Patent Description:
The drag that acts on an aircraft can be classified into two of pressure drag and friction drag.

Of those, the pressure drag is force that pulls an object backward by separating air around the object, creating vortexes backwards, and reducing the pressure. The pressure drag is one kind of profile drag that varies depending on only the shape of an object.

In a transonic airliner, approximately <NUM>% of the pressure drag occurs by a main wing, and thus a reduction in pressure drag of the main wing plays a significant role in reduction of the whole drag of the airliner.

The main wing of transonic airliners that is currently operated has an sweptback angle to delay the influence of compressibleness of the air generated on the wings, that is, the damage caused by a shock wave. Furthermore, transonic airfoils such as a peaky airfoil, a rear-loading airfoil, and a supercritical airfoil, in which a cross-sectional shape of a wing, i.e., an airfoil, has a flat upper wing surface and is configured to suppress acceleration, are employed to devise generation of a gentle shock wave.

The peaky airfoil, which is a representative transonic airfoil, is a transonic airfoil but causes no shock wave. Meanwhile, the supercritical airfoil causes a shock wave, but the shock wave is very weak (see Non-Patent Literatures <NUM> to <NUM> and Patent Literatures <NUM> to <NUM>).

However, to improve sustainability of environment and resources in air transport, which is predicted to continue expanding, it is necessary to further reduce aerodynamic drag.

For one means of such reduction, the inventors of the present invention are examining the technology in which the technology of designing a natural laminar flow wing for reducing friction drag is applied to a concept design of a transonic aircraft (see Patent Literature <NUM> and the like). In the process of the examination, the inventors of the present invention have discovered an airfoil that significantly reduces pressure drag (Non-Patent Literature <NUM>).

Thus, it is an object of the present invention to provide a transonic airfoil capable of reducing pressure drag more than before, a wing having such an airfoil, and an aircraft including such a wing.

The following facts have been known: thrust occurs if the leading edge of a wing is twisted down; and a curvature distribution that forms an airfoil has high sensitivity with respect to a pressure distribution.

However, the shape of the leading edge of the airfoil, and its change in the chord direction, which are points for reducing the pressure drag of the transonic airfoil, are not clarified. The inventors of the present invention have been keenly examined to clarify those points and then solved the above problems.

A transonic airfoil according to an example not according to the claims has a shape in which a pressure coefficient Cp of a static pressure in a chord direction of a leading edge is -<NUM> or less at z/c=<NUM>, where z represents a coordinate in a direction perpendicular to an airflow direction within a plane that forms an airfoil, with the leading edge being as a reference (an upper wing surface direction is positive, and a lower wing surface direction is negative), and c represents a chord length.

The transonic airfoil may further have a shape in which the pressure coefficient Cp of the static pressure in the chord direction of the leading edge is -<NUM> or less at z/c=<NUM>.

In the transonic airfoil according to one embodiment of the present invention, κ has a local maximal value of <NUM> or more in an upwardly convex curve in a range of -<NUM><s/c<<NUM>, Ks is <NUM> or more in a range from s/c=-<NUM> to s/c=<NUM>, and κ is <NUM> or less in a range from s/c=<NUM> to s/c=<NUM>, where s represents a surface length along a surface of the airfoil, with the leading edge being as a reference (the upper wing surface direction is positive, and the lower wing surface direction is negative), κ represents a curvature that is made dimensionless by a reciprocal of the chord length, and Ks represents an integral value of the curvature κ.

Furthermore, in the transonic airfoil, κ may be less than <NUM> at s/c=<NUM>, and the κ being less than <NUM> may increase to be <NUM> or more at s/c=<NUM>.

Furthermore, in the transonic airfoil, κ may have a local maximal value of <NUM> or more in an upwardly convex curve in a range from s/c=-<NUM> or more to a position of a trailing edge.

In the transonic airfoil, κ may monotonically decrease in a range from a stagnation point to a crest position of a lower wing surface, and Ks may be <NUM> or more in a range from s/c=-<NUM> to s/c=-<NUM>.

Furthermore, in the transonic airfoil, κ may have a mean value of <NUM> or less in a range from s/c=-<NUM> to s/c=-<NUM>, and κ may be <NUM> or less at s/c=-<NUM>.

Furthermore, in the transonic airfoil, a distribution of κ may monotonically increase to <NUM> or more in a range from s/c=-<NUM> or less to a position of a trailing edge.

