The present invention relates to a gas turbine engine and, more particularly, to a case therefor.
Gas turbine engines, such as those that power modern commercial and military aircraft, generally include a compressor section to pressurize an airflow, a combustor section to burn a hydrocarbon fuel in the presence of the pressurized air, and a turbine section to extract energy from the resultant combustion gases.
The combustor section typically includes an outer shell lined with heat shields to form a combustion chamber. The combustion chamber is surrounded by a diffuser case formed of an inner and outer case, where the inner case defines a pre-diffuser and the outer case serves as structural containment. Together the cases form the flowpath and necessary volume to mitigate unrecoverable compressor surge. Although effective, the diffuser case includes multiple through-holes which may form undesirable stress concentrations.
A case for a gas turbine engine, according to one aspect of the present invention, is claimed in claim 1. The wall may be an outer wall of a diffuser case. A method of reducing stress in the case, according to another aspect of the present invention, is claimed in claim 3.
BRIEF DESCRIPTION OF THE DRAWINGS
Various features will become apparent to those skilled in the art from the following detailed description of the disclosed examples and embodiments. The drawings that accompany the detailed description can be briefly described as follows:
FIG. 1 is a schematic cross-section of an example gas turbine engine architecture;
FIG. 2 is a schematic cross-section of another example gas turbine engine architecture;
FIG. 3 is an expanded longitudinal schematic sectional view of a combustor section according to one embodiment that may be used with the example gas turbine engine architectures shown in FIGS. 1 and 2;
FIG. 4 is a schematic view of a gas turbine engine case assembly;
FIG. 5 is an expanded schematic view of a case;
FIG. 6 is an expanded outer perspective view of a through-hole in the case;
FIG. 7 is an expanded inner perspective view of a through-hole in the engine case and pockets formed therein;
FIG. 8 is an inner perspective view of a through hole with corresponding pockets defined in the case assembly according to an embodiment of the present invention;
FIG. 9 is a lateral sectional view showing the pockets of FIG. 8; and
FIG. 10 is a perspective sectional view showing example stress concentrations adjacent to the pockets of FIG. 8.
FIG. 1 schematically illustrates a gas turbine engine 20. The gas turbine engine 20 is disclosed herein as a two-spool turbo fan that generally incorporates a fan section 22, a compressor section 24, a combustor section 26 and a turbine section 28. Another alternative engine architecture 200 might include an augmentor section 12, an exhaust duct section 14 and a nozzle section 16 in addition to the fan section 22', compressor section 24', combustor section 26' and turbine section 28' (see FIG. 2). Although depicted as an aero engine in the disclosed examples, it should be understood that the concepts described herein are not so limited and the teachings may be applied to other types of turbine engines such as a turbojets, turboshafts, and three-spool (plus fan) turbofans with an intermediate spool.
Referring to FIG. 1, the fan section 22 drives air along a bypass flowpath and into the compressor section 24. The compressor section 24 drives air along a core flowpath for compression and communication into the combustor section 26, which then expands and directs the air through the turbine section 28. The engine 20 generally includes a low spool 30 and a high spool 32 mounted for rotation about an engine central longitudinal axis A relative to an engine case assembly 36 via several bearing structures 38. The low spool 30 generally includes an inner shaft 40 that interconnects a fan 42, a low pressure compressor ("LPC") 44 and a low pressure turbine ("LPT") 46. The inner shaft 40 may drive the fan 42 directly (see FIG. 2) or through a geared architecture 48 (see FIG. 1) to drive the fan 42 at a lower speed than the low spool 30. An exemplary reduction transmission is an epicyclic transmission, namely a planetary or star gear system.
The high spool 32 includes an outer shaft 50 that interconnects a high pressure compressor ("HPC") 52 and a high pressure turbine ("HPT") 54. A combustor 56 is arranged between the HPC 52 and the HPT 54. The inner shaft 40 and the outer shaft 50 are concentric and rotate about the engine central longitudinal axis A which is collinear with their longitudinal axes.
Core airflow is compressed by the LPC 44 then the HPC 52, mixed with the fuel and burned in the combustor 56, then expanded over the HPT 54 and the LPT 46. The LPT 46 and the HPT 54 rotationally drive the respective low spool 30 and the high spool 32 in response to the expansion. The main engine shafts 40, 50 are supported at a plurality of points by the bearing structures 38 within the static structure 36.
