EP4649012A1

EP4649012A1 - Aircraft surface

Info

Publication number: EP4649012A1
Application number: EP24701527.4A
Authority: EP
Inventors: Marios KOTSONIS; Alberto Felipe RIUS VIDALES; Giulia ZOPPINI; Jordi CASACUBERTA PUIG; Theodoros MICHELIS; Sven Hendrikus Johannes WESTERBEEK
Original assignee: Technische Universiteit Delft
Current assignee: Technische Universiteit Delft
Priority date: 2023-01-14
Filing date: 2024-01-15
Publication date: 2025-11-19
Also published as: WO2024151166A1; NL2033960B1

Abstract

The invention relates to a swept wing of an aircraft including an aerodynamical wing surface and at least one elongated wing part integrally formed with or attached to the aerodynamical wing surface, the at least one elongated wing part extending in a longitudinal direction at an acute angle relative to a line parallel to a leading edge of the aerodynamical wing surface, wherein the at least one elongated wing part is configured to form a longitudinal protrusion on and/or a longitudinal depression in the aerodynamical wing surface respectively increasing or decreasing the local thickness of the swept wing. The at least one elongated wing part has a longitudinal leading edge and a longitudinal trailing edge and wherein at least one of the longitudinal leading edge and the longitudinal trailing edge of the at least one wing part has a curved shape profile.

Description

AIRCRAFT SURFACE

The present disclosure relates to a swept wing (for instance, a main (lifting) wing, a winglet, a horizontal and/or vertical tail, a control surface, or any other general lift-producing element) on an aircraft and to an aircraft comprising one or more of such swept wings.

Modern aircraft wings are swept backwards to delay the occurrence of a shock on the suction side of the wing. The characteristics of a boundary layer over the wing of the aircraft determines the lift, drag, and affects the efficiency of the vehicle in operation.

The airflow interacting with the leading edge of the aerodynamical wing surface of the swept wing forms a boundary layer at the wing surface. The boundary layer is defined as a thin region of airflow near the wing surface where the flow is dominated by effects of viscous nature. This boundary layer starts in a laminar state in which disturbances naturally develop. These disturbances amplify during the movement of the air over the wing surface and eventually lead to a transition of the boundary layer to a turbulent state.

Turbulence in the boundary layer airflows on the swept wing of an airplane is known to increase drag. A turbulent boundary layer airflow may therefore result in reduced efficiency in energy usage and thus fuel consumption of the aircraft. Flying under such conditions can result in a less economical, more expensive, and more polluting flight. An improved efficiency could be achieved if the boundary layer was completely laminar.

Controlling turbulence in a fluid flow has been a challenge in many fields of science and technology. In aerospace engineering, maintaining a laminar airflow on the aircraft wings is a known objective. The desire to reduce the drag of aircraft unites those seeking the expansion of economic margins and those aiming for a greener aviation industry.

Known methods and devices for turbulent flow control to reduce the negative effect of turbulence in the boundary layer of the airflow on swept wings include, for example, the application of small-sized riblets extending either parallel or almost parallel to the general airflow direction. However, these small-sized riblets are typically located in a turbulent boundary layer and work by inhibiting the spanwise turbulent motion and therefore breaking the turbulence autogeneration cycle. The overall technical effect of these small-sized riblets is the reduction of turbulent skin friction drag.

More generally, techniques applied to control a turbulent boundary layer on swept wings are referred to as turbulent flow control. These techniques aim at reducing the turbulent skinfriction drag produced by an already turbulent boundary layer. Active and passive techniques based on actuation or surface roughness manipulations have been developed for these applications as

SUBSTITUTE SHEET (RULE 26) well. These devices and methods are developed to deal with a relatively inefficient and non-ideal constraint, turbulence in the boundary layer flow, in the most efficient way possible.

An alternative approach to reducing the drag on a swept wing is by decreasing the portion of the wing of the aircraft covered by a turbulent boundary layer airflow. The present invention aims at delaying the boundary layer transition of laminar flow to turbulent flow. Delaying here means shifting the location at which the laminar flow transitions into a turbulent flow over the wing surface towards the trailing edge of the aircraft wing. In other words, delaying the boundary layer transition means reducing the areas of fully turbulent flow in the boundary layer. If the transition is delayed, the surface area on the wing where the flow is laminar is increased and the surface area where the flow is turbulent is decreased. Accordingly, the drag experienced by the aircraft wing may be reduced considerably.

Laminar flow control aims at delaying the laminar-to-turbulent transition of the boundary layer, obtaining a reduction of the aerodynamic drag by preventing or delaying the occurrence of turbulent flow. Methods and devices that aim to delay the boundary layer flow transition act in a different flow regime than turbulent flow control. Methods that aim to delay transition are applied to laminar boundary layers and affect the development of boundary layer instabilities. Methods that have been developed in this respect include active (such as blowing and suction) and passive techniques (such as surface roughness manipulations).

The transition of laminar to turbulent flow in the boundary layer on a swept wing is governed by a specific type of boundary layer flow component called Crossflow and an associated type of instability called Crossflow Instability (CFI). CFI develops as co-rotating vortices in the boundary layer near the wing surface. As these vortices grow and increase in amplitude, they eventually trigger the transition to turbulence. If the CFI is stabilized, then the laminar-to-turbulent transition can be delayed. Stabilization refers to either decay or inhibited growth of the amplitude of the instability (the instability generally being in the form of waves that are dampened out after interacting with the wing surface) as it propagates over the wing surface. Active techniques applied in these methods and devices involve complications. Namely, they require extra components and energy to operate and are more prone to wear and tear.

