GB2434785A

GB2434785A - Twin fuselage aircraft

Info

Publication number: GB2434785A
Application number: GB0706518A
Authority: GB
Inventors: John Edward Randell
Original assignee: Individual
Current assignee: Individual
Priority date: 2003-06-24
Filing date: 2007-04-04
Publication date: 2007-08-08
Also published as: GB0706518D0

Abstract

The present invention relates to a fixed-wing twin-fuselage aircraft comprising two fuselages 37a and a pair of wings 1a. One wing is attached to one fuselage and the other wing is attached to the other fuselage, the two fuselages being connected by a centre wing 1c. The centre wing 1c may extend for approximately only half the length of the aircraft, or it may extend for approximately the whole length of the aircraft minus the nose length. The aircraft may comprise sustaining means for sustaining a stream wise vortex adjacent a root of each wing.

Description

A FIXED-WING TWIN-BODY AIRCRAFT

The present application relates to a fixed-wing twin-body aircraft.

The reader is referred to UK Patent Application No. 0130860.0 in the name of the applicant.

The present invention relates to "twin-body" aircraft, i.e. aircraft in which two bodies (also known as fuselages) are connected by a centre wing. Twin-body aircraft have some clear advantages over single body aircraft.

A significant lift force can be produced on the two bodies and the centre wing of a twin-body aircraft by their combination, since the bodies and wing assembly so formed behaves as a large wing of small aspect ratio (of span to breadth). Wings of small aspect ratio are known not to stall readily and the displacement of air downwards by a centre wing, between the bodies, towards the tail of the aircraft (along the centre) reduces the tendency for flow separation to occur from the main wings at high angles of pitch. This reduces the risk of stall at high pitch and this permits a reduction in the length of runway required for takeoff by the aircraft. However, whilst twin-body aircraft have been constructed in the past (the reader is referred to the Blériot 125 and the North American F-82G Twin Mustang), they have not been developed as modern passenger aircraft. p

A disincentive to development of twin-body aircraft is aerodynamic complexity, which is due to the additional variables involved in the aircraft design. This complexity can render testing prohibitively extensive.

It is not possible to sustain a single rotational vortex about the nose of a single body aircraft, unless the resultant uni-directional torque is countered, for example, by ailerons.

Therefore, a single body aircraft generates positive pressure which acts against the nose of the body. Such positive pressure acts as a drag on the single body aircraft.

It is an aim of the invention to provide a practicable fixed-wing twin-body aircraft.

According to a first aspect of the invention there is provided a fixed-wing twin-body aircraft comprising two bodies and a pair of wings, one wing being attached to one body, the other wing being attached to the other body, wherein the aircraft comprises sustaining means for sustaining a stream wise vortex adjacent a root of each wing.

"Body" can be used interchangeably with the term "fuselage". By "stream wise" vortex it is meant that the axis of the vortex lies in the direction of motion of the aircraft.

"Sustaining" is intended to mean keep substantially uniform.

A vortex is preferably generated about the nose of each body. This generates suction on each nose, instead of the positive pressure generated on the nose of a single body aircraft.

In this way, resistance to movement through the air is reduced, which reduces fuel

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consumption. The division of the oncoming air into vortex flows about the twin noses permits air to flow smoothly past the aircraft with little turbulence and loss of energy provided the aircraft is designed appropriately in other respects.

The two bodies are preferably separated. Preferably, the bodies are separated by, and connected by, a centre wing. More preferably, the minimum distance between the bodies is substantially equal to the diameter of a body.

The wing is preferably dimensionally designed in the region of the root to give rise to the sustaining of a stream wise vortex at or adjacent each wing root. Preferably, the sustaining means compnses two subsidiary wings, a first subsidiary wing being connected to the main wing part of a first wing, the second subsidiary wing being connected to the main wing part of the second wing, and each subsidiary wing being inclined relative to its respective main wing part.

In a preferred embodiment of the invention, the subsidiary wings are inclined downwards from the inbound ends of the associated main wing parts.

The subsidiary wing is beneficially of a length equal to approximately half the diameter of a body. In this way, a vortex core substantially equal in diameter to the diameter of the body can form.

Bodies used in twin-body aircraft are generally of small diameter relative to the bodies used in single body aircraft. Therefore, obstruction of the perimeter of the bodies of twin-body aircraft by wings is relatively greater. This greater obstruction can cause difficulty in evacuation and loading.

It is another aim of the invention to overcome this major disadvantage of twin-body aircraft having conventional wing arrangements by permitting the main wings to be raised to a level above the bodies in an arrangement that maintains a high aerodynamic efficiency.

In preferred embodiments of the invention the subsidiary wings are oriented at between degrees to 75 degrees to the horizontal plane through the longitudinal axis of the bodies. Preferably, the subsidiary wings extend outwards radially from the body (outer surface) at angles of +30 degrees and -30 degrees respectively to the vertical plane through the longitudinal axis of the respective body. Stable vortex cores are then formed centred on the respective junctions of main wing parts and subsidiary wings.

In the preferred form of the invention, the sustaining means fixes stream wise vortexes at or adjacent the leading edge of the main wing parts.

Preferably, the main wing part and the sustaining means, which is in the form of a subsidiary wing, form a junction on which a vortex can centre.

In one preferred embodiment of the invention, the subsidiary wings are of aerofoil section (i.e. shaped to produce more lift than drag). In an alternative preferred embodiment of the invention, the subsidiary wings are of aerodynamic (but not aerofoil) section. For example, the wing section may have a line of symmetry about a chord between the leading and the trailing edge.

Flaps may be provided on the subsidiary wings for flight control.

Preferably, each wing comprises a first main wing part and a second subsidiary wing part.

The first wing part and the second wing part are preferably of generally aerofoil cross section. The first wing part, comprising the wing tip, is substantially longer than the second wing part. The first wing part extends in a first plane and the second wing part preferably extends at least partly in a second plane at an angle to the first plane, the angle being between 105 and 135 degrees. In one preferred embodiment, the angle is approximately 120 degrees.

According to a second aspect of the invention there is provided a fixed-wing twin-body aircraft, in which the two bodies are spaced by a centre wing, and each body has a respective outer wing extending from it, wherein the outer wings are inclined upwardly where they extend from the bodies.

According to a third aspect of the invention there is provided a fixed-wing twin-body aircraft comprising a pair of bodies and associated wings, each of the wings comprises a main wing part and sustaining means, wherein the sustaining means is designed to fix stream wise vortexes at or adjacent the leading edge of the main wing parts.

According to a fourth aspect of the invention there is provided a fixed-wing twin-body aircraft comprising a pair of bodies and associated wings, each wing extends outwards from a body and comprises a subsidiary wing, wherein each subsidiary wing extends from the body at an angle of approximately thirty degrees to a vertical plane taken through a central longitudinal axis of the body.

Preferably, the subsidiary wings extend radially from the body.

According to a fifth aspect of the invention there is provided a fixed-wing twin-body aircraft, the aircraft comprising a pair of bodies and a pair of wings, each body having a longitudinal axis, each wing having a first wing part and a second wing part, both wing parts being of aerofoil cross section, wherein the first wing part of a first wing extends outwards from a first body at an angle of 15 to 45 degrees to a vertical plane through the longitudinal axis of the first body, the first wing part of the second wing extends outwards from a second body at an angle of minus 15 to 45 degrees to a vertical plane through the longitudinal axis of the second body.

Preferably, the second wing parts extