WO2013083570A1

WO2013083570A1 - Airplane propeller system acting on the wingtip vortex

Info

Publication number: WO2013083570A1
Application number: PCT/EP2012/074376
Authority: WO
Inventors: Cesar Raul LOPEZ ORBEA; Norberto Manfredo SCHNEIDER
Original assignee: Lopez Orbea Cesar Raul; Schneider Norberto Manfredo
Priority date: 2011-12-06
Filing date: 2012-12-04
Publication date: 2013-06-13
Also published as: AR088412A1

Abstract

An aircraft propulsion system acting on the wingtip vortex consisting of a fixed shaft (2) placed at the end chord or wingtip of each of the wings of the aircraft and a plurality of propellers (4,5,6) mounted on each of the shafts (2). Both shafts are driven by a single central engine (8) or a single gear box (10). Each wing shows a slot or cutout portion (11) that allows at least one central wingtip propeller (5) to freely rotate through the wing to enhance the vortex cancellation effect.

Description

AIRPLANE PROPELLER SYSTEM ACTING ON THE WINGTIP VORTEX

CROSS REFERENCE TO RELATED APPLICATIONS

[0001] This application claims the benefits of patent application AR Pl l 01 04554 filed on

December 6, 201 1 , by the present inventors and the application is hereby incorporated by reference in its entirety.

STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT

[0002] Not Applicable NAMES OF PARTIES TO JOINT RESEARCH AGREEMENT

[0003] Not Applicable

REFERENCE TO SEQUENCE LISTING

[0004] Not Applicable

DESCRIPTION OF ATTACHED APPENDIX [0005] Not Applicable

BACKGROUND OF THE INVENTION 1. Field of the Invention

[0006] The field of the invention is the one related to the aircraft design and more precisely to the design of a new propeller design that acts on the wingtip vortex. 2. Description of Related Art

[0007] It is well known by the aeronautic engineers that a moving wing generates uplift because its shape (profile) produces an overpressure at the lower part (intrados) and underpressure at the upper part (extrados) (in the NACA profiles, this is done in an approximate proportion of 5: 1 , that is, the wing is sucked 5 times stronger from the upper side than being pushed from the lower side). Figure 1 shows a front view of a wing, wherein the + signs indicate overpressure and the - signs indicate underpressure. Due to the fact that the wing is finite, at the wingtip (or tip chord) a strong pressure gradient takes place, generating the well known wingtip vortex (see figure 1), which is responsible of generating the part of the drag to the forward movement called "induced drag", being approximately 25% of the total drag. This induced drag also adds other problems; for example, at the leading edge it adds an ascending vertical component to the incident speed, by which the effective attack angle is greater than the apparent one, which not only increases the forward movement drag, but also drives the wing closer to the stall angle. The mentioned issue is well known since the first aircraft was made to fly, and several solutions were attempted:

[0008] 1.- The most basic one is to construct the wingtip shaped as a cone quarter (see fig. 2)

[0009] 2.- The above mentioned solution was improved by using the "winglet" (see fig. 3).

[0010] Several documents are known in the art that attempt to solve the problem of the wingtip vortex.

[0011] Document US 2,485,218 discloses a rigid wing of an aircraft comprising a rotor positioned with its rotating axis substantially parallel with the direction of flight and forming a wing tip, the rotor having circular cross-section and a longitudinal contour corresponding substantially to that of the cross-sectional shape of the airfoil at its outer end, bearing means for the rotor mounted in the wing and means for driving the rotor in the direction from the low pressure side to the high pressure side of the wing. However, this long overdue design does not show the convenient novel features of the present invention because the rotor has proved to be a quite weak vortex destroying shape, compared to a propeller. Moreover, the cited document discloses the use of an electric engine or other motor 18 near each of the wing tips (column 1 , lines 42 - 43), which is quite inconvenient for the following reasons: a) the motors should not be placed near the wing tips to avoid an flexural overstress on the cantilever mounted wings' structure and b) two separate power engines should not be used to drive both vortex- destroying elements to avoid that the failure of one of the motors severely unbalances the aircraft during flight. This risk is even higher it the power engines are mounted at the wingtips due to the great torque unbalance that may result, urging the aircraft to rotate in a horizontal plane around its central vertical axis (yaw).