The current state of the art and industrial practice indicates that surface structures should be avoided or be kept as small as possible (typically with a height not larger than 0.2 times the local boundary layer thickness or even smaller and/or typically with a width of not more than one time the local boundary layer thickness) to delay transition to turbulence. Smooth and polished surfaces are currently the goal of the aerospace industry.

It is an object of the present disclosure to provide a swept wing of an aircraft wherein at least one of the above-mentioned disadvantages and inconveniences has at least partially been removed. It is a further object to provide a swept wing showing a reduced drag on the area on the aerodynamical wing surface and /or an increased efficiency in energy usage and thus fuel consumption of the aircraft.

SUMMARY

According to a first aspect a swept wing of an aircraft as claimed in appended claim 1 is provided. The swept wing may comprise an aerodynamical wing surface and at least one elongated wing part integrally formed with or attached to the, the at least one elongated wing part extending in a longitudinal direction at an acute angle relative to a line parallel to a leading edge of the aerodynamical wing surface, wherein the at least one elongated wing part is configured to form a longitudinal protrusion on and/or a longitudinal depression in the aerodynamical wing surface, thereby altering the local surface curvature and, consequently, increasing or decreasing the local thickness of the swept wing respectively; wherein the at least one elongated wing part has a longitudinal leading edge and a longitudinal trailing edge and wherein at least one of the longitudinal leading edge and the longitudinal trailing edge of the at least one wing part, or, preferably, both the longitudinal leading edge and the longitudinal trailing edge of the at least one wing part, has/have a curved shape, in particular a curved cross-sectional profile.

The term ‘wing’ is used in this document to refer to swept wings, either swept forwards and swept backwards, which can be any aerodynamical surface of the aircraft, including main wings providing the lifting force for the aircraft, winglets, (horizontal and/or vertical) tail parts, fins (vertical stabilizers, horizontal stabilizers (tailplanes), parts of an H-tail, dorsal fins, parts of a V-tail, parts of a T-tail, etc.). The curved shape, herein also referred to as smooth, elongated, or rounded shape, has been proven to allow for a smooth transition between the aerodynamical wing surface and the elongated wing part. The curved shape is defined by the cross section of at least one of the longitudinal edges of the wing part, wherein the cross-sectional shape, at the aerodynamical surface, is curving upwards, protruding from the wing, or curving downwards, forming a depression in the wing.

In a preferred embodiment the cross-section of at least one of the wing parts has a height that in operation is larger than 0.2 times the local boundary layer thickness and smaller than 5 times the local boundary layer thickness. The combination of the curved shape of the longitudinal leading edge, preferably also the longitudinal trailing edge, and the large size of the wing parts, has enabled the elongated wing part to at least partially stabilize boundary layer instabilities which are responsible for laminar-turbulent transition. The stabilization of the boundary layer instabilities makes it possible to delay the laminar-to-turbulent transition of the boundary layer flow. The height is determined by the distance between the regular aerodynamical surface of the wing without the wing part, and the protrusion or depression of the wing part from the regular aerodynamical surface. The local height of the wing part varies along the width of the wing part. Usually, when referring to the height of the wing part in general, the maximum height (depression or protrusion) is meant. The width of the wing part is defined by its extension along an imaginary axis from the leading edge to the trailing edge.

The local boundary layer thickness (d) or at least its order of magnitude can be determined by the following expression: with x the distance from the leading edge (LE) of the wing and U the flight velocity, specifically the cruise flight velocity of the aircraft.

In preferred embodiments of the present disclosure the acute angle of the wing part ranges from 0 to 45 degrees, preferably 5 to 40 degrees, more preferably 15-35 degrees with respect to a line parallel to the wing leading edge.

Additionally, the present device provides, in a preferred embodiment, at least one elongated wing part, wherein the cross-section of at least a portion of the wing part has a width, measured from the leading edge to trailing edge, that in operation is larger than 5 times the local boundary layer thickness and smaller than 300 times the local boundary layer thickness. The technical effects of such dimensions are the ease of fabrication and maintenance of the wing part, robustness and resilience to damages, debris, ware and tare, and insect strikes.

Depending on the boundary layer and constraints, optimization of the height and width of the wing part can maximize the stabilizing effect of the wing parts on the boundary layer. The boundary layer is a thin region of fluid in the vicinity of the solid aerodynamic surface of the wing in which the flow velocity is lower than the bulk velocity outside the boundary layer in the frame of reference of the aircraft or wing. The thickness of typical boundary layers encountered on modern transport aircraft can be in the range between 0.3 mm and 200 mm. Furthermore, the dimensions of the wing part are relatively large compared to known surface geometries for flow control on swept wings, which makes the wing parts described in the present disclosure relatively easy to produce, apply and maintain.

Throughout the disclosure, the aerodynamic surface of a swept wing may comprise a tail surface, a control surface, or any other lift-producing aerodynamic surface with varying functionality. Furthermore, a swept wing may comprise a main lifting wing, a winglet, a horizontal and/or vertical tail, control surfaces or any other general lift-producing element. DESCRIPTION EXEMPLIFYING EMBODIMENTS

In a first aspect, the present disclosure provides a swept wing of an aircraft, the swept wing comprising an aerodynamical wing surface and at least one elongated wing part integrally formed with or attached to the aerodynamical wing surface, the at least one elongated wing part extending in a longitudinal direction at an acute angle relative to a line parallel to a leading edge of the aerodynamical wing surface, wherein the at least one elongated wing part is configured to form a longitudinal protrusion on and/or a longitudinal depression in the aerodynamical wing surface respectively increasing or decreasing the local thickness of the swept wing. Furthermore, the at least one elongated wing part has a longitudinal leading edge and a longitudinal trailing edge. At least one of the longitudinal leading edge and the longitudinal trailing edge of the at least one wing part may have a curved shape in a cross-sectional profile, wherein only smooth geometries are present and sharp interfaces are avoided.

