Patent Description:
Environmental criteria for aircrafts have been gradually strengthened by requests for conservation of living environment or the like. Noise from an aircraft gas turbine engine (i.e., aircraft jet engine) is one of the objects of the criteria and is also required to be reduced.

Gas turbine engines installed in commercial aircrafts in recent years are mainly turbofan engines that provide good propulsion performance and fuel efficiency. The turbofan engine has a fan to obtain forward thrust. Stator vanes (fan stator vanes) of a fan are provided rearward of rotor blades (fan blades) of the fan. During the operation of the fan, noise (rotor-stator interaction sound) is generated by the aerodynamic interaction of both airfoils.

Patent Literature <NUM> discloses an outlet guide vane intended to reduce the above-described rotor-stator interaction sound.

Patent Literature <NUM> relates to an annular array of turning vanes in a duct of a gas turbine engine. The annular array of turning vanes comprises aerodynamic vanes and strut-vanes. The strut-vanes have a greater chord length and extend further axially downstream than the aerodynamic vanes. The leading edge of the strut-vanes is upstream of the trailing edge of the aerodynamic vanes. The strut-vanes provide flow turning.

Patent Literature <NUM> relates to a propulsor for a gas turbine engine that comprises a case including a duct disposed along an axis to define a flow path, a rotor including a row of propulsor blades extending in a generally radial direction outwardly from a hub. The hub is rotatable about the axis such that the propulsor blades deliver airflow into the flow path, and a row of guide vanes situated in the flow path. A first guide vane comprises a radially inner length, a radially outer length, and a midspan length. The first guide vane has a first dimensional relationship defined as LO/LM and a second dimensional relationship defined as LI/LM. The first dimensional relationship and/or the second dimensional relationship is greater than <NUM>.

Patent Literature <NUM> relates to fan downstream guide vane profiles that have an optimized form of skeleton line angle distribution in an area situated between an upper and a lower limitation as well as a specific thickness distribution superimposed on the respective skeleton line angle distribution.

Patent Literature <NUM> relates to a turbofan engine that includes a row of fan blades disposed upstream from a row of stator vanes. The fan blades are serrated for mixing wakes generated therefrom to attenuate fan noise.

The rotor-stator interaction sound is generated by periodic interaction between a wake (i.e., a velocity defect region) of a rotor blade and a stator vane (e.g., an outlet guide vane) provided rearward of the rotor blade. In addition, the rotor-stator interaction sound is generated not only in the fan but also in other rotating machinery such as a compressor and a turbine. Generally, it is known that the sound pressure level of the rotor-stator interaction sound increases in proportion to the sixth power of a flow velocity when a sound source is a dipole sound source, and increases in proportion to the eighth power of a flow velocity when a sound source is a quadrupole sound source. On the other hand, a decrease in exhaust velocity directly leads to a decrease in thrust. Therefore, it is required to reduce noise while avoiding the fluctuation of the exhaust velocity.

It is an object of the present disclosure to provide a stator vane of an aircraft gas turbine engine and an aircraft gas turbine engine capable of reducing noise generated when a rotating machinery such as a fan is operated.

The invention is defined by an aircraft gas turbine engine according to claim <NUM>.

According to the present invention it is possible to provide a stator vane of an aircraft gas turbine engine and an aircraft gas turbine engine capable of reducing noise generated when a rotating machinery such as a fan is operated.

Hereinafter, exemplary embodiments will be described with reference to the drawings. For convenience of explanation, a turbofan engine is adopted as an example of an aircraft gas turbine engine according to the present embodiment. Further, the turbofan engine is simply referred to as an "engine". It should be noted that the turbofan engine according to the present embodiment may be a geared turbofan engine or other gas turbine engines having a fan. In any case, the bypass ratio does not matter. Furthermore, the stator vane according to the present embodiment can be applied not only to a fan which is rotating machinery (axial flow machine) but also to other rotating machinery (axial flow machines) such as a low-pressure compressor, a high-pressure compressor, a high-pressure turbine, and a low-pressure turbine.

