposed within a gear shaft pocket  137  of the aft flange  136 , and may be disposed in the gear shaft pocket  137  coaxial with the cylindrical wall  138 . 
     Each planet gear  105  may further comprise one or more gears carried by the cylindrical shaft  123  between the forward end portion  141  and the aft end portion  142 . The gears may be, for example, a sun gear engaging portion  125  of the planet gear  105  and/or a ring gear engaging portion  127  of the planet gear  105 . 
     The planet gear housing assembly  111  and/or the forward planet carrier assembly  161  may further comprise a stiffening member  171 . The stiffening member  171  may comprise an annular body  172  defining a central aperture  173 . The stiffening member  171  may further comprise a plurality of radial flanges  174  extending radially outward from the annular body  172 . Each of the radial flanges  174  may at least partly define a gear-facing surface  175 . In some embodiments the stiffening member  171  comprises seven radial flanges  174  and defines seven gear-facing surfaces  175 . 
     The stiffening member  171  may be positioned between the aft planet carrier assembly  131  and the forward planet carrier assembly  161 . In some embodiments the stiffening member  171  may abut one or both of the forward flange  162  and the aft flange  136 . In some embodiments the stiffening member  171  is coupled to one or both of the forward flange  162  and the aft flange  136  by a plurality of bolts, pins, or other fasteners  176 . 
     When the epicyclical gear system  100  is fully assembled, a respective planet gear  105  may be carried by the forward flange  162  and the aft flange  136 , and may have a gear portion such as the ring gear engaging portion  127  positioned proximate the gear-facing surface  175  of the stiffening member  171 . The stiffening member  171  may be coupled between the forward flange  162  and the aft flange  136  to improve torsional stiffness of the forward planet carrier assembly  161  with reference to the aft planet carrier assembly  131 . 
     The epicyclical gear system  100  may be an epicyclical gear system of a gas turbine engine. The gas turbine engine may comprise an engine casing  177 , a portion of which is illustrated in cross section at  FIG. 7 . The aft flange  136  may be coupled to the engine casing  177  to define an aft torsional stiffness. The forward flange  162  may be coupled to the aft flange  136  to define a forward torsional stiffness. The forward flange  162  may be coupled to the aft flange  136 , for example, with the forward facing mounting surface  134  of the mounting flange  133  positioned relative to the aft facing mounting surface  147  of the mounting flange  143 . The forward flange  162  may be coupled to the aft flange  136  by a static curvic joint. 
     In some embodiments the forward torsional stiffness may be between 60% and 80% of the aft torsional stiffness. The aft torsional stiffness may be greater due to the mounting of the aft flange  136  to the aft casing  177  at a greater radius than the forward flange  162  is mounted to the aft flange  136 . More broadly, the forward torsional stiffness may be between 50% and 90% of the aft torsional stiffness. In other embodiments the forward torsional stiffness may be between 65% and 75% of the aft torsional stiffness. 
       FIG. 9  is a flow diagram of a method  900  of reducing relative movement between static components of an epicyclical gear system in accordance with some embodiments of the present disclosure. Method  900  starts at Block  901 . The steps of method  900 , presented at Blocks  901  through  919 , may be performed in the order presented in  FIG. 9  or in another order. One or more steps of the method  900  may not be performed. 
     At Block  903  a static forward planet carrier assembly  161  and a static aft planet carrier assembly  131  may be provided. The forward planet carrier assembly  161  may comprise a forward flange  162  defining a central aperture  144  and a plurality of gear shaft pockets  145  positioned about the circumference of the central aperture  144 . The forward planet carrier assembly  161  may further comprise an annular mounting flange  143  extending from said forward flange  162 , said annular mounting flange  143  positioned aft of and coaxial with said central aperture  144 . The mounting flange  143  may form an aft facing mounting surface  147 . 
     The aft planet carrier assembly  131  may comprise an aft flange  136  defining a central aperture  132  and a plurality of gear shaft pockets  137  positioned about the circumference and radially outward of said central aperture  132 . The aft planet carrier assembly  131  may further comprise an annular mounting flange  133  extending from said aft flange  136 , said annular mounting flange  133  positioned forward of and coaxial with said central aperture  132 . The mounting flange  133  may form a forward facing mounting surface  134 . 