A wing according to one embodiment of the present invention has the transonic airfoil described above. Additionally, an aircraft according to one embodiment of the present invention includes a main wing having the transonic airfoil described above.

According to the present invention, it is possible to reduce pressure drag more than before.

<FIG> is a schematic perspective view of an aircraft according to one embodiment of the present invention.

An aircraft <NUM> includes a main wing <NUM>, an empennage <NUM>, and the like provided to a fuselage <NUM>.

The main wing <NUM> has a transonic airfoil according to the present invention.

<FIG> is a diagram showing a dimensionless airfoil of the main wing <NUM>.

A reference numeral <NUM> represents a two-dimensional airfoil (airfoil) of the main wing <NUM>. The two-dimensional airfoil <NUM> includes two-dimensional elements in a chord direction, which are arranged in a wingspan direction to configure three-dimensional elements (three-dimensional wing) attached to mainly generate lift in the aircraft <NUM>.

A reference numeral <NUM> represents a leading edge, and a reference numeral <NUM> represents a trailing edge. The leading edge <NUM> and the trailing edge <NUM> are two-dimensional elements at positions having a minimum value and a maximum value of the coordinate in the chord direction.

In the diagram, the upper side from a line segment <NUM> connecting the leading edge <NUM> and the trailing edge <NUM> is the upper surface of the main wing <NUM>, and the lower side from the line segment <NUM> is the lower surface of the main wing <NUM>.

A reference symbol x represents a coordinate in an airflow direction with the leading edge <NUM> being as a reference, a reference symbol y represents a coordinate in the wingspan direction orthogonal to the airfoil <NUM>, and a reference symbol z represents a coordinate in a direction perpendicular to x within the plane that forms the airfoil <NUM>, with the leading edge <NUM> being as a reference.

A reference symbol c represents a chord length, that is, a maximum length between any two points on the airfoil <NUM>.

In the diagram, the unit of each of the x-axis and the z-axis is x/c and z/c, respectively, which are dimensionless.

A reference symbol θ represents an angle defined by a line <NUM>, which connects the center of the airfoil <NUM> (x/c=<NUM>, z/c=<NUM>) and any point on the airfoil <NUM>, and a line satisfying z/c=<NUM> (X-axis). The upper surface side is assumed as positive, and the lower surface side is assumed as negative.

A reference symbol s represents a surface length along the surface of the airfoil <NUM> with the leading edge <NUM> being as a reference. The upper surface side is assumed as positive, and the lower surface side is assumed as negative.

A reference numeral <NUM> represents a stagnation point, a reference numeral <NUM> represents an upper surface crest, and a reference numeral <NUM> represents a lower surface crest. The stagnation point <NUM> is a position, at which the velocity of the fluid is zero, on the surface of the two-dimensional element into the airflow. The stagnation point <NUM> is located near the leading edge <NUM> in an actual flow with viscosity. The crest means a position at which the z coordinate is maximum or minimum on the airfoil <NUM>. The maximum position is referred to as an upper surface crest, and the minimum position is referred to as a lower surface crest.

A reference numeral <NUM> represents a mid-chord. The mid-chord <NUM> is a middle region between the leading edge <NUM> and the trailing edge <NUM> of the two-dimensional airfoil <NUM>.

Additionally, a reference numeral <NUM> represents lift, a reference numeral <NUM> represents drag, and a reference numeral <NUM> represents thrust. The lift <NUM> is the force in the airflow direction and the perpendicular direction that acts by movement of the two-dimensional element in the air. The drag <NUM> is the force in the airflow direction that acts by movement of the two-dimensional element in the air. The thrust <NUM> is the force in the direction opposite to the airflow direction that acts by movement of the two-dimensional element in the air. Note that pressure drag is, in the drag <NUM>, drag generated by the pressure of the surface of the two-dimensional element, and pressure thrust is, in the thrust <NUM>, thrust pressure generated by the pressure of the surface of the two-dimensional element.

In this specification, a reference symbol κ represents a curvature that is made dimensionless by the reciprocal of the chord length c, and a reference symbol K is an integral value of the curvature θ. Here, Kθ and Ks are as follows.

<FIG> is a graph (part <NUM>) showing a pressure distribution of a static pressure in the chord direction of the airfoils <NUM> according to this embodiment and airfoils as reference examples.