<FIG> is a schematic cross-sectional view of an engine <NUM> according to this embodiment. As shown in this figure, the engine <NUM> includes a core engine <NUM> and a fan <NUM> provided forward of the core engine <NUM>. The core engine <NUM> includes a low-pressure compressor <NUM>, a high-pressure compressor <NUM>, a combustor <NUM>, a high-pressure turbine <NUM>, a low-pressure turbine <NUM>, and a core nozzle <NUM>. These are housed in a core case <NUM> and arranged along the axis <NUM>. In other words, they are arranged from an upstream side (i.e., a forward or a left side in <FIG>) to a downstream side (i.e., a rearward or a right side in <FIG>) of the mainstream of the working fluid (i.e., air or combustion gas). The core engine <NUM> according to the present embodiment is a multi-stage turbine engine. Therefore, the number of stages of compressors and turbines may be, for example, two or three. For convenience of explanation, an extending direction of the axis <NUM> is defined as the axial direction AD. A circumferential direction about the axis <NUM> is defined as a circumferential direction CD. The rotational direction RD of each rotor blade (including a fan blade) described later is assumed to coincide with the circumferential direction CD.

The low-pressure compressor <NUM> includes rotor blades fixed to a low-pressure shaft 16a and stator vanes fixed to an outer wall of the low-pressure compressor <NUM>. The stator vanes and the rotor blades of the low-pressure compressor <NUM> are alternately disposed along the axis <NUM>, and both are arranged in the circumferential direction CD. The low-pressure compressor <NUM> compresses the working fluid flowing into a front core passage <NUM> and supplies it to the high-pressure compressor <NUM>.

The high-pressure compressor <NUM> is provided rearward of the low-pressure compressor <NUM>. The high-pressure compressor <NUM> includes rotor blades fixed to a high-pressure shaft 16b and stator vanes fixed to an outer wall of the high-pressure compressor <NUM>. Similar to the low-pressure compressor <NUM>, the stator vanes and the rotor blades of the high-pressure compressor <NUM> are alternately disposed along the axis <NUM>, and both are arranged in the circumferential direction CD. The high-pressure compressor <NUM> further compresses the working fluid compressed by the low-pressure compressor <NUM> and supplies it to the combustor <NUM>.

The combustor <NUM> is connected with a fuel supply system (not shown). The combustor <NUM> includes an ignitor (not shown), mixes the working fluid compressed by the high-pressure compressor <NUM> with fuel, and combusts the mixed gas. The generated combustion gas is discharged to the high-pressure turbine <NUM>.

The high-pressure turbine <NUM> is provided rearward of the combustor <NUM>. The high-pressure turbine <NUM> includes rotor blades fixed to the high-pressure shaft 16b and stator vanes fixed to an outer wall of the high-pressure turbine <NUM>. The rotor blades and the stator vanes of the high-pressure turbine <NUM> are alternately disposed along the axis <NUM>, and both are arranged in the circumferential direction CD. The combustion gas passes through the rotor blades and the stator vanes of the high-pressure turbine <NUM> while being expanding. In the process of passing, the combustion gas rotates the rotor blades of the high-pressure turbine <NUM>, and this rotational force is transmitted to the high-pressure compressor <NUM> via the high-pressure shaft 16b. Accordingly, the rotor blades of the high-pressure compressor <NUM> is rotated to compress the working fluid.

The low-pressure turbine <NUM> is provided rearward of the high-pressure turbine <NUM>. The low-pressure turbine <NUM> includes rotor blades fixed to the low-pressure shaft 16a and stator vanes fixed to the outer wall of the low-pressure turbine <NUM>. The rotor blades and the stator vanes of the low-pressure turbine <NUM> are alternately disposed along the axis <NUM>, and both are arranged in the circumferential direction CD. The combustion gas discharged from the high-pressure turbine <NUM> passes through the rotor blades and the stator vanes of the low-pressure turbine <NUM> while being expanding. In the process of passing, the combustion gas rotates the rotor blades of the low-pressure turbine <NUM>, and this rotational force is transmitted to the low-pressure compressor <NUM> via the low-pressure shaft 16a. Accordingly, the rotor blades of the low-pressure compressor <NUM> is rotated to compress the working fluid.