In one non-limiting example, the gas turbine engine 20 is a high-bypass geared aircraft engine with a bypass ratio greater than about six (6:1). The geared architecture 48 can include an epicyclic gear train, such as a planetary gear system or other gear system. The example epicyclic gear train has a gear reduction ratio of greater than about 2.3:1, and in another example, is greater than about 2.5:1. The geared turbofan enables operation of the low spool 30 at higher speeds which can increase the operational efficiency of the LPC 44 and the LPT 46 to render increased pressure in a fewer number of stages.
A pressure ratio associated with the LPT 46 is pressure measured prior to the inlet of the LPT 46 as related to the pressure at the outlet of the LPT 46 prior to an exhaust nozzle of the gas turbine engine 20. In another non-limiting example, the bypass ratio of the gas turbine engine 20 is greater than about ten (10:1), the fan diameter is significantly larger than that of the LPC 44, and the LPT 46 has a pressure ratio greater than about five (5:1). It should be appreciated, however, that the above parameters are only exemplary of a geared architecture engine and that the present disclosure is applicable to other gas turbine engines including direct drive turbofans. In an example high-bypass turbofan configuration, significant thrust is provided by the bypass flow path due to the high bypass ratio as the fan section 22 is designed for a particular flight condition - typically cruise at about 0.8 Mach and about 35,000 feet (10668 meters).
This flight condition, with the gas turbine engine 20 at its best fuel consumption, is also known as bucket cruise Thrust Specific Fuel Consumption (TSFC). TSFC is an industry standard parameter of fuel consumption per unit of thrust.
Fan Pressure Ratio is the pressure ratio across a fan blade of the fan section 22 without the use of a Fan Exit Guide Vane system. The low Fan Pressure Ratio according to one example gas turbine engine 20 is less than 1.45. Low Corrected Fan Tip Speed is the actual fan tip speed divided by an industry standard temperature correction of ("Tram" / 518.7)0.5
. The Low Corrected Fan Tip Speed according to the example gas turbine engine 20 is less than about 1150 fps (351 m/s).
With reference to FIG. 3, the combustor section 26 generally includes a combustor 56 with an outer combustor wall assembly 60, an inner combustor wall assembly 62 and a diffuser case 64. The outer combustor wall assembly 60 and the inner combustor wall assembly 62 are spaced apart such that an annular combustion chamber 66 is defined therebetween.
The outer combustor wall assembly 60 is spaced radially inward from an outer diffuser case 64A of the diffuser case 64 to define an outer annular plenum 76. The inner combustor wall assembly 62 is spaced radially outward from an inner diffuser case 64B of the diffuser case 64 to define an inner annular plenum 78. It should be understood that although a particular combustor is illustrated, other combustor types with various combustor wall and diffuser case arrangements will also benefit herefrom.
The combustor wall assemblies 60, 62 contain the combustion products for direction toward the turbine section 28. Each combustor wall assembly 60, 62 generally includes a respective support shell 68, 70 which supports one or more liner panels 72, 74 mounted within the respective support shell 68, 70. Each of the liner panels 72, 74 may be generally rectilinear with a circumferential arc and manufactured of, for example, a nickel based super alloy, ceramic or other temperature resistant material and are arranged to form a liner array. In one disclosed non-limiting example, the liner array includes a multiple of forward liner panels 72A and a multiple of aft liner panels 72B that are circumferentially staggered to line the outer shell 68. A multiple of forward liner panels 74A and a multiple of aft liner panels 74B are circumferentially staggered to line the inner shell 70.
The combustor 56 further includes a forward assembly 80 immediately downstream of the compressor section 24 (see FIG. 1) to receive compressed airflow therefrom. The forward assembly 80 generally includes an annular hood 82 and a bulkhead assembly 84 that support a multiple of fuel nozzles 86 (one shown) and a multiple of swirlers 90 (one shown). The annular hood 82 extends radially between, and is secured to, the forwardmost ends of the combustor wall assemblies 60, 62. The annular hood 82 includes a multiple of circumferentially distributed hood ports 94 that accommodate the respective fuel nozzle 86 and introduce air into the forward end of the combustion chamber 66 through a respective swirler 90. The bulkhead assembly 84 includes a bulkhead support shell 96 secured to the combustor wall assemblies 60, 62, and a multiple of circumferentially distributed bulkhead liner panels 98 secured to the bulkhead support shell 96. Each fuel nozzle 86 may be secured to the diffuser case 64 and project through one of the hood ports 94 and respective swirlers 90.