[0012] Document US 4,533,101 discloses a device that seeks to increase the efficiency of aircrafts based on a wingtip pusher propeller 12 positioned behind the wingtip 18 to rotate in the crossflow of the wingtip vortex. The propeller 12 rotates against the vortex swirl in order to create additional thrust from the wingtip vortex and attenuating it when extracting its energy to convert it into propeller movement. However, the cited document discloses the use of an engine nacelle 20 positioned at the wingtip 18 (column 3, lines 36 - 37), which is quite inconvenient for the following reasons: a) the motors should not be placed near the wing tips to avoid an flexural overstress on the cantilever mounted wings' structure and b) two separate power engines should not be used to drive both vortex-destroying propellers to avoid that the failure of one of the motors unbalances the aircraft during flight. Also, the document mentions in column 2, lines 35 - 41 and in claims 7 and 8 that the propeller is positioned aft of the trailing edge to operate in the crossflow of the wingtip vortex, and in detail aft of the engine nacelle. This position, however, is not the best placing of the propellers for cancelling the vortex. Another difference of the cited document compared to the present invention is that the document defines in claim 1 that the propeller is designed to induce a wingtip vortex having a core and a crossflow and for achieving this, claim 4 defines very particular blade diameters and that the propeller is positioned right inside the crossflow zone. However, in the present invention the propellers, and particularly those mounted in the intermediate positions of the end chord, are designed to take advantage of the core vortex but not a crossflow, because the vortex intensity in the core is maximum and this will achieve a maximum traction. Also, in the cited document the shaft mounted at the wing tip is fixed, by which it is not possible to change the pitch angle and this poses a strong limitation to the device's adjustment settings to meet the vortex cancellation requirements.

[0013] Document US 4. ,917,332 discloses a means for extracting rotational energy from the vortex created at aircraft wing tips which consists of a turbine with blades located in the crossflow of the vortex and attached downstream of the wingtip. The turbine 30 has blades 40, 41 , 42 and 43 attached to a core 45. When the aircraft is in motion, rotation of core 45 transmits energy to a centrally attached shaft 50. The rotational energy thus generated may be put to use within the airfoil 20 or aircraft fuselage 10. However, unlike the present invention, the element designed for cancelling the wingtip vortex of the cited document is not driven by any engine but, instead, is designed to supply mechanical energy to an electrical generator or the like.

[0014] Document US 5, 100,085 describes a rotor which is mounted to an aircraft wing to recover induced drag associated with a wingtip vortex. The document indicates that, when the rotor is placed in the vortex stream, the transverse component of relative wind encountered at selected increasing spanwise locations along the blade changes sign due to an increasing transverse component of blade velocity and decreasing transverse component of vortex velocity with decreasing distance from the axis of rotation. The document also describes that, in order to maximize induced drag recovery, the blade is twisted in a spanwise direction so that the inner portion of the blade drives the blade, while the radially outward portion acts as a propeller to resist rotation and, the induced drag recovery is in the form of thrust generated by the rotor blade. In this way the inner blade portions next to the hub urge the propeller to rotate, while the outer portions of the blades next to the blade tips generate a traction force. However, unlike the present invention, the cited document uses a free rotating propeller with no active engine to drive its rotational movement.

[0015] Document US 5,918,835 discloses a wingtip vortex device installed at the wingtips of an aircraft for induced drag reduction and vortex cancellation formed by a spinner 52 having a number of radial fins 54 coupled to a shaft. . However, the document describes that the device is self-contained and powered by the wingtip vortex but is not powered by any engine whatsoever. Also, unlike the present invention, figs. 1 - 3 of the cited document show that the spinner 54 is not placed right at the wingtip 30 but displaced inwardly so as to be rotated by the descending portion of the vortex. This shifted placement would not be effective with the vortex cancelling propeller of the present invention.