The low-pressure shaft 16a is located radially inward of the high-pressure shaft 16b. The low-pressure shaft 16a and the high-pressure shaft 16b are coaxially located about the axis <NUM>, and are rotatably supported by support members such as bearings (not shown). As described above, the low-pressure shaft 16a connects between the low-pressure compressor <NUM> (the rotor blades of the low-pressure compressor <NUM>) and the low-pressure turbine <NUM> (the rotor blades of the low-pressure turbine <NUM>). Similarly, the high-pressure shaft 16b connects between the high-pressure compressor <NUM> (the rotor blades of the high-pressure compressor <NUM>) and the high-pressure turbine <NUM> (the rotor blades of the high-pressure turbine <NUM>).

The core nozzle <NUM> is provided on the downstream side of the low-pressure turbine <NUM>. The core nozzle <NUM> is an annular flow passage formed of a tail cone <NUM> provided at the center thereof and a rearmost part of the core case <NUM>. The core nozzle <NUM> discharges the combustion gas flowing out of the low-pressure turbine <NUM> toward the rear of the core engine <NUM>.

As shown in <FIG>, the fan <NUM> includes rotor blades (fan blades) <NUM> and a fan case <NUM>. The rotor blades <NUM> are attached to a fan rotor <NUM> and radially arranged around the axis <NUM>. The fan rotor <NUM> is connected to the low-pressure shaft 16a. As the low-pressure shaft 16a rotates, the rotor blades <NUM> and the fan rotor <NUM> are rotated integrally. With the rotation of the rotor blades <NUM>, the working fluid flows into the nacelle <NUM> from the outside of the engine <NUM>, and part of the working fluid is guided into the core passage <NUM> in the core case <NUM>.

The fan case <NUM> is a hollow cylindrical member extending along the axis <NUM>, and surrounds a row of rotor blades <NUM>. That is, the maximum diameter of the fan case <NUM> is set to be larger than a diameter of a circle including tips of the rotor blades <NUM>. The length of the fan case <NUM> along the axis <NUM> has a length that accommodates at least the rotor blades <NUM>, an upstream part of the core case <NUM>, and stator vanes <NUM>. That is, the fan case <NUM> accommodates not only the rotor blades (fan blades) <NUM>, but also part of the core engine <NUM> provided rearward of the rotor blades <NUM>, and defines a bypass passage <NUM> with respect to the core case <NUM>. The fan case <NUM> is attached and housed in the nacelle <NUM>. The stator vanes (fan stator vanes) <NUM> are provided in the bypass passage <NUM>.

The core case <NUM> accommodates (covers) rotating machinery such as the low-pressure compressor <NUM> and the combustor <NUM>, which constitutes the core engine <NUM>. The core case <NUM> has a tubular (hollow cylindrical) shape centered on the axis <NUM>. The core case <NUM> has an inner surface 15a and an outer surface 15b. The inner surface 15a constitutes a flow passage of the working fluid taken into the core engine <NUM>, that is, part of the wall surface of the core passage <NUM>. On the other hand, the outer surface 15b is located radially outward of the inner surface 15a and served as a wall surface constituting the bypass passage <NUM>.