Embodiments could comprise wings with varying functionality such as main lifting wing, winglets, horizontal and vertical tails, control surfaces and any general lift-producing element.

The present disclosure poses an alternative for the passive delay of transition from laminar to turbulent flow in the boundary layer on swept wings. The device is intended for applications having laminar boundary layer flow subject to instabilities causing its transition to a turbulent boundary layer flow. The elongated wing part results in the stabilization of boundary layer instabilities, which are responsible for laminar to turbulent transition through the mechanisms described below. Stabilization of these instabilities eventually leads to transition delay and thus increasing the portion of the wing featuring laminar flow. In experiments, the inventors have demonstrated significant transition delay on a swept wing by using this device.

The advantage of using a smoothly shaped protrusion or depression is manyfold. Firstly, it is a passive device, which requires no complex or moving components and no energy to operate. It is therefore more robust and durable than active laminar flow control methods or devices. Secondly, this device extends the surface area on the swept wing on which the boundary layer is not turbulent. Instead of managing an already turbulent boundary layer and disadvantageous conditions, this device extends the range of the laminar boundary layer flow and thus increases advantageous conditions and corresponding effects. Furthermore, an elongated wing part with a smoothly shaped cross-sectional geometry has the advantage of being more efficient for laminar flow control than rectangular protrusions and depressions in swept wings. Furthermore, the elongated wing part is more robust and resilient to damage, debris, wear and tear, and insect strikes than smaller or rectangular protrusions and depressions. Furthermore, in practice the elongated wing part may be easier to maintain, clean and repair than smaller or rectangular protrusions and depressions. Common knowledge and intuition dictate that when the wing is not perfectly smooth, but comprises surface geometry or roughness features, the transition from laminar to turbulent flow in the boundary layer on the swept wing is advanced, and thus the portion of the boundary layer on the swept wing that is turbulent increases. The current state of the art and industrial practice indicates that surface structures should be avoided to delay transition to turbulence. Smoothly flat and polished surfaces maybe purposedly designed to achieve this effect. Hence these are currently the goals of the aerospace industry. The present disclosure describes surface geometry features of a swept wing that can be used for passive laminar flow control, with important potential impact on aerodynamic drag reduction of aircraft and subsequent fuel consumption.

For the sake of completeness, it is noted that the application of riblets and surface modifications are well-known for turbulent boundary layer drag reduction, as was mentioned before. However, the device according to the current disclosure is foremostly intended for applications involving laminar boundary layers, preferably specifically intended for application on the wing at a location where the boundary layer is usually laminar during operation, as opposed to turbulent. Additionally, the working mechanisms of these devices are fundamentally different, and the systems are not interchangeable, because laminar and turbulent boundary layers are fundamentally different and the device and method according to the current disclosure aims to delay the formation of a turbulent boundary layer.

An additional advantage of the device disclosed in the present application is the theoretical framework behind it. The main working principle of the stabilizing elongated wing part is identified as a wave interference phenomenon; this interference effect yields to a modification of the energy-transfer mechanisms of the flow and to the eventual decay or reduced growth of the energy of the instabilities that lead to transition. A linear interaction mechanism was found that results in a stabilization or reduced growth of the primary crossflow instability through energytransfer mechanisms. The shape and position of the device is critical for the success of laminarization. In some embodiments the device has smooth features and avoids any sharp interface for the stabilization of incoming crossflow instabilities by smooth surface geometries on the wing. In other embodiments the device may have non-smooth, for instance, sharp features. The working mechanisms of the device are known theoretically and predictable using modelling. This more fundamental understanding allows more efficient and case specific applications of the disclosed technique.

In preferred embodiments the elongated wing part is arranged on the aerodynamical wing surface, preferably on the upper surface of a horizontal wing. As mentioned before, a wing may be a main wing configured for proving the lifting force needed to keep the aircraft in the air. The upper and lower surface or upper and lower side of such main wing can be defined the suction side and the pressure side, respectively. More specifically, the main wings may be defined as the wings that provide for a lifting force (i.e. a lifting force that is sufficient to keep the aircraft in the air). The wing may also be a tailplane (i.e. a horizontal stabilizer) located on the tail (empennage) behind the main wings. A tailplane also generates a lifting force. In case the wing is comprised of a vertical tail part of the aircraft (i.e. the vertical stabilizer), this definition of suction side and pressure side may also apply although there may be conditions in which such tail parts do not produce such a lifting force. In preferred embodiments of the present disclosure, both the longitudinal leading edge and the longitudinal trailing edge of the at least one wing part are curved in the cross-sectional profile.

For example, in an embodiment wherein the swept wing comprises an elongated wing part configured to form an elongated protrusion on the aerodynamic surface of the wing, at the leading edge of the protrusion, the surface is curved upwards, and, at the trailing edge of the protrusion, it is curved downwards. When the elongated wing part is configured to form a depression on the surface, at the leading edge, the surface curves downwards and at the trailing edge, the surface curves upwards.

In these embodiments, sharp interfaces or surface geometries in the cross-sectional profile are generally avoided, which increases the beneficial effects of the elongated wing part on the stabilization of disturbances in the boundary layer flow.

In alternative embodiments, the cross-section of at least one of the wing parts has a height that in operation is preferably larger than 0.2 times the local boundary layer thickness and smaller than 5 times the local boundary layer thickness. Furthermore, in preferred embodiments, the crosssection of at least a portion of the wing part has a width, measured from its leading longitudinal edge to its trailing longitudinal edge, that in operation is larger than the local boundary layer thickness and smaller than 100 times the local boundary layer thickness.

