Abstract: A blade for a gas turbine engine includes an aerofoil portion and a root portion defined by concave and convex walls having opposing inner surfaces. A reinforcing member is located between the concave and convex walls and is bonded to the inner surfaces of the concave and convex walls, and the root portion includes an unbonded region in which the reinforcing member contacts an inner surface of one of the concave and convex walls but is not bonded to the concave and convex walls.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprises a generally cylindrical, or frustoconical, metal casing (40). The metal casing (40) has at least one arched portion (62,64,66) and each arched portion (62,64,66) extends circumferentially around the metal casing (40). Each arched portion (62,64,66) comprises a radially inwardly extending arch (68) in which the trough (74) of the arch (68) is radially inwardly of a straight line (76) interconnecting the axial ends (70,72) of the arch (68). The trough (74) of the arch (68) is radially inwardly of the at least one of the axial ends (70,72) of the at least one arch (68). The arched portions (62,64,66) introduce compressive stresses into the metal casing (40) during a fan blade 34 impact and enable the metal casing (40) to withstand the impact. The fan blade containment assembly (38) is lighter for large diameter turbofan gas turbine engine (10).

Abstract: The fan cowl (12) of a ducted fan gas turbine engine (10) has a downstream portion including an inner casing (30) that surrounds a honeycomb structure (36) in abutting engagement. Honeycomb structure (36) is made up of a number of closely spaced honeycomb blocks (36a) fixed about a liner (40), one or more of which blocks, if struck by a broken fan blade portion, will move relative to those not struck, in a radially outward direction, thus reducing the crushing force of the blow, and only flexing casing (30) instead of permanently deforming or puncturing it.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprises a containment casing (40) and the containment casing (40) has a plurality of axially spaced ribs (66). The ribs (66) extend circumferentially and radially inwardly from the containment casing (40) to define a plurality of circumferentially extending grooves (76) and a perforate sheet (70) is secured to the radially inner ends (68) of the ribs (66). In the arrangement of the present invention the blade containment function and the acoustic treatment function are combined.

Abstract: A blade (10, FIG. 1) for a gas turbine engine includes an aerofoil portion (12, FIG. 1) and a root portion 14 defined by concave and convex walls 16, 18 having opposing inner surfaces 16a, 18a. A reinforcing member 20 is located between the concave and convex walls 16, 18 and is bonded to the inner surfaces 16a, 18a thereof, and the root portion 14 includes an unbonded region 24a, 24b in which the reinforcing member 20 contacts an inner surface 16a, 18a of one of the concave and convex walls 16, 18 but is not bonded thereto.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprises a generally cylindrical, or frustoconical, metal casing (40). The metal casing (40) has at least one arched portion (62, 64, 66) and each arched portion (62, 64, 66) extends circumferentially around the metal casing (40). Each arched portion (62, 64, 66) comprises a radially inwardly extending arch (68) in which the trough (74) of the arch (68) is radially inwardly of a straight line (76) interconnecting the axial ends (70, 72) of the arch (68). The trough (74) of the arch (68) is radially inwardly of the at least one of the axial ends (70, 72) of the at least one arch (68). The arched portions (62, 64, 66) introduce compressive stresses into the metal casing (40) during a fan blade 34 impact and enable the metal casing (40) to withstand the impact. The fan blade containment assembly (38) is lighter for large diameter turbofan gas turbine engine (10).