The engine <NUM> (in other words, the fan <NUM>) includes the stator vane (fan stator vanes) <NUM> according to the present embodiment. The stator vanes <NUM> are arranged in the circumferential direction CD to regulate the flow of the working fluid discharged from the rotor blades <NUM>. The stator vanes <NUM> are located rearward of the rotor blades <NUM> and extend from the outer surface 15b of the core case <NUM> to the inner surface 22a of the fan case <NUM>. The stator vanes <NUM> are provided in the bypass passage <NUM>, for example, as an outlet guide vane (OGV). In this case, the hub 30a of the stator vane <NUM> is attached to the outer surface 15b of the core case <NUM>, and the tip 30b of the stator vane <NUM> is attached to the inner surface 22a of the fan case <NUM>. However, the hub 30a and tip 30b of the stator vane <NUM> may be supported by corresponding other structural members.

<FIG> is a view illustrating the stator vane <NUM> according to the present embodiment. <FIG> is also a development diagram in the circumferential direction CD. <FIG> is a view illustrating a relationship between an inclination of the rotor blade <NUM> and a position of the maximum airfoil thickness of the stator vane <NUM> (hereinafter, the maximum airfoil thickness position) M with respect to the axial direction AD. <FIG> is a sectional view taken along line IIIA-IIIA (a cross-sectional view on the tip side) as shown in <FIG>. <FIG> is a sectional view taken along line IIIB-IIIB (a cross-sectional view on the hub side) as shown in <FIG>. In other words, <FIG> is a sectional view of a portion of the stator vane <NUM> near the tip 30b. <FIG> is a sectional view of a portion of the stator vane <NUM> near the hub 30a. As shown in <FIG>, <FIG> are cross-sectional views parallel to the axis <NUM>.

The stator vane <NUM> includes an airfoil body <NUM> having an airfoil profile (airfoil cross section) <NUM> shown in <FIG>. The stator vanes <NUM> are arranged in the circumferential direction CD at a predetermined pitch P. The airfoil body <NUM> includes a leading edge 32a, a trailing edge 32b, a suction side (negative pressure surface) 32c, and a pressure side (positive pressure surface) 32d. The suction side 32c and the pressure side 32d extend from the leading edge 32a to the trailing edge 32b. The suction side 32c is a convex surface curved generally toward the rotational direction RD (see <FIG>) of the rotor blade <NUM> (i.e., forward of the circumferential direction CD). The pressure side 32d is a concave surface also curved generally toward the rotational direction RD (see <FIG>) of the rotor blade <NUM> (i.e., forward of the circumferential direction CD). That is, the suction side 32c and the pressure side 32d are both curved in the same direction.

The airfoil profile <NUM> of the airfoil body <NUM> satisfies the following conditions at least on the tip 30b side. In other words, the airfoil body <NUM> has the airfoil profile <NUM> satisfying the following conditions at least on the tip 30b side.

On a plane expanded (developed) in the circumferential direction CD in which the stator vanes <NUM> are arranged, the maximum airfoil thickness position M of the airfoil body <NUM> is:.

In other words, the maximum airfoil thickness position M is located in a third region <NUM> where the first region <NUM> and the second region <NUM> overlap each other. Here, the line <NUM> is a virtual line, which is parallel to an extension line <NUM> of a camber line (airfoil centerline) <NUM> of one rotor blade <NUM> at the trailing edge 21b and passes through a leading edge 40a of another stator vane <NUM> adjacent in the circumferential direction CD (see <FIG>). The extension line <NUM> is tangent to the camber line <NUM> at the trailing edge 21b and extends rearward from the trailing edge 21b. The chord ratio is a value obtained by dividing a distance from the leading edge 32a of the airfoil body <NUM> to an arbitrary position on the chord of the airfoil body <NUM> by the chord length of the airfoil body <NUM>. The stator vane <NUM> described above is one of the stator vanes <NUM> arranged in the circumferential direction CD, and is located forward in the circumferential direction CD by a pitch P from the stator vane <NUM> of interest. The lower limit value and the upper limit value of the chord ratio in the second region <NUM> are set to suppress induction of separation caused by the reason in that a distance from the maximum airfoil thickness position M to the leading edge 32a or a distance from the maximum airfoil thickness position M to the trailing edge 32b becomes extremely short.