The boundary layer is a thin region of fluid in the vicinity of the solid aerodynamic surface of the wing in which the flow velocity is lower than the bulk velocity outside the boundary layer. The thickness of typical boundary layers encountered on modern transport aircraft can be in the range between 50 micrometers and 200 mm.

In some preferred embodiments, the dimensions of the wing part are determined relative to the dimensions of the wing. In a preferred embodiment, the wing part has a height in the range of 10 micrometers to 200 millimeter, preferably 20 micrometers to 100 millimeters, most preferably 30 micrometers to 50 millimeters. For example, between 30 to 60 micrometer, or for example between 60 to 600 micrometers, or for example between 0.6 and 1 millimeter, or for example between 1 and 100 millimeters, or for example between 100 and 200 millimeters. In these preferred embodiments, the width of the wing part is in the range of 0.5 mm to 1 meter, preferably 1 millimeters to 500 millimeters, and most preferably 5 millimeters to 200 millimeters.

In embodiments of the present disclosure the height of the wing part is based on the relative chordwise location defined as the distance from the leading edge of the wing as percentage of the total wing width. The total wing width may be defined as the distance between the leading edge and trailing edge of the wing. The height therefore may vary according to the chordwise location. Generally, the height may first increase from the leading edge towards the trailing edge of the wing and then be reduced again.

In specific embodiments the wing part in a relative chord wise location range of 0-5%, has a height center value of about 0.6 mm, in a relative chordwise location range of 5-10% has a height center value of about 1 mm, and in a relative chordwise location range of 10-20% has a height center value of about 2 mm. More specifically, in further embodiments, the wing part in a relative chordwise location range of 20-30% has a height center value of about 3 mm, in a relative chordwise location range of 30-40% has a height center value of about 3.8 mm, in a relative chordwise location range of 40-50% has a height center value of about 4.1 mm, in a relative chordwise location range of 50-60% has a height center value of about 5 mm, and/or in a relative chordwise location range of 60-100% has a height center value of about 8 mm.

The actual height of the swept of the wing part may vary in a range between 0.2-2 times any of the above-mentioned height center values.

It is to be understood that within the above-identified ranges the dimensions of the shape of the wing part of course may vary across the wing, for instance vary in the spanwise direction.

In other preferred embodiment, one dimension of the wing part, such as the height or width, is defined relative to the wing or relative to the boundary layer, and another dimension of the wing part, such as the width or height, is defined in absolute measurements.

Depending on the boundary layer and constraints, an optimization of the height and width of the wing part can maximize the stabilizing effect of the wing parts on the boundary layer. Furthermore, the dimensions of the elongated wing part are relatively large compared to known surface geometries for flow control on swept wings, which makes these wing parts relatively easy to produce, apply and maintain.

In some embodiments, the longitudinal leading edge and longitudinal trailing edge of the at least one elongated wing part on the wing extend at a substantially constant angle relative to a line parallel to the leading edge of the aerodynamical wing surface.

In such embodiments, the leading and trailing edge of the elongated wing part extend at a constant angle with a line parallel to the leading edge of the wing. In other words, the longitudinal extension of the elongated wing part is not curved, but straight (at one angle) with respect to the leading edge of the wing.

These surface geometries on the wing could prove to be sufficient for significant extension of the laminar flow in the boundary layer on the swept wing, whilst requiring a relatively simple modification of the common aircraft wings. The use of a straight elongated wing parts is easy in regard to fabrication by common manufacturing methods.

In other embodiments, the longitudinal extension of the elongated wing part on the wing is curved. For example, as the elongated wing part extends along the wing in longitudinal direction, it might bend away from the leading edge of the wing, creating an acute angle between the longitudinal extension of the edge of the elongated wing part and a line parallel to the edge of the wing.

For example, the swept wing might comprise an elongated wing part, whose trailing edge curves towards the trailing edge of the swept wing, as the elongated wing part extends along a longitudinal direction on the swept wing. Similarly, the leading edge of the elongated wing part might curve towards the leading edge of the wing, as the elongated wing part extends along a longitudinal direction on the swept wing, making the elongated wing part wider as it extends. Alternatively, the leading and/or trailing edge of the elongated wing part might curve ‘inwards’; the leading edge might bend towards the trailing edge of the swept wing, and/or the trailing edge of the elongated wing part might bend towards the leading edge of the swept wing. When both the leading and trailing edge of the elongated wing part are curved, they might bend in the same direction or in opposite directions. Additionally, their curvatures might be evenly strong, or one of the leading edge and the trailing edge of the elongated wing part might have a stronger curvature than the other.

By providing a wing part with a continuous curve of one or both of the longitudinal leading edge and the longitudinal trailing edge, the effect of the elongated wing part on the characteristics of the flow in the boundary layer can be increased. In particular, these may be optimized for local characteristics and non-uniformities of the boundary layer. Furthermore, a curved longitudinal extension of the elongated wing part can be more effective on swept wings which have a curved leading edge or curved trailing edge or both. Furthermore, a curved longitudinal extension of the elongated wing part can be more effective on swept wings which have a non-constant geometric angle of attack, or non-constant chord or non-constant dihedral angle or any combination of the aforementioned features.

In preferred embodiments, the swept wing comprises a plurality of elongated wing parts, wherein the elongated wing parts preferably extend substantially parallel to each other. For example, the swept wing might comprise two or more elongated wing parts extending in the longitudinal direction of the wing, comprising a depression or a protrusion in the wing.

By adding multiple elongated parts to the wing, the region of the wing featuring a laminar boundary layer flow can be extended even further, moving the transition location of the boundary layer further towards the wing’s trailing edge. Consecutive, e.g. as a series of, protrusions or depressions in the aerodynamical wing surface can consecutively cause a stabilization of the flow in the boundary layer. Since the stabilizing effect is repeated as the flow progresses on the aerodynamical wing surface, the transition from laminar to turbulent is even further delayed.