In <FIG>, a thick solid line A indicates a pressure distribution of an airfoil <NUM> in a first mode according to this embodiment, a medium solid line B indicates a pressure distribution of an airfoil <NUM> in a second mode according to this embodiment, a thin solid line C indicates a pressure distribution of an airfoil <NUM> in a third mode according to this embodiment. Additionally, a dotted line D indicates a pressure distribution of an RAE <NUM> airfoil (see Non-Patent Literature <NUM>), a chain line E indicates a pressure distribution of a CRM airfoil (see Non-Patent Literature <NUM>), and a chain double-dashed line F indicates a pressure distribution of a Baseline airfoil (see Non-Patent Literature <NUM>).

Note that, also in the graphs to be shown below, the thick solid line A indicates data regarding the airfoil <NUM> in the first mode according to this embodiment, the medium solid line B indicates data regarding the airfoil <NUM> in the second mode according to this embodiment, the thin solid line C indicates data regarding the airfoil <NUM> in the third mode according to this embodiment, the dotted line D indicates data regarding the RAE <NUM> airfoil, the chain line E indicates data regarding the CRM airfoil, and the chain double-dashed line F indicates data regarding the Baseline airfoil.

The airfoils <NUM> in the first to third modes according to this embodiment have a shape in which a pressure coefficient Cp of a static pressure in the chord direction of the leading edge <NUM> is -<NUM> or less at z/c=<NUM>.

With this shape, the airfoils <NUM> in the first to third modes according to this embodiment have a sharp rise of the pressure distribution as compared with the airfoils illustrated as reference examples, and can thus decrease the pressure drag. Decreasing the pressure drag leads to decreasing the drag <NUM> and increasing the thrust <NUM>.

Note that the airfoils <NUM> in the first to third modes according to this embodiment favorably have a shape in which the pressure coefficient Cp of the static pressure in the chord direction of the leading edge <NUM> is -<NUM> or more at z/c=<NUM>. This is because an extremely low Cp is highly likely to generate a large adverse pressure gradient in the downstream, cause boundary layer separation, and thus increase the pressure drag.

Additionally, <FIG> is a graph of the pressure distribution of the airfoil <NUM> in the first mode, which is extracted from <FIG>. As a hatched region (inverted region) denoted by a reference symbol S becomes larger, the thrust increases. As seen from <FIG>, the airfoils <NUM> in the first to third modes according to this embodiment have a larger area of the inverted region than those of the airfoils illustrated as reference examples. This results from the shape in which the pressure coefficient Cp of the static pressure in the chord direction of the leading edge <NUM> is -<NUM> or less at z/c=<NUM>. Therefore, the airfoils <NUM> in the first to third modes according to this embodiment have an increased thrust also by an increase of the inverted region, as compared with the airfoils illustrated as reference examples.

According to the knowledge of the inventors of the present invention, it has been found that the area of the inverted region of the airfoil <NUM> according to this embodiment increases by approximately <NUM>% to <NUM>% as compared with the related art, and thus the thrust corresponding thereto increases.

<FIG> is a graph (part <NUM>) showing the pressure distribution of the static pressure in the chord direction of the airfoils <NUM> according to this embodiment and airfoils as reference examples.

The airfoils <NUM> in the first to third modes according to this embodiment further have a shape in which the pressure coefficient Cp of the static pressure in the chord direction of the leading edge <NUM> is -<NUM> or less at z/c=<NUM>.

Thus, the airfoils <NUM> in the first to third modes according to this embodiment have a further expanded inverted region, and the effect of increasing the thrust can be enhanced.

As described above, the airfoil <NUM> according to this embodiment is characterized by having a sharp rise of the pressure distribution. The modes in shape of the airfoil <NUM> for such a purpose will be described below.

<FIG> is a graph (part <NUM>) showing a relationship between s/c and κ of the airfoils <NUM> according to this embodiment and the airfoils illustrated as reference examples. <FIG> is a graph (part <NUM>) showing a relationship between s/c and Ks of the airfoils <NUM> according to this embodiment and the airfoils illustrated as reference examples. <FIG> is a graph (part <NUM>) showing a relationship between s/c and κ of the airfoils <NUM> according to this embodiment and the airfoils illustrated as reference examples.

The airfoils <NUM> according to this embodiment have a shape in which, as shown in <FIG>, κ has a local maximal value of <NUM> or more in an upwardly convex curve in the range of -<NUM><s/c<<NUM>, and as shown in <FIG>, Ks is <NUM> or more in the range from s/c=-<NUM> to s/c=<NUM>. Such a shape can sharply decrease the pressure and can increase the thrust.