     At Block  905  a stiffening member  171  may be positioned between said aft planet carrier assembly  131  and said forward planet carrier assembly  161 . The stiffening member  171  may comprise an annular body  172  defining a central aperture  173  and a plurality of radial flanges  174  extending radially outward from said annular body  172 . Each of said plurality of radial flanges  174  may partly define a gear-facing surface  175 . 
     At Block  907  a bearing  119 ,  121  may be positioned at least partly in a gear shaft pocket  145 ,  137  of the forward planet carrier assembly  161  and/or the aft planet carrier assembly  131 . 
     At Block  909  a planet gear  105  of a plurality of planet gears  105  may be positioned in each of the plurality of axially aligned gear shaft pocket pairs formed between the pocket  145 ,  137  of the forward planet carrier assembly  161  and/or the aft planet carrier assembly  131 . Each planet gear  105  may comprise a cylindrical shaft  123  having a forward end portion  141  disposed in a gear shaft pocket  145  of said forward planet carrier assembly  161  coaxially with the cylindrical wall  146  of said pocket  145 , an aft end portion  142  disposed within a gear shaft pocket  137  of said aft planet carrier assembly  131  coaxially with the cylindrical wall  138  of said pocket  137 . The planet gear  105  may further comprise one or more gears  125 ,  127  carried by said shaft  123  between said forward and aft end portions  141 ,  142 . 
     At Block  911 , a portion of a cylindrical shaft  123  of a planet gear  105  may be carried with said bearing  119 ,  121 . 
     At Block  913  the method may further comprise positioning a mounting surface  134  of the aft planet carrier assembly  131  relative to a mounting surface  147  of the forward planet carrier assembly  161 . 
     At Block  915  the forward planet carrier assembly  161  may be coupled to the aft planet carrier assembly  131 . 
     At Block  917  the cylindrical shaft  123  of each planet gear  105  may be rotated. The rotation may be driven by a sun gear  101  of a compound star gear assembly  100 . 
     Method  900  ends at Block  919 . 
     The presently disclosed systems and methods provide numerous advantages over prior art systems. By providing a stiffening member positioned between and joining a forward and aft planet carrier assembly, the disclosed planet gear housing assembly reduces fore-to-aft misalignment of planet gear shafts caused by relative movement between the forward and aft planet carrier assemblies. The aft planet carrier assembly may be rigidly mounted to the engine casing at the outer diameter of the aft flange. Since the forward flange may be mounted to the aft flange at an inner diameter, the forward-to-aft coupling of flanges may be inherently less stiff and therefore allow increased torsional deflection for an equal bearing reaction load. 
     By coupling the aft planet carrier assembly and forward planet carrier assembly with a stiffening member between them, the forward planet carrier assembly may be held to the aft planet carrier assembly with less relative movement between the two than in a standard coupling. While typical epicyclical gear systems may have a forward torsional stiffness of 25-40% of the aft torsional stiffness, the disclosed stiffening member provides for an increase to a forward torsional stiffness of 60-80% of the aft torsional stiffness. Less relative movement is advantageous during assembly as well as operation of the epicyclical gear system. 
     The present disclosure may be used in combination with the disclosure of one or more of the related applications listed above. In particular, the present disclosure may be used in combination with “Static Curvic Joint for an Epicyclical Gear System Housing Assembly.” A static curvic joint assures bore-to-bore alignment during machining and assembly but may decrease the forward torsional stiffness. The stiffening member described herein may counter this decrease in forward torsional stiffness, such that the combination of a static curvic joint and stiffening member in an epicyclical gear system is advantageous. The combination may assure bore-to-bore alignment during manufacture, machining, assembly, and operation. 
     Although examples are illustrated and described herein, embodiments are nevertheless not limited to the details shown, since various modifications and structural changes may be made therein by those of ordinary skill within the scope and range of equivalents of the claims.