In a preferred embodiment, the swept wing comprises a plurality of substantially parallel wing parts, wherein the wing parts are spaced apart over a distance between 1 and 10 times, preferably between 2 and 8 times, the width of at least one of the wing parts.

For example, the swept wing might comprise two or more elongated wing parts spaced apart by 5, or anything between 2 and 10, times the width of the elongated wing part. When the swept wing comprises more than two elongated wing parts, they might be spaced equally or irregularly.

By adding multiple elongated parts to the wing, the transition from laminar to turbulent flow in the boundary layer can be extended even further. Two elongated wing parts cannot be positioned too close together, or the stabilizing effect will be reduced or diminished, since it takes some space for the amplitude of the CFI to reach its minimum after interaction with the surface geometry. Additionally, the protrusions and/or depressions should not be positioned too far apart, since that will result in the increased development of CFI in the boundary layer, and, consequently, the transition into turbulent flow. The correct spacing of elongated wing parts according to the present disclosure can result in the subsequent extension of the laminar flow regime, at every other elongated wing part, in the boundary layer on the swept wing during operation.

Swept wing as claimed in any of the preceding claims, wherein the elongated wing part has a smooth and continuous shape in cross-section.

In embodiments of the present disclosure, the elongated wing part comprises only curved edges and no sharp interfaces or rectangular shape geometries.

In some embodiments, the elongated wing part has a cross-sectional shape that is essentially symmetric relative to a central imaginary line of symmetry extending in longitudinal direction.

For example, when the swept wing is viewed at its cross section, the surface geometry of the elongated part might be symmetrical with respect to the imaginary line that extends from the center of the wing in cross section through the maximum or minimum of the shape geometry of the elongated wing part. In this way, a significant effect can be achieved using a relatively simple wing part that requires a relatively easy modification of the common aircraft wings. Furthermore, the fabrication of a symmetric elongated wing part is more straightforward with common manufacturing techniques.

In other embodiments, the elongated wing part has a cross-sectional shape that is asymmetric relative to a central imaginary line of symmetry extending in longitudinal direction.

For example, the swept wing might comprise at least one elongated wing part, wherein the elongated wing part has a cross-sectional shape that rises faster than it lowers (in the direction of the airflow over the aerodynamical wing surface) or other way around. When one or more elongated wing parts comprise a depression, the elongated wing part could have a cross-sectional shape that drops faster than it rises.

Using an asymmetric protrusion or depression in one or more elongated parts of the swept wing in the right position, the beneficial effect of the elongated wing part on the development in CFI in the boundary layer can be increased. The increase of beneficial effect is derived by the optimization of the energy exchange mechanisms earlier identified in the theoretical framework of the disclosure. The asymmetry can be used to optimally suppress the development of boundary layer instabilities over the entire region of interaction with the wing part.

In a preferred embodiment, the swept wing comprises at least one elongated wing part wherein at least one of the end surfaces of the elongated wing part, in the direction of the flow, preferably both end surfaces of the elongated wing part, is configured to smoothly converge to an original shape of the swept wing. An original shape of the swept wing is a shape of the wing without a wing part according to the present disclosure. In other words, the wing part converges to the remainder of the swept wing, the part of the swept wing outside the wing part.

One advantage of using elongated wing parts that converge into the original shape of the swept wing is that existing wings can be modified to include such wing parts in the application. Furthermore, the operation of converging wing parts prevents the unwanted presence of sharp edges and geometries in the application of this technique, which could reduce its stabilizing effect on the boundary layer flow significantly.

In some embodiments, the swept wing of the present disclosure comprises one or more elongated wing parts wherein at least one of the end surfaces of the elongated wing part is tapered.

Swept wing as claimed in any of the preceding claims, wherein the cross-sectional shape geometry of the one or more elongated wing parts is defined by an arbitrary algebraic polynomial of the form: where:

0 < x < w : independent x coordinate w : max width of the elongated wing part (measured in the direction substantially perpendicular to the leading edge of the wing) y : dependent y coordinate n : polynomial order cik G R : polynomial coefficients

For example, the cross section of the elongated wing part might be described with a sixth order polynomial.

The advantage of using this shape geometry is that it allows for a large variety of smooth protrusions and depressions. Furthermore, the advantage of this shape geometry is that it is mathematically always smooth and differentiable, facilitating ease of programming in automated manufacturing processes.

In some embodiments, if there is more than one elongated wing part on the wing, the cross-sectional shape geometry of the at least two of the elongated wing parts differ.

For example, the swept wing might comprise one elongated wing part that has the cross- sectional shape geometry of a fifth order polynomial, and one elongated wing part with the cross- sectional shape geometry of a fourth order polynomial. Furthermore, at least one of the elongated wing parts might be a protrusion on the wing, whilst at least one other elongated wing part is a depression in the wing.

The use of different shaped wing parts in a wing can allow the optimization of each individual wing part to the local characteristics of the boundary layer interacting with each wing part, contributing to the stabilizing effect of the wing parts on the boundary layer.

In further embodiments, a swept wing elongated wing part has a smooth and continuous shape in longitudinal section.

In further embodiments, the cross section of one or more wing parts has a varying width and/or a varying thickness/height along the longitudinal direction.

In further embodiments, the shape geometry of the cross section of at least one elongated wing part varies along the longitudinal direction of the elongated wing part.

In further embodiments, the one or more wing parts of the swept wing are located in a laminar and/or transitioning boundary layer of the airflow during operation.

During operation of the aircraft, the airflows around the swept wing and interacts with the wing and a boundary layer flow forms around the wing of the aircraft.