Compared the airfoil profile <NUM> of the present embodiment with a conventional airfoil profile at the same span position in a span direction (i.e., radial direction) of their stator vanes, the maximum airfoil thickness position M of the present embodiment is shifted closer to the trailing edge 32b than the maximum airfoil thickness position of the conventional stator vane because of the above conditions. If the maximum airfoil thicknesses of these airfoil profiles are the same, the leading edge 32a of the present embodiment is sharper than the blunt leading edge formed in the conventional stator vane. That is, according to the present embodiment, as compared with the conventional stator vane, the portion <NUM> having a thin airfoil thickness formed near the leading edge 32a is enlarged from the leading edge 32a toward the trailing edge 32b.

When the working fluid passes through the row (cascade) of rotor blades, a wake is generated behind each rotor blade. Since the wake and the mainstream of the working fluid alternately collide with the stator vane, the pressure fluctuates periodically in the vicinity of the stator vane, and a sound (so-called rotor-stator interaction sound) is generated. The generated sound propagates back and forth in the bypass passage, and leaks to the outside of the engine, thereby becoming noise.

A dipole sound source or a quadrupole sound source can be assumed as a sound source model of the above-mentioned sound. The dipole or quadrupole source is a source of pressure oscillations caused by flow disturbances such as wakes and vortices. It is known that the sound pressure level of the dipole sound source is proportional to the sixth power of the velocity of the working fluid. Similarly, it is known that the sound pressure level of the quadrupole sound source is proportional to the eighth power of the velocity of the working fluid. In the present embodiment, by enlarging the portion <NUM> having a thin rotor blade thickness based on the above two conditions, the working fluid flowing in the vicinity of the leading edge 32a of the stator vane <NUM> can be decelerated and the sound pressure can be reduced.

<FIG> is graphs showing surface Mach number distributions on a suction side (SS) and a pressure side (PS) of a stator vane <NUM> according to the present embodiment and on those of a stator vane of a conventional example. The ordinate represents the surface Mach number, and the abscissa represents the chord ratio as described above. The solid line shows the surface Mach number distribution of the stator vane <NUM> according to the present embodiment. The dotted line shows the surface Mach number distribution of the stator vane of the conventional example. In the figure, "~% span" refers to a distance from the hub (base) along the span direction based on the span length as the standard. Thus, <NUM>% span, <NUM>% span, and <NUM>% span refer to positions near the tip of the airfoil body, at the center of the airfoil body, and near the hub of the airfoil body, respectively.

The stator vane <NUM> according to the present embodiment satisfies conditions (a) and (b), and the maximum airfoil thickness position at <NUM>% span is located at a position (i.e., <NUM>% chord length) of the chord ratio <NUM>. The stator vane of the conventional example has the same maximum airfoil thickness as that of the stator vane <NUM> according to the present embodiment and satisfies the above-mentioned condition (b). However, it does not satisfy the condition (a). That is, the maximum airfoil thickness position of the stator vane of the conventional example is located in a region closer to the leading edge than the intersection of the stator vane of the conventional example and a line in the conventional example corresponding to the line <NUM> of the present embodiment.

<FIG> shows the surface Mach number distributions at <NUM>% span on the suction side (SS) and pressure side (PS) of the stator vane <NUM> according to the present embodiment and those of the stator vane of the conventional example. As shown in this figure, it is understood that the surface Mach number distribution in the vicinity of the leading edge 32a of the stator vane <NUM> according to the present embodiment is smaller than that of the stator vane of the conventional example. This decrease appears in both the suction side 32c and the pressure side 32d. According to the computational fluid dynamics (CFD) analysis, the sound pressure distribution in the vicinity of the leading edge 32a also decreases compared with the conventional example. That is, according to the present embodiment, by defining the maximum airfoil thickness position that satisfies the above two conditions, noise can be reduced more than that of the stator vane that does not satisfy the above two conditions.