The forward assembly 80 introduces core combustion air into the forward section of the combustion chamber 66 while the remainder enters the outer annular plenum 76 and the inner annular plenum 78. The multiple of fuel nozzles 86 and adjacent structure generate a fuel-air mixture that supports stable combustion in the combustion chamber 66.
Opposite the forward assembly 80, the outer and inner support shells 68, 70 are mounted to a first row of Nozzle Guide Vanes (NGVs) 54A in the HPT 54 (see FIG. 1). The NGVs 54A are static engine components which direct core airflow combustion gases onto turbine blades in the turbine section 28 to facilitate the conversion of pressure energy into kinetic energy. The core airflow combustion gases are also accelerated by the NGVs 54A because of their convergent shape and are typically given a "spin" or a "swirl" in the direction of turbine rotor rotation.
With reference to FIG. 4, the engine case assembly 36 generally includes a multiple of cases or modules in addition to the outer diffuser case 64A to include, for example, a fan case 100, an intermediate case 102, a HPC case 104, the outer diffuser case 64A, a HPT case 106, a mid turbine frame (MTF) case 108, a LPT case 110, and a Turbine Exhaust Case (TEC) 112. The fan case 100 is bolted to the intermediate case 102, which is bolted to the HPC case 104, which is bolted to the outer diffuser case 64A, which is bolted to the HPT case 106, which is bolted to the MTF case 108, which is bolted to the LPT case 110, which is bolted to the TEC 112 each at a respective flange. It should be understood that the order of assembly may not necessarily follow the disclosed description and that various additional or alternative cases may be provided.
With reference to FIG. 5, the outer diffuser case 64A generally includes a multiple of through-holes 120 which penetrate through a wall 122 typical of holes configured to receive instrumentation such as a borescope, threaded holes for bolts to mount various components such as the fuel injectors and other types of apertures. The through-holes 120 are defined through a boss 124 which extends from an outer surface 126 of the wall 122 (see FIG. 6).
The outer diffuser case 64A is pressurized, which produces hoop stresses in the wall 122. At the holes 120, stresses are relatively high. The through-holes 120 create high stress concentrations in the wall material that may otherwise reduce the strength and life of the component. To reduce these stresses, an inner surface 128 of the wall 122 includes pockets 130 which are operatively disposed adjacent to and circumferentially flank each through-hole 120 (see FIG. 7). A single through-hole 120 requires two pockets 130. That is, the pockets 130 are located on either side of the through-hole 120 in the hoop direction along a hoop line H such that the pockets 130 are aligned with the stress state such as the hoop stresses to break or otherwise shield the through-hole 120 from the nominal local stresses. It will be appreciated by those skilled in the art that such pockets 130 provide space for material of wall 122 to deform, expand, and/or contract, and thus reduce stresses in the wall material defining the through-holes 120.
With reference to FIG. 7, the pockets 130 have a circular outer edge.
With reference to FIG. 8, the pockets 130 are generally larger than the through-hole 120 and are of a diameter D 100%-500% of a diameter D of the through-hole 120. The pockets 130 may be dimple shaped. The pockets extend from the inner surface 128 for a depth T of between 10%-50% a thickness 't' of the wall 122 (FIG. 9). It should be appreciated that the pockets 130 may alternatively include a flat bottom, a curved bottom, or be spherical in shape. An outer periphery 132 of each of the pockets 130 are also circumferentially spaced a distance Z from an outer periphery 134 of the through-hole 120 along the hoop line H from between 10%-100% the diameter of the through-hole 120.
With reference to FIG. 10, the pockets 130 reduce the stress in and around material disposed about an inner periphery of the through-hole 120, which increases the strength and reduces crack initiation at that location. The pockets 130, in one tested example, reduce stress by approximately 10% at the critical stress location 140, but increase stress in non-critical areas, locations 142 and 144. Relatively deeper pockets 130 drive higher stress into locations 142 and 144 to reduce stress at location 140 such that optimal pocket 130 depth T results in a desired balance of stress. The pockets 130 also advantageously reduce weight.