By positioning the one or more elongated wing parts on the area of the swept wing that corresponds to the area of the laminar flow regime in the boundary layer on the swept wing during operation, the provided elongated wing parts affect the development of CFI in the boundary layer and delay the transition of laminar to turbulent flow. This position is essential for the successful application of the swept wing with elongated wing parts.

In some embodiments, the elongated wing part is oriented substantially perpendicular to the local general flow direction of air flowing along the wing surface when the swept wing of the aircraft is in operation.

During operation, the air flows around the swept wing and interacts with the wing and a boundary layer of the flow around the wing of the aircraft is formed.

By positioning the elongated wing part on the wing such that the longitudinal extension of the wing part is substantially perpendicular to the local general flow direction of air flowing along the wing surface, the stabilizing effect of the elongated wing part is increased.

The wing part as disclosed may be applied in swept wings of an aircraft.

By providing one swept wing or, preferably, a plurality of swept wings, with one or more elongated wing parts as described in the present disclosure, the above-discussed effect is increased, and the efficiency of the aircraft is improved.

BRIEF DESCRIPTION OF DRAWINGS

Figure 1 shows one swept wing of an aircraft.

Figure 2 shows the boundary layer of the airflow on the swept wing of an aircraft in operation.

Figure 3 shows an embodiment of a swept wing with an elongated wing part on the swept wing of an aircraft.

Figure 4 shows the effect of the elongated wing part on the swept wing on the laminar and turbulent flow regimes in the boundary layer on the swept wing.

Figure 5 shows cross-sectional shape geometries of the swept wing with at least one elongated wing part in some possible embodiments.

Figure 6 shows the possible cross-sectional shape geometries of elongated wing parts.

Figure 7 shows some possible embodiments of swept wings with multiple elongated wing parts from a perspective.

DETAILED DESCRIPTION

In figure 1, an example of a swept wing 1 of an aircraft 2 is illustrated. The wing is tilted backward slightly from its root, rather than extending in a straight sideways direction. The aircraft wing is illustrated as in operation, which means that it is moving through air. Specifically, in this example the aircraft is flying. The phrasing “in operation” is used for any situation wherein the aircraft is in moving in a forward direction through air; and thus also refers to situations where the aircraft is moving on the ground. However, “in operation” is applied mainly to refer to flying aircraft.

In figure 2 the swept wing of an aircraft is shown in detail. The wing of the aircraft is illustrated as when flying. In the figure, the boundary layer 3 of the airflow around the swept wing is shown for one cross section of the wing. In the boundary layer, the different flow regimes are shown. The boundary layer starts in a laminar state 4 on which disturbances 5 and 7 naturally develop. These disturbances amplify during the movement of the air over the wing surface and eventually lead to a transition of the boundary layer to a turbulent state 6.

The disturbances are indicated with lines within the boundary layer 3. A separate frame inside the figure shows the transition 5 from the laminar to turbulent flow in more detail. The transition of laminar to turbulent flow in the boundary layer on a swept wing is governed by a specific type of boundary layer flow component called crossflow and an associated type of instability called Crossflow Instability (CFI). CFI develops as co-rotating streamwise aligned vortices developing in the boundary layer. As these vortices grow, they increase in amplitude and eventually break down to turbulence 8. In the figure the turbulence is indicated with swirl-shaped lines 8.

In Figures 3A and 3B (not to scale), an example of an embodiment of the present disclosure is shown.

Figure 3A shows an aircraft with a wing comprising one elongated wing part 10 from a perspective. The elongated wing part 10 is attached to the aerodynamical wing surface 9 and extends on the aerodynamic surface of the wing in a longitudinal direction, substantially parallel to the leading edge of the wing 20. In operation, the direction of the longitudinal extension of the elongated wing part 10 will be substantially perpendicular to the direction of the flow 22. The elongated wing part is configured to form a longitudinal protrusion 17 and locally increases the local thickness of the swept wing 1. The longitudinal extension direction 13 is indicated with an arrow.

In the figure, the elongated wing part has a longitudinal leading edge 11 and a longitudinal trailing edge 12. In some embodiments, at least one of the longitudinal leading edge 11 and the longitudinal trailing edge 12 of the wing part has a curved shape in cross-sectional profile. At the leading edge 11 of the elongated protrusion, the surface curves upwards, and, at the trailing edge of the elongated protrusion the surface curves back into the original shape of the plane. In figure 3B, the same embodiment as in Figure 3A is shown, from a cross-sectional perspective of the wing. The aircraft is shown behind the cross section of the wing. In orientation shown in the figure, the leading edge of the wing is on the left, the trailing edge of the wing is on the right side of the wing. In operation, the direction of the flow 22 will be from left to right, from the leading edge of the wing to the trailing edge of the wing. The flow direction will be more or less perpendicular to the longitudinal direction 15 of the elongated wing part. The cross-sectional inside of the wing 23 with elongated wing part is indicated with dashed lines. The elongated wing part is configured to form a longitudinal protrusion 17 and locally increases the local thickness of the swept wing. The longitudinal extension of the elongated wing part is not visible in this figure since the longitudinal direction 15 is perpendicular to the cross-section of the elongated wing part. The position of the longitudinal leading edge 11 in the cross section of the elongated wing part on the wing is indicated with an arrow on the left of the protrusion, and the position of the longitudinal trailing edge 12 in the cross section of the elongated wing part on the wing is indicated with an arrow on the right of the protrusion.

Figures 3A and 3B show a minimal embodiment of the present disclosure, comprising only one elongated wing part. Other embodiments can comprise multiple elongated wing parts. Furthermore, in other embodiments, the elongated wing parts could be integrally formed with or attached to the aerodynamical wing surface and configured to form a longitudinal protrusion on and/or a longitudinal depression in the aerodynamical wing surface respectively increasing or decreasing the local thickness of the swept wing.