Note that, in the airfoils <NUM> according to this embodiment, κ favorably has a local maximal value of <NUM> or less in the range of -<NUM><s/c<<NUM>. Additionally, Ks is favorably <NUM> or less in the range from s/c=-<NUM> to s/c=<NUM>. This is because an extremely sharp shape of the leading edge is highly likely to cause a stall if the angle of attack of the airframe is changed.

Additionally, the airfoils <NUM> according to this embodiment have a shape in which κ is <NUM> or less in the range from s/c=<NUM> to s/c=<NUM>, which is near a position where a shock wave is generated, as shown in <FIG>. Thus, the pressure lowered with the shape shown in <FIG> and <FIG> is kept low also at that position (from s/c=<NUM> to s/c=<NUM>), and thus the thrust further increases.

Note that, in the airfoils <NUM> according to this embodiment, κ is favorably <NUM> or more in the range from s/c=<NUM> to s/c=<NUM>. This is because a flat shape or a recessed shape with a negative curvature in this region is highly likely to cause the boundary layer separation and thus increase the pressure drag.

The airfoils <NUM> according to this embodiment have the shape described above and can set the pressure coefficient Cp of the static pressure in the chord direction of the leading edge <NUM> to be -<NUM> or less at z/c=<NUM> and further set the pressure coefficient Cp of the static pressure in the chord direction of the leading edge <NUM> to be -<NUM> or less at z/c=<NUM>. This can reduce the pressure drag and also enlarge the inverted region, and thus increase the thrust.

<FIG> is a graph (part <NUM>) showing a relationship between s/c and κ of the airfoils <NUM> according to this embodiment and the airfoils illustrated as reference examples.

The airfoils <NUM> according to this embodiment have a shape in which, as shown in <FIG>, κ being near the position of the upper surface crest <NUM> is less than <NUM> at s/c=<NUM>, and κ, which is less than <NUM>, increases to be <NUM> or more at s/c=<NUM>. Positions of s/c=<NUM> and s/c=<NUM> are around the position at which a shock wave occurs.

With such a shape, the pressure in the rear of the position of the upper surface crest <NUM> rises, and the thrust further increases.

Note that κ is favorably <NUM> or more at s/c=<NUM>. This is because a flat shape or a recessed shape with a negative curvature in this region is highly likely to cause the boundary layer separation and thus increase the pressure drag. Additionally, it is favorable that κ increases to <NUM> or less at s/c=<NUM>. This is because an excessively large curvature is highly likely to generate a large adverse pressure gradient in the downstream, cause the boundary layer separation, and thus increase the pressure drag.

The airfoils <NUM> according to this embodiment have a shape in which, as shown in <FIG>, κ has a local maximal value of <NUM> or more in an upwardly convex curve in the range from s/c=<NUM> or more to the position of the trailing edge.

With this shape, the pressure, which is kept low from s/c=<NUM> to s/c=<NUM>, increases by the shock wave and then further increases at that position (position of local maximal value), and thus the drag decreases.

Note that κ favorably has a local maximal value of <NUM> or less. This is because an extremely large curvature at the trailing edge is highly likely to generate a large adverse pressure gradient in the downstream, cause the boundary layer separation, and thus increase the pressure drag.

<FIG> is a graph (part <NUM>) showing a relationship between s/c and Ks of the airfoils <NUM> according to this embodiment and the airfoils illustrated as reference examples.

The airfoils <NUM> according to this embodiment have a shape in which, as shown in <FIG>, κ monotonically decreases in the range from the stagnation point <NUM> to the crest position <NUM> of the lower wing surface, and Ks, which is the integral value of the curvature, is <NUM> or more in the range from s/c=-<NUM> to s/c=-<NUM>.

Thus, the pressure sharply decreases, and thus the thrust increases.

Note that Ks, which is the integral value of the curvature, is favorably <NUM> or less in the range from s/c=-<NUM> to s/c=-<NUM>. This is because an extremely sharp shape of the leading edge is highly likely to cause a stall if the angle of attack of the airframe is changed.

The airfoils <NUM> according to this embodiment have a shape in which, as shown in <FIG>, κ has a mean value of <NUM> or less in the range from s/c=-<NUM> to s/c=-<NUM> near the crest position <NUM> of the lower wing surface, and κ is <NUM> or less at s/c=-<NUM>.

Thus, the pressure lowered with the shape shown in <FIG> is kept low in that position (from s/c=-<NUM> to s/c=-<NUM>), and thus the thrust further increases.