[0016] Document US 5,934,612 discloses a wingtip vortex device for induced drag reduction and vortex cancellation, to be installed at the aircraft wingtips and comprising an impeller located ahead of the wingtip chord, a wind turbine device located behind the wingtip chord, and an intermediate gear device for coupling the impeller with the wind turbine and/or generating power. Thus, unlike the present invention, the impeller is powered by the vortex-driven turbine and no external power source or external driving engine is used. [0017] However, it is the inventor's understanding that the solutions disclosed in the prior art documents show a flaw in their very origin: they are conceptually wrong in that they only attempt to hinder the vortex generation, when in fact, the correct thing to do seems to be to take advantage of it. By means of the advantages of the present invention, the vortex might diminish, disappear, or even be inverted. Once, the tip vortex disappears, the induced forward movement drag will also disappear.

[0018] SUMMARY OF THE INVENTION

[0019] The present invention consists of a shaft or mechanical axis placed at the wingtip of each of the wings of the aircraft and a set of one, two, three or more propellers mounted on each of the shafts. Both shafts are driven by a single central engine and each shaft supporting the propellers receives the mechanical power to rotate these. The preferred embodiment shown in the figures refer to mechanical driving power means. However, the skilled in the art will understand that these can also be replaced by electrical o hydraulic means while keeping the spirit of the invention. As a result of the propeller action, a second vortex is created, opposite to the wing tip vortex caused by the air flow around the wing tip, by which the wing tip vortex gets a) attenuated; b) totally cancelled or c) inverted. The resulting effect on the aircraft is that the induced forward movement drag due to the vortex effect gets strongly reduced or even cancelled.

BRIEF DESCRIPTION OF THE DRAWINGS

[0020] FIG. 1 is a schematic drawing of a wing profile showing the wing tip vortex.

[0021] FIG. 2 is a schematic drawing of a wing having a wing tip shaped as a quarter of a cone. [0022] FIG. 3 is a schematic drawing of a wing having a winglet shaped wing tip.

[0023] FIG 4. is a plant view of a preferred embodiment of a right angled wing according to the present invention.

[0024] FIG. 5 is a front view of the wing and the vortex cancelling propeller of Fig. 4.

[0025] FIG. 6 shows an extended schematic front view of a blade according to present invention.

[0026] FIG. 7 shows different embodiments of blade cross sectional views of the blade of Figure 6.

[0027] FIG. 8 shows an upper plant view of the wing shaft assembly of the present invention.

[0028[ FIG. 9 shows a plant cross sectional view of the wingtip propeller blades.

[0029] FIG. 10 shows a perspective view of the wingtip propellers of a first embodiment of the present invention under rotation.

[0030] FIG 1 1. shows a perspective view of the wingtip propellers of a second embodiment of the present invention under rotation.

[0031] FIG. 12 shows an upper plant view of a third embodiment of the present invention.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0032] Detailed descriptions of the preferred embodiment are provided herein. It is to be understood, however, that the present invention may be embodied in various forms. Therefore, specific details disclosed herein are not to be interpreted as limiting, but rather as a basis for the claims and as a representative basis for teaching one skilled in the art to employ the present invention. [0033] In the following discussion, the terms "including" and "comprising" are used in an open-ended fashion, and thus should be interpreted to mean "including, but not limited to... ."

[0034] The disclosed invention provides a relatively simple and inexpensive device for taking advantage of the wingtip vortex. As stated before, by means of the invention, the end vortex might diminish, disappear, or even be inverted.

[0035] FIG. 4 shows a wing portion of a first preferred embodiment of a right angle-winged aircraft, that is, an aircraft in which the lengthwise symmetry axis of each wing is perpendicular to the aircraft's main body. In FIG. 4 it may be seen that each set of wingtip propellers is formed by three propellers 4, 5, 6, mounted on the common shaft 2. A schematic plant view of a wing 9 may be seen in which the power torque transmission shaft 1 is driven by the central engine 8 (only shown in FIG.8) to the wing tip. The drawing shows the shaft or mechanical transmission axis 2 on which the at least one propellers are mounted. The black circle 3 represents the coupling means among both shafts, and may be based on a pair of conical gears, a worm gear reducer o any other means that may allow certain angle in the power transmission (in this particular case of 90°). In a preferred embodiment, propeller 4 is installed at the leading edge of the wing chord; propeller 5 is installed in a slot 11. In a first preferred embodiment, single slot 11 is centered in the middle of the wing chord between the leading and trailing edge to allow the free blade rotation and propeller 6 is installed at the trailing edge of the wingtip. The arrow 13 shows the forward advance direction of the wing. A perspective view of this embodiment may be seen in FIG. 10.