Abstract: A joint assembly for limiting an extension of the joint in the direction of a load path derived from an impact comprising a first member having a portion and a second member having a portion, the portions overlapping one another and arranged generally parallel to one another and secured together via securing means disposed through corresponding holes defined therein. One of the overlapping portions further defines, sequentially in the direction of extension, a shear neck, a pocket and a catcher portion. In the event of a worst-case impact load the securing means shears through the shear neck and the pocket and is arrested by the catcher portion, thereby the extension of the joint assembly is limited and the joint assembly remains integral.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) has an upstream portion (56), a transition portion (58) and a blade containment portion (54) and a downstream portion (60). The upstream portion (56) has a flange (42) connecting the metal casing (40) to a flange (48) on axially adjacent casing (46). The blade containment portion (54) has a greater thickness (T2) than the thickness (T1) of the upstream portion (54) and the downstream portion (60). The downstream portion (60) has impact protection means (64) located on its inner surface (62) to protect the downstream portion (60) of the containment casing (40). The impact protection means (64) comprises a plurality of radially inwardly and circumferentially extending ribs (80) on the inner surface (62) of the downstream portion (60) to act as spacer between an inner portion of a detached fan blade (34) and the downstream portion (60).

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) which has an inner surface (62), an outer surface (64), an upstream flange (42) and a downstream flange (52). The outer surface (62) of the metal casing (40) is provided with a pattern of blind apertures (66) which extend radially inwardly from the outer surface (62) of the metal casing (40) between the flanges (42,52). The fan blade containment has reduced weight of the metal casing (40) of the fan blade containment assembly (38) for a large diameter turbofan gas turbine engine (10) but has unimpaired stiffness and penetration resistance.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) has an upstream portion (56), a transition portion (58) and a blade containment portion (54). The upstream portion (56) has a flange (42) connecting the metal casing (40) to a flange (48) on axially adjacent casing (46). The blade containment portion (54) has a greater thickness (T2) than the thickness (T1) of the upstream portion (54). The transition portion (58) connects the blade containment portion (54) and the upstream portion (56) to transmit loads from the blade containment portion (54) to the flange (42) on the upstream portion (56). The ratio of the thickness (T4) of the flange (42) on the upstream portion (56) to the thickness (T1) of the upstream portion (56) is between 3 to 1 and 6 to 1.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) which has an inner surface (62), an outer surface (64), an upstream flange (42) and a downstream flange (52). The outer surface (62) of the metal casing (40) is provided with a pattern of blind apertures (66) which extend radially inwardly from the outer surface (62) of the metal casing (40) between the flanges (42,52). The fan blade containment has reduced weight of the metal casing (40) of the fan blade containment assembly (38) for a large diameter turbofan gas turbine engine (10) but has unimpaired stiffness and penetration resistance.

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor (32) carrying a plurality of radially extending fan blades (34). A fan blade containment assembly (38) surrounds the fan blades (34) and the fan blade containment assembly (34) comprises a generally annular, or frustoconical, cross-section casing (52). At least one corrugated sheet metal ring surrounds the casing (52) wherein the corrugations extend with axial and/or circumferential components.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) which has an inner surface (62), an outer surface (64), an upstream flange (42) and a downstream flange (52). The outer surface (62) of the metal casing (40) is provided with a pattern of blind apertures (66) which extend radially inwardly from the outer surface (62) of the metal casing (40) between the flanges (42,52). The fan blade containment has reduced weight of the metal casing (40) of the fan blade containment assembly (38) for a large diameter turbofan gas turbine engine (10) but has unimpaired stiffness and penetration resistance.

Abstract: A joint assembly for limiting an extension of the joint in the direction of a load path derived from an impact comprising a first member having a portion and a second member having a portion, the portions overlapping one another and arranged generally parallel to one another and secured together via securing means disposed through corresponding holes defined therein. One of the overlapping portions further defines, sequentially in the direction of extension, a shear neck, a pocket and a catcher portion. In the event of a worst-case impact load the securing means shears through the shear neck and the pocket and is arrested by the catcher portion, thereby the extension of the joint assembly is limited and the joint assembly remains integral.

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) has an upstream portion (56), a transition portion (58) and a blade containment portion (54) and a downstream portion (60). The upstream portion (56) has a flange (42) connecting the metal casing (40) to a flange (48) on axially adjacent casing (46). The blade containment portion (54) has a greater thickness (T2) than the thickness (T1) of the upstream portion (54) and the downstream portion (60). The downstream portion (60) has impact protection means (64B) located on its inner surface (62) to protect the downstream portion (60) of the containment casing (40). The impact protection means (64) comprises a plurality of radially inwardly and circumferentially extending ribs (80) on the inner surface (62) of the downstream portion (60) to act as spacer between an inner portion of a detached fan blade (34) and the downstream portion (60).