Note that κ favorably has a mean value of <NUM> or more in the range from s/c=-<NUM> to s/c=-<NUM>. This is because a flat shape or a recessed shape with a negative curvature in this region is highly likely to cause the boundary layer separation and thus increase the pressure drag.

The airfoils <NUM> according to this embodiment have a shape in which, as shown in <FIG>, the distribution of κ monotonically increases to <NUM> or more in the range from s/c=-<NUM> or less to the position of the trailing edge <NUM>.

Thus, the pressure rises, and the drag decreases.

<FIG> is a graph showing the dimensionless airfoils <NUM> of this embodiment and dimensionless airfoils used as reference examples. <FIG> is a graph showing the vicinity of the leading edge of the <FIG> in an enlarged manner.

The airfoils <NUM> according to the embodiment described above can reduce the pressure drag of a transonic wing having the airfoil <NUM> by approximately <NUM>% of the whole aerodynamic drag of the transonic aircraft. This corresponds to approximately ten times the friction drag reduced by achieving a natural laminar flow.

Note that the airfoils <NUM> according to the embodiment described above have a shape in which the pressure coefficient Cp of the static pressure in the chord direction of the leading edge <NUM> is -<NUM> or less at z/c=<NUM>. This is expressed using the relationship between Cp and x/c, which corresponds to the fact that Cp takes a negative value at x/c=<NUM>, as shown in <FIG>. Note that <FIG> is a graph showing the relationship between Cp and x/c of the airfoils <NUM> shown in <FIG> in the vicinity of the leading edge in an enlarged manner.

<FIG> show a relationship between x/c and Cp, a relationship between x/c and z/c, a relationship between z/c and Cp, a relationship between z/c and κ, a relationship between the angle from the center θ and Kθ, a relationship between the angle from the center θ and Cp, and a relationship between the angle from the center θ and κ, respectively, of the airfoil <NUM> in the first mode according to the this embodiment.

<FIG> show a relationship between x/c and Cp, a relationship between x/c and z/c, a relationship between z/c and Cp, a relationship between z/c and κ, a relationship between the angle from the center θ and Kθ, a relationship between the angle from the center θ and Cp, and a relationship between the angle from the center θ and κ, respectively, of the airfoil <NUM> in the second mode according to the this embodiment.

<FIG> show a relationship between x/c and Cp, a relationship between x/c and z/c, a relationship between z/c and Cp, a relationship between z/c and κ, a relationship between the angle from the center θ and Kθ, a relationship between the angle from the center θ and Cp, and a relationship between the angle from the center θ and κ, respectively, of the airfoil <NUM> in the third mode according to the this embodiment.

<FIG> show a relationship between x/c and Cp, a relationship between x/c and z/c, a relationship between z/c and Cp, a relationship between z/c and κ, a relationship between the angle from the center θ and Kθ, a relationship between the angle from the center θ and Cp, and a relationship between the angle from the center θ and κ, respectively, of the RAE <NUM> airfoil.

<FIG> show a relationship between x/c and Cp, a relationship between x/c and z/c, a relationship between z/c and Cp, a relationship between z/c and κ, a relationship between the angle from the center θ and Kθ, a relationship between the angle from the center θ and Cp, and a relationship between the angle from the center θ and κ, respectively, of the CRM airfoil.

<FIG> show a relationship between x/c and Cp, a relationship between x/c and z/c, a relationship between z/c and Cp, a relationship between z/c and κ, a relationship between the angle from the center θ and Kθ, a relationship between the angle from the center θ and Cp, and a relationship between the angle from the center θ and κ, respectively, of the Baseline airfoil.

Claim 1:
A transonic airfoil (<NUM>), in which
a curvature that is made dimensionless by a reciprocal of the chord length has a local maximal value of <NUM> or more in an upwardly convex curve in a range of -<NUM><s/c<<NUM>,
an integral value of the curvature is <NUM> or more in a range from s/c=-<NUM> to s/c=<NUM>, and
the curvature is <NUM> or less in a range from s/c=<NUM> to s/c=<NUM>, where
s represents a surface length along a surface of the airfoil (<NUM>), with a leading edge (<NUM>) being as a reference (an upper wing surface direction is positive, and a lower wing surface direction is negative), and
c represents the chord length, wherein
the curvature is less than <NUM> at s/c=<NUM>, and the curvature increases to be <NUM> or more at s/c=<NUM>, and
a distribution of the curvature monotonically increases to <NUM> or more in a range from s/c=-<NUM> or less to a position of a trailing edge (<NUM>).