Stator vane and aircraft gas turbine engine
A stator vane provided rearward of a rotor blade, includes: an airfoil body having an airfoil profile. The maximum airfoil thickness position of the airfoil body on the airfoil profile satisfies following conditions at least on a tip side of the airfoil body: (a) on a plane expanded in a circumferential direction in which the stator vane is arranged, the position is located in a first region which is close to a trailing edge of the airfoil body from an intersection of the airfoil body and a line which is parallel to an extension line of a camber line of the rotor blade at a trailing edge of the rotor blade and passes through a leading edge of another stator vane adjacent in the circumferential direction, and (b) the position is located in a second region having a chord ratio from 0.2 to 0.8.
1. An assembly comprising a plurality of stator vanes provided rearward of a plurality of rotor blades, a stator vane of the plurality of stator vanes comprising:
an airfoil body having an airfoil cross section, wherein
a maximum airfoil thickness position of the airfoil body in the airfoil cross section satisfies following conditions at least on a tip side of the airfoil body:
(a) on a plane expanded in a circumferential direction in which the plurality of stator vanes are arranged, the maximum airfoil thickness position is located in a region between a trailing edge of the airfoil body and an intersection of the airfoil body, wherein the intersection is between the airfoil body and a line which is parallel to an extension line of a camber line of a rotor blade of the plurality of rotor blades at a trailing edge of the rotor blade of the plurality of rotor blades, the extension line also passes through a leading edge of a circumferentially adjacent stator vane of the plurality of stator vanes, and
(b) the maximum airfoil thickness position is located in a region between a position of a chord ratio 0.2 and a chord ratio 0.8, and
the chord ratio is a value obtained by dividing a distance from the leading edge of the airfoil body to a position on a chord of the airfoil body by a full length of the airfoil body chord.
2. The assembly according to claim 1 , wherein
the airfoil body has the airfoil cross section in which the maximum airfoil thickness position satisfying the conditions, from the tip side of the airfoil body to a hub side of the airfoil body.
3. The assembly according to claim 2 , wherein
the chord ratio of the maximum airfoil thickness position on the tip side of the airfoil body is larger than the chord ratio of the maximum airfoil thickness position on the hub side of the airfoil body.
4. The assembly according to claim 3 , wherein
the plurality of stator vanes are fan stator vanes.
5. The assembly according to claim 2 , wherein
the plurality of stator vanes are fan stator vanes.
6. The assembly according to claim 1 , wherein
the plurality of stator vanes are fan stator vanes.
7. An aircraft gas turbine engine comprising:
the assembly according to claim 1 , the plurality of stator vanes being a plurality of fan stator vanes, and the plurality of rotor blades being a plurality of fan blades;
a fan case accommodating the plurality of fan blades; and
a core case configured to accommodate a core engine provided rearward of the plurality of fan blades and defines a bypass passage of a working fluid between the core case and the fan case, wherein
the plurality of stator vanes are provided in the bypass passage.