The airfoil body <NUM> may have the airfoil profile <NUM> which satisfies the above conditions and is provided from the tip 30b side of the airfoil body <NUM> to the hub 30a side of the airfoil body <NUM>. For example, the airfoil profile <NUM> at <NUM>% span assumed in <FIG> satisfies conditions (a) and (b), and the maximum airfoil thickness position is located at the position of the chord ratio <NUM> (i.e., <NUM>% chord length). Similarly, the airfoil profile <NUM> at the <NUM>% span assumed in <FIG> satisfies conditions (a) and (b), and the maximum airfoil thickness position is located at the position of the chord ratio <NUM> (i.e., <NUM>% chord length).

As shown in <FIG>, it can be seen that the surface Mach number distributions in the vicinity of the leading edge 32a at the center and on the hub side of the airfoil body <NUM> are smaller than those of the stator vane of the conventional example. In addition, the reduction of the sound pressure levels is also obtained in the computational fluid dynamics (CFD) analysis, as same as the result at <NUM>% span. Therefore, the sound pressure level can be further reduced by forming the airfoil profile <NUM> satisfying the above conditions over the entire area in the span direction of the airfoil body <NUM>. <FIG> shows a graph as the evidence. The graph shows the results of numerical analysis of the sound pressure levels of the stator vane <NUM> according to the present embodiment and those of the stator vane of the conventional example. The stator vane <NUM> assumed in <FIG> has the airfoil profile <NUM> satisfying the above conditions over the entire area in the span direction. <FIG> shows the sound pressure levels of a harmonic sound of the blade passage frequency (BPF), which is one component of the rotor-stator interaction sound. The left side of the figure shows a comparison result of the sound pressure levels in front of the stator vane (in other words, the sound pressure levels of frontal sounds). The right side of the figure shows a comparison result of the sound pressure levels behind the stator vane (in other words, the sound pressure levels of back sounds). As shown in this figure, the sound pressure levels are decreased both in front of and behind the stator vane.

The chord ratio of the maximum airfoil thickness position M on the tip 30b side of the airfoil body <NUM> may be larger than that on the hub 30a side of the airfoil body <NUM>. For example, the chord ratio of the maximum airfoil thickness position M may continuously increase from the hub 30a toward the tip 30b. As described above, the sound source model for sounds generated by the flow of the working fluid between the rotating rotor blades <NUM> and the stator vanes <NUM> may assume a dipole or quadrupole sound source. The sound pressure levels of these sound sources are proportional to the sixth and eighth power of the flow velocity, respectively. On the other hand, the velocity of the rotor blade <NUM> at each position in the span direction increases in proportion to the distance from the axis <NUM>. That is, the flow velocity of the working fluid is larger on the tip 30b side of the rotor blade <NUM> than on the hub 30a side of the rotor blade <NUM>. Therefore, by enlarging the portion <NUM> on the tip 30b side of the stator vane <NUM> toward the trailing edge 32b more than the portion <NUM> on the hub 30a side, it is possible to promote the suppression of noise on the tip 30b side while suppressing the separation of the working fluid on the hub 30a side.

As described above, the stator vane according to the present embodiment can be applied to any one of the low-pressure compressor <NUM>, the high-pressure compressor <NUM>, the high-pressure turbine <NUM>, and the low-pressure turbine <NUM>. That is, at least one of the rotating machinery may include rotor blades and stator vanes, the stator vane being provided rearward of the rotor blades and each having an airfoil body satisfying conditions (a) and (b) as described above.

Claim 1:
An aircraft gas turbine engine comprising:
a plurality of rotor blades;
a plurality of stator vanes provided rearward of the plurality of rotor blades,
characterized in that
a stator vane of the plurality of stator vanes comprises:
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 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 and passes through a leading edge of another stator vane of the plurality of stator vanes adjacent in the circumferential direction, and
(b) the maximum airfoil thickness position is located in a region between a position of a chord ratio <NUM> and a position of chord ratio <NUM>,
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 chord.