[0036J Turning now to FIG. 5 and FIG. 6 and referring to the design of the propeller blades, these are non-ducted turbine type and may have a plurality of blades. In the shown example each propeller comprises four blades. It is well known that, in normal propeller blades, if dividing these in equal thirds, the portion that most contributes in generating impulse for lift is the most external one because its tangent linear speed is the highest. The inner portion of the blades, that is the nearest one to the hub, is made with an elliptical profile and, due to structural reasons, the tangent linear speed of this portion is too low for generating enough impulse for obtaining lift.

[0037] However, for the particular design of the present invention, if the at least one propeller is installed at the wingtip, it can be designed comprising iso-lift blades, that is, blades showing the same high lift profile all along its length, from the hub to the tip. FIG. 6 shows the blade divided in equal thirds: an inner third A closest to the hub, a middle third B and an outer third C near the blade's tip. The iso-lift blade's working principle is the following: the inner third A generates lift because right next to the wingtip the vortex is very intense; the middle third B generates lift in part due to its tangent linear speed and in part due to the wingtip vortex intensity at that point; and the external third C generates lift due to its high linear tangent speed. FIG. 7 shows the blade^'s cross sectional profile when moving from the hub to the tip: the inner third A shows a thick and curved profile in order to show good aerodynamic and structural strength; the middle third B shows an intermediate thickness profile and the external third C is made with as thin profile. All three thirds will thus generate a substantially better traction thanks to the wingtip vortex thrust. Also, thanks to its relatively thick and curved profile, the inner third A generates a strong inverse vortex effect. Therefore, when installing at least one propeller at the wing tip, the end vortex will be weakened, cancelled or inverted, by which the induced forward movement drag will also be weakened or cancelled. [0038] As was already mentioned, the use of two or more independent engines placed at the wing tips, to drive vortex cancelling elements as may be seen in most prior art solutions, is very risky. In the event that one of the engines stops working, a great yaw torque will result, urging the aircraft to rotate in a horizontal plane around a central vertical axis. The present invention avoids this issue basing the power to drive both propellers in a central power means and transmitting it via a gear box to the corresponding power transmission shafts and from there to each of the wingtip propeller shafts. In this way only two scenarios are possible: either both wing tip propeller assemblies work or none of these do so. The skilled in the art will understand that, the central power means may comprise a single engine or, in order to enhance reliability, a second central backup engine may be used. Typically, the central power means may be an internal explosion motor supplied by diesel oil, fuel oil, regular gasoline, aircraft gasoline or else a turbine based engine.

[0039] Due to the fact that wings are cantilever-mounted structures, they act as a flexing springs that are forced to vibrate by the vertical wind blasts. Therefore, in order to make sure that the power transmission shafts follow the wings' flexing deformations, the former are mounted on several bearings fixed to the wings. In this way, also part of the wings' shear load and flexure stress loads may be absorbed by these shafts.

[0040] FIG. 8 shows a plant view of the aircraft fuselage and its wings, in which the small rectangles 7 represent the bearings supporting both power torque transmission shafts 1. In this figure, engine 8, mounted inside the aircraft's main body 14 transfers its mechanical power to a central gear box 10 which in turn transfers the mechanical power to both power torque transmission shafts 1 and these drive the gear boxes 3 (in this figure the vortex propellers are not shown). Both power torque transmission shafts 1 may be directly coupled to the single engine and transmit the power torque to corresponding gear boxes 3. Also, in another embodiment the central power plant may comprise an electrically driven power engine or alternatively a central hydraulically driven engine. FIG. 8 also shows the symmetry geometrical axis 15 of the wings 9 which, in this embodiment form part of a right angled-wing aircraft.

[0041] The propeller placement shown in FIG. 4 is important for obtaining the features of the present invention. Each blade may be considered under a structural point of view as being a cantilever-mounted beam and, while the blades are rotating for generating thrust, the variable aerodynamic loads force them to vibrate. It is important to take in account that, in the lower vibration modes (1^st, 2^nd, and maybe in 3^rd) the amplitude of this vibration is maximum. Therefore, the criterion for placing the leading and trailing edge propellers 4, 6 (particularly those of the trailing edge) is to position the propeller plane as close as possible to the edge (in order to assure that maximum vortex interaction) but far enough to allow the propellers to vibrate without interfering with the mentioned edge. Also, the careful slot 11 dimensioning is important for obtaining an enhanced vortex cancellation. In a preferred embodiment, typical vortex cancelling propellers 4, 5, 6 may have a blade length of 60 - 80 cm and a cross sectional chord of 20 - 25 cm. In order to take in account the blade flexure deformation, the slot 11 is dimensioned to keep a 5 - 10 cm free space on each side of the propellers' blades and, therefore, the slot may have a total width of approximately 45 - 50 cm and a length of 80 - 100 cm.

[0042] Turning now to the propeller 5 mounted on an intermediate spot of the wingtip shaft, this propeller is built as a fixed step propeller because there seems to exist no practical way of letting the wing slot 11 change its width to accommodate a variable pitch propeller. The slot 11 must be designed and dimensioned with the same criterion as the one mentioned before: its size and shape should be adequate to allow the free propeller rotation with the blades under their maximum vibration deflection amplitude. The placement of slot 11 is also important. Preferably, in a first embodiment it will be aligned and placed on the geometrical center axis of the wing's length. However, in alternative embodiments, slot 11 might be shifted to be nearer to either the leading edge or the trailing edge of the wing to take account of different particular designs of the wingtip vortex cancellation.

[0043] FIG. 11 shows an alternative embodiment in which the vortex cancelling effect may be enhanced by using four propellers and, for this, also two parallel slots 11 in each wing may be used, each slot 11 having a wingtip vortex cancelling propeller 5 rotating inside and sharing the fixed shaft 2. The skilled in the art may understand that more propellers and slots might be used if required for enhancing the vortex cancelling effect.

[0044] When the wingtip propellers rotate, as they pass nearby the external wingtip semicircle, they create a downflowing air current which generates an upward thrust. As the wingtip propellers pass nearby the wingtip interior semicircle, they generate an upflowing air current which, combined with the forward moving speed, creates a high pressure zone behind the slot, and this enhances the lift. As a result, the lift diagram as a function of the wing length will no longer be a half ellipse but will instead turn into a trapezium, in which the diagram extends right to the wingtip itself. Therefore, as a side effect result, when mounting the propellers as shown in figure 4, the overall wing lift of the aircraft will also be enhanced. The propeller blade portions next to the hub overlap in order to allow the inverse vortex generation at the vortex core to be continuous. A flattened projection of a portion of the propeller circle with its corresponding blades is shown in FIG. 9.

[0045] The above description was aimed to aircrafts using right angled wings, that is, the symmetry axis 15 is perpendicular to the aircraft's main body 14. However, a great number of airplanes use the so-called swept wings, that is, wings usually forming an angle that slants these usually backwards (although well known, aircrafts using front-angled swept wings are less used). FIG. 12 shows a typically swept wing 12 in which each of the wing's symmetry axis 15 is slanted backwards respect of the aircraft's main body 14, forming an angle a which is less than 90°. Although not shown in FIG 12, a first embodiment includes a single slot 11 placed at the middle of the wingtip of the swept wing, halfway between the leading edge and the trailing edge. In FIG. 12 it may be seen that, by a way of example, three slots 11 accommodate a leading propeller 4 and two central propellers 5, being the trailing propeller 6 external to the wing. As may be noted in the drawing, the position of slots 11 in the swept wing is quite unexpected and non- obvious for the skilled in the art. The fact that the propellers 4, 5, 6 must be placed on the shaft 2 so that they face the wingtip vortex in a maximum cancellation direction, teaches the slot 11 positioning as shown in the drawing. In other words, the slots 11 do not follow the wings' angle but are placed instead perpendicular to the airplane's main body 14. Therefore, if the wing is swept, leading propeller 4 cannot be mounted outside the leading edge as shown in FIG. 4 following the wings' s angle, because it would be slanted respect of the vortex central rotation vector and, therefore, propeller 4 must be placed behind this edge and, as stated before, keeping a right angled position respect of the aircraft's main body 14. [0046] The ultimate goal of the wingtip vortex cancelling propeller system of the present invention is, whenever possible, to even invert the wingtip vortex, in order to cancel the induced drag and, also, to enhance wing lift.