Figures 4A and 4B show some plotted results from the model of a possible embodiment of the present disclosure.

In Figure 4 A, the flow regimes of the boundary layers are shown for different embodiments of an aircraft wing in operation. The upper plot on the left shows the range of the laminar 4 and the turbulent 6 flow regimes on the aerodynamic surface a swept wing according to the state of the art. The lower plot on the left shows the range of the laminar and the turbulent flow regimes on the aerodynamic surface a swept wing according to the present disclosure. The plot in Figure 4 A on the left shows the difference, by means of subtraction, between the two situations.

In Figure 4B contains a plot showing the development of the amplitude of the disturbances 7 on the two types of wings; the swept wing according to the state of the art and the swept wing according to the present disclosure. The development of the amplitude of the disturbances as a function of the progression of the flow on the aerodynamic surface of the wing according to the state of the art and the wing according to the present disclosure is illustrated in the plot. These are the result of a model describing current swept wing as available. According to the model, the use of a swept wing according to the present disclosure shows a significant reduction of the development of the amplification of the disturbances in the boundary layer flow.

Figure 5 shows some possible types of embodiments of the elongated wing part from a cross section. Figure 5A shows an embodiment of a wing of an aircraft according to the present disclosure, comprising an elongated wing part attached to the aerodynamical wing surface and configured to form a longitudinal protrusion 17 on the aerodynamical wing surface, increasing the local thickness of the swept wing.

Figure 5B shows an embodiment of a wing of an aircraft according to the present disclosure, comprising an elongated wing part integrally formed with the aerodynamical wing surface and configured to form a longitudinal depression 16 in the aerodynamical wing surface, decreasing the local thickness of the swept wing.

Figure 5C shows an embodiment of a wing of an aircraft according to the present disclosure, comprising two elongated wing parts from a cross section. The first elongated wing part forms a protrusion 17 on the aircraft wing, the second elongated wing part forms a depression 16 in the wing.

Other possible embodiments can also comprise multiple elongated protrusions or multiple elongated depressions or a combination thereof.

Figure 6 shows some possible shape geometries of the cross section of the elongated wing parts of a few of the possible embodiments of the present disclosure. In the upper row of the table, the cross-sectional shape geometries of some possible elongated wing parts configured to form a protrusion on the wing are shown. On the left, a symmetric cross-sectional shape geometry of the elongated protrusion of a possible embodiment is shown. On the right, two asymmetric cross- sectional shape geometries of elongated protrusions of possible embodiments are shown. The first asymmetrical protrusion 152 first rises relatively slowly and then drops steeper than it rises. The second asymmetrical protrusion 153 rises more steeply than it descends.

In the lower row of the table, the cross-sectional shape geometries of some possible elongated wing parts configured to form a depression on the wing are shown. On the left, a symmetric cross-sectional shape geometry of the elongated depression 161 of a possible embodiment is shown. On the right, two asymmetric cross-sectional shape geometries of elongated depressions of possible embodiments are shown. The first asymmetrical depression 162 first drops relatively slowly and then rises steeper than it rose. The second asymmetrical depression 163 lowers more steeply than it rose.

Figures 7A-7D show some possible embodiments of the present disclosure from a perspective of a cross section of the swept wing. The wing 1 is indicated with thick continuous lines. On the wing, the dash-dotted lines indicate the ‘imaginary’ lines parallel to the leading edge of the swept wing 201. The inside of the wing 130 with elongated wing parts is shown in cross section. The elongated wing parts 10 are shown to increase or decrease the local thickness of the wing. The longitudinal direction 15 of the elongated wing parts is indicated with a dashed line. In the figures, multiple possible embodiments are shown, comprising curved as well as straight longitudinal extensions of protrusions/depressions with respect to the leading edge of the wing, and parallel configurations of elongated wing parts as well as non-parallel configurations of multiple elongated wing parts on the wing. The elongated wing part may be curved in the generally longitudinal direction 15 as well, i.e. in plan view rather than in cross-sectional view. This means that the leading edge 11 or the trailing edge 12 of the elongated wing part or both can bend either towards the leading edge of the wing 20 or towards the trailing edge of the wing 21, making at acute angle a with a line parallel to the leading edge of the wing. In a preferred embodiment, the acute angle is in the range of 0 to 45 degrees, preferably in a range of 5 to 40 degrees, more preferably in a range of 15 to 35 degrees. In an embodiment, the acute angle is in the range of 0 to 25 degrees with respect to the leading edge of the wing.

Although the present invention has been described above with reference to specific embodiments, it is not intended to be limited to the specific form set forth herein. Rather, the invention is limited only by the accompanying claims and other embodiments than the specific above are equally possible within the scope of these appended claims.

Furthermore, although exemplary embodiments have been described above in some exemplary combination of components and/or functions, it should be appreciated that, alternative embodiments may be provided by different combinations of members and/or functions without departing from the scope of the present disclosure. In addition, it is specifically contemplated that a particular feature described, either individually or as part of an embodiment, can be combined with other individually described features, or parts of other embodiments.

Claims

1. Swept wing of an aircraft, the swept wing comprising an aerodynamical wing surface and at least one elongated wing part integrally formed with or attached to the aerodynamical wing surface configured to decrease drag on the aerodynamical wing surface, the at least one elongated wing part extending in a longitudinal direction at an acute angle relative to a line parallel to a leading edge of the aerodynamical wing surface, wherein the at least one elongated wing part is configured to form a longitudinal protrusion on and/or a longitudinal depression in the aerodynamical wing surface respectively increasing or decreasing the local thickness of the swept wing; wherein the at least one elongated wing part has a longitudinal leading edge and a longitudinal trailing edge and wherein at least one of the longitudinal leading edge and the longitudinal trailing edge of the at least one wing part has a curved shape in cross-section, wherein preferably the at least one elongated wing part is arranged at the top side of the swept wing.