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) has an upstream portion (56), a transition portion (58) and a blade containment portion (54). The upstream portion (56) has a flange (42) connecting the metal casing (40) to a flange (48) on axially adjacent casing (46). The blade containment portion (54) has a greater thickness (T2) than the thickness (T1) of the upstream portion (54). The transition portion (58) connects the blade containment portion (54) and the upstream portion (56) to transmit loads from the blade containment portion (54) to the flange (42) on the upstream portion (56). The ratio of the thickness (T4) of the flange (42) on the upstream portion (56) to the thickness (T1) of the upstream portion (56) is between 3 to 1 and 6 to 1.

Abstract: A gas turbine engine rotor blade containment assembly (38) comprising a generally cylindrical, or frustoconical, first metal casing (40) and a plurality of generally cylindrical, or frustoconical axially spaced second metal casing is (54, 56 & 58). The second metal casings (54, 56, 58) surrounded and abut the first metal casing (40). The first metal casing (40) is relatively hard and tough to contain a fan blade (34). The second metal casings (54, 56, 58 are relatively soft and lightweight to stiffen the first metal casing (40) and fan blade containment regions (A, B & C). The second metal casing (54, 56, 58) are arranged coaxially with the first metal casing (40). One of the second metal casing means (56) is in a plane (Y) containing the fan blades (34) of the gas turbine engine (10). The first metal casing (40) comprises steel and the second metal casing is (54, 56, 58) comprise aluminum. The fan blade containment assembly (38) is lighter for large diameter turbofan gas turbine engine (10).

Abstract: A turbofan gas turbine engine (10) comprises a fan rotor (32) carrying a plurality of radially extending fan blades (34). A fan blade containment assembly (38) surrounds the fan blades (34) and the fan blade containment assembly (38) comprises a generally cylindrical, or frustoconical, metal casing (40). The casing (40) comprises an annular hook (54) positioned axially upstream of the tip (37) of the fan blade (34). The annular hook (54) extends in a radially inwardly and axially downstream direction from the metal casing (40) towards the tip (37) of the fan blade (34). The metal casing (40) has a containment portion (A) downstream of the annular hook (54) and substantially in the plane of the leading edge of the tip (37) of the fan blade (34). The containment portion (A) of the casing (40) has a greater diameter than the diameter of the casing (40) at the annular hook (54) to restrain upstream movement of a detached fan blade (34).

Abstract: A gas turbine engine fan blade containment assembly (38) comprising a generally cylindrical, or frustoconical, metal casing (40) which has an inner surface (62), an outer surface (64), an upstream flange (42) and a downstream flange (52). The outer surface (62) of the metal casing (40) is provided with a pattern of blind apertures (66) which extend radially inwardly from the outer surface (62) of the metal casing (40) between the flanges (42,52). The fan blade containment has reduced weight of the metal casing (40) of the fan blade containment assembly (38) for a large diameter turbofan gas turbine engine (10) but has unimpaired stiffness and penetration resistance.

Abstract: A containment casing for a gas turbine engine comprising a substantially rigid casing shell arranged in operation coaxially with an axis of rotation of the gas turbine engine. The casing extending circumferentially around an array of fan rotor blades which are arranged to rotate about the engine axis. In the region of the casing where it is predicted blade impact in the event of a fan blade failure, there is at least one reinforcing rib extending substantially radially from the casing and circumscribing the outer periphery of the casing shell. The ribs providing improved strengthening of the casing. Preferably a first rib is positioned about the casing between the operational position of the trailing edge and the mid chord point of the fan blade. A second rib is axially positioned between the operational position of the mid chord point and the leading edge of the fan blade. A third rib may also be positioned axially forward of the operational position of the leading edge of the fan blade.