The use of the terms "a" and "an" and "the" and similar references in the context of description (especially in the context of the following claims) are to be construed to cover both the singular and the plural, unless otherwise indicated herein or specifically contradicted by context. The modifier "about" used in connection with a quantity is inclusive of the stated value and has the meaning dictated by the context (e.g., it includes the degree of error associated with measurement of the particular quantity). All ranges disclosed herein are inclusive of the endpoints, and the endpoints are independently combinable with each other. It should be appreciated that relative positional terms such as "forward," "aft," "upper," "lower," "above," "below," and the like are with reference to the normal operational attitude of the vehicle and should not be considered otherwise limiting.
The foregoing description is exemplary rather than defined by the features within. Various examples and embodiments are disclosed herein; however, one of ordinary skill in the art would recognize that various modifications and variations in light of the above teachings will fall within the scope of the appended claims. It is therefore to be appreciated that within the scope of the appended claims, the disclosure may be practiced other than as specifically described. For that reason the appended claims should be studied to determine true scope and content.
Gehäuse (64A) für einen Gasturbinentriebwerk (20), wobei das Gehäuse Folgendes umfasst:
eine Wand (122), die ein Durchgangsloch (120) definiert; und
erste und zweite Taschen (130) neben und auf gegenüberliegenden Seiten des Durchgangslochs (120), wobei die ersten und zweiten Taschen (130) das Durchgangsloch umlaufend flankieren und in einer Innenfläche (128) der Wand (122) definiert sind,
dadurch gekennzeichnet, dass:
mindestens eine der ersten oder zweiten Taschen (130) eine kreisförmige Peripherie einschließt und das Durchgangsloch (120) durch einen Vorsprung (124) lokalisiert wird;
mindestens eine der ersten und zweiten Taschen (130) sich über eine Tiefe (T) von der Innenfläche (128) der Wand (122) zwischen 10 %-50 % einer Dicke der Wand (122) erstreckt und einen Durchmesser (D) von 100 %-500 % eines Durchmessers (d) des Durchgangslochs (120) aufweist; und
jede Tasche (130) eine Außenkante (132) aufweist, die umlaufend einen Abstand (Z) von einer das Durchgangsloch (120) definierenden Oberfläche entlang einer Umreifungslinie (H) aufweist, und der Abstand (Z) zwischen 10 %-100 % des Durchmessers (d) des Durchgangslochs (120) liegt.
2. Gehäuse nach Anspruch 1, wobei die Wand (122) eine Außenwand eines Diffusorgehäuses (64A) ist.
Verfahren zum Reduzieren von Spannungen in einem Gehäuse (64A) eines Gasturbinentriebwerks (20) nach Anspruch 1 oder 2, umfassend:
Reduzieren von Spannungen um ein Durchgangsloch (120) durch Bereitstellen der ersten und zweiten Taschen (130) auf jeder Seite des Durchgangslochs (120); und
Lokalisieren des Durchgangslochs (120) durch einen Vorsprung (124).
Carter (64A) pour un moteur à turbine à gaz (20), le carter comprenant :
une paroi (122) définissant un trou traversant (120) ; et
des première et seconde poches (130) adjacentes à, et sur des côtés opposés, du trou traversant (120), dans lequel la première et la seconde poches (130) flanquent circonférentiellement le trou traversant et sont définies dans une surface interne (128) de la paroi (122),
caractérisé en ce que :
au moins l'une de la première ou de la seconde poches (130) comporte une périphérie circulaire, et le trou traversant (120) est situé à travers un bossage (124) ;
au moins l'une des première et seconde poches (130) s'étend sur une profondeur (T) depuis la surface interne (128) de la paroi (122) entre 10 % et 50 % d'une épaisseur de la paroi (122) et a un diamètre (D) de 100 % à 500 % d'un diamètre (d) du trou traversant (120) ; et
chaque poche (130) a un bord externe (132) espacé circonférentiellement d'une distance (Z) d'une surface définissant le trou traversant (120) le long d'une ligne de cercle (H), et la distance (Z) est comprise entre 10 % et 100 % du diamètre (d) du trou traversant (120).
2. Carter selon la revendication 1, dans lequel la paroi (122) est une paroi externe d'un carter de diffuseur (64A).
Procédé de réduction de contraintes dans un carter (64A) d'un moteur à turbine à gaz (20) selon la revendication 1 ou 2, comprenant :
la réduction de contraintes autour d'un trou traversant (120) en fournissant les première et seconde poches (130) sur chaque côté du trou traversant (120) ; et
le positionnement du trou traversant (120) à travers un bossage (124) .