2. Swept wing as claimed in claim 1, wherein both the longitudinal leading edge and the longitudinal trailing edge of the at least one wing part are curved.

3. Swept wing as claimed in any of the preceding claims, wherein the cross-section of at least one of the wing parts has a height that in operation is larger than 0.2 times the local boundary layer thickness and smaller than 5 times the local boundary layer thickness.

4. Swept wing as claimed in any of the preceding claims, wherein the cross-section of at least a portion of the wing part has a width, measured from its leading longitudinal edge to its trailing longitudinal edge, that in operation is larger than 5 times the local boundary layer thickness and smaller than 300 times the local boundary layer thickness.

5. Swept wing as claimed in any of the preceding claims, wherein the longitudinal leading edge and a longitudinal trailing edge of the at least one elongated wing part extend at a substantially constant angle relative to a line parallel to the leading edge of the aerodynamical wing surface.

6. Swept wing as claimed in any of the preceding claims, wherein the swept wing comprises a plurality of elongated wing parts, wherein the elongated wing parts preferably extend substantially parallel to each other.

7. Swept wing as claimed in claim 6, wherein the wing parts are spaced apart over a distance between 1 and 10 times, preferably between 2 and 8 times, the width of at least one of the wing parts.

8. Swept wing as claimed in any of the preceding claims, wherein the elongated wing part has a smooth and continuous shape in cross-section.

9. Swept wing as claimed in any of the preceding claims, wherein the elongated wing part has a cross-sectional shape that is essentially symmetric relative to a central imaginary line of symmetry extending in longitudinal direction.

10. Swept wing as claimed in any of the claims 1-8, wherein the elongated wing part has a cross-sectional shape that is asymmetric relative to a central imaginary line of symmetry extending in longitudinal direction.

11. Swept wing as claimed in any of the preceding claims, wherein at least one of the end surfaces of the elongated wing part, preferably both end surfaces of the elongated wing part, is configured to smoothly converge to the original shape of the swept wing.

12. Swept wing as claimed in any of the preceding claims, wherein at least one of the end surfaces of the elongated wing part is tapered.

13. Swept wing as claimed in any of the preceding claims, wherein the cross-sectional shape geometry of the one or more elongated wing parts is defined by an arbitrary algebraic polynomial of the form: where:

0 < x < w : independent x coordinate w : maximum width of the elongated wing part (measured in the direction substantially perpendicular to the leading edge of the wing) y : dependent y coordinate n : polynomial order a_k £ i : polynomial coefficients

14. Swept wing as claimed in any of the preceding claims, wherein, if there is more than one elongated wing part, the cross-sectional shape geometry of at least two of the elongated wing parts differ.

15. Swept wing as claimed in any of the preceding claims, wherein the elongated wing part has a smooth and continuous shape in longitudinal section.

16. Swept wing as claimed in any of the preceding claims, wherein the cross section of one or more wing parts has a varying width and/or a varying thickness along the longitudinal direction.

17. Swept wing as claimed in any of the preceding claims, wherein the shape geometry of the cross section of at least one elongated wing part varies along the longitudinal direction of the elongated wing part.

18. Swept wing as claimed in any of the preceding claims, wherein the one or more wing parts of the swept wing are located in a laminar and/or transitioning boundary layer of the airflow during operation.

19. Swept wing as claimed in any of the preceding claims, wherein the elongated wing part is oriented substantially perpendicular to the local general flow direction of air flowing along the wing surface when the swept wing of the aircraft is in operation.

20. Swept wing as claimed in any of the preceding claims, wherein the swept wing is a swept forward wing or, preferably, a swept backward wing.

21. Swept wing as claimed in any of the preceding claims, wherein the acute angle is in a range of 0 to 45 degrees, preferably in a range of 5 to 40 degrees, more preferably in a range of 15 to 35 degrees.

22. Swept wing as claimed in any of the preceding claims, wherein the swept wing is a main lifting wing, a winglet, a horizontal and/or vertical tail, a control surface, or any other general lift-producing element.

23. Swept wing as claimed in any of the preceding claims, wherein a local boundary layer thickness (d) is in the order of magnitude as: with x the distance from the leading edge (LE) of the wing and U the flight velocity, specifically the cruise flight velocity of the aircraft.

24. Swept wing as claimed in any of the preceding claims, wherein the wing part has a height in the range of 10 micrometers to 200 millimeter, preferably 20 micrometers to 100 millimeters, most preferably 30 micrometers to 50 millimeters.

25. Swept wing as claimed in any of the preceding claims, wherein the height of the wing part is based on the relative chordwise location defined as the distance from the leading edge of the wing as percentage of the total wing width.

26. Swept wing as claimed in claim 25, wherein the at least one elongated wing part in a relative chordwise location range of 0-5%, has a height center value of about 0.6 mm, in a relative chordwise location range of 5-10% has a height center value of about 1 mm, and in a relative chordwise location range of 10-20% has a height center value of about 2 mm.

27. Swept wind as claimed in claim 26, wherein the at least one elongated wing part in a relative chordwise location range of 20-30% has a height center value of about 3 mm, in a relative chordwise location range of 30-40% has a height center value of about 3.8 mm, in a relative chordwise location range of 40-50% has a height center value of about 4.1 mm, in a relative chordwise location range of 50-60% has a height center value of about 5 mm, and/or in a relative chordwise location range of 60-100% has a height center value of about 8 mm.

28. Swept wind as claimed in claim 26 or 27, wherein the actual height of the wing part may vary in a range between 0.2-2 times the height center value.

29. An aircraft comprising one or more swept wings as claimed in any of the preceding claims.