Claims

CLAIMS What we claim is:

1. An aircraft propulsion system acting on the vortex formed at the tip of the wings (9, 12) of an aircraft, said aircraft having a main body (14), said aircraft propulsion system comprising: a fixed shaft (2) mounted on each tip of said wings, wingtip propeller means (4, 5, 6) mounted on each of said fixed shafts (2), power generating means (8, 10) for driving said propellers means (4, 5, 6), power torque transmission shafts (1) for transmitting the power of said power generating means (8, 10) to said propeller means (4, 5, 6), characterized in that:

said power generating means (8,10) comprises a central engine means (8) and a central power gear means (10);

said wingtip propeller means (4, 5, 6) comprising at least one propeller (4, 5, 6) placed at the tip of each wing;

wherein each wing comprises at least one slot (11) and wherein each slot (1 1) accommodates one of said at least one wingtip propellers (4, 5, 6) to rotate inside therein.

2. The aircraft propulsion system of claim 1 , wherein said aircraft is straight winged, being the central symmetry axis (15) of each of said wings (9) perpendicular to said main body (14), and wherein each of said fixed shafts (2) is mounted on each wingtip parallel to the aircraft's main body (14).

3. The aircraft propulsion system of claim 1 , wherein said aircraft is swept winged, being the central symmetry axis (15) of each of said wings (12) slanted respect of said main body (14) and wherein each of said fixed shafts (2) is mounted on each wingtip parallel to the aircraft's main body (14).

4. The aircraft propulsion system of claim 2, wherein each wing (9) comprises a single wingtip propeller (5) and a single slot (11).

5. The aircraft propulsion system of claim 3, wherein each wing (12) comprises a single wingtip propeller (5) and a single slot (11).

6. The aircraft propulsion system of claim 4, wherein said single slot (11) is centrally aligned with said central symmetry geometrical wing's axis (15).

7. The aircraft propulsion system of claim 4, wherein said single slot (11) is displaced respect of said central geometrical symmetry wing axis (15), nearer to the wing's leading edge.

8. The aircraft propulsion system of claim 4, wherein said single slot (1 1) is displaced respect of the central geometrical symmetry wing axis (15), nearer to the wing's trailing edge.

9. The aircraft propulsion system of any of claims 4 or 5, wherein each of said wings (9) comprises more than one slot (11).

10. The aircraft propulsion system of any of the preceding claims, wherein the direction of the length of the at least one slot (11) is perpendicular to said aircraft's main body (14).

11. The aircraft propulsion system of any of the preceding claims, wherein each of said at least one slot (1 1) has a length of 90 - 100 cm, a width of 45 - 50 cm and accommodates a vortex cancelling propeller having a blade that is 30 - 40 cm long and having a chord width of 20 - 25 cm.

12. The aircraft propulsion system of claim 1 , wherein said central engine means (8) comprises at least one central engine and said central gear means (10) comprises a central gear box (10).

13. The aircraft propulsion system of claim 12, wherein said central engine (8) comprises an internal explosion motor supplied by a fuel selected from the group formed by: diesel oil, fuel oil, regular gasoline, aircraft gasoline.

14. The aircraft propulsion system of claim 12, wherein said central engine (8) comprises a turbine based engine.

15. The aircraft propulsion system of claim 12, wherein said central engine (8) comprises an electrically driven power engine or a central hydraulically driven engine.

16. The aircraft propulsion system of claim 1, wherein said power torque transmission shafts (1) transmit its mechanical power to said fixed shaft (2) by means of a conical gear box (3).

17. The aircraft propulsion system of claim 1 , wherein said wingtip propellers (4, 5, 6) rotate in a direction against the corresponding wingtip vortexes' rotation direction.

18. The aircraft propulsion system of claim 1, wherein each of said power torque transmission shafts (1) is supported by a plurality of bearings (7) fixed on the corresponding aircraft's wings (9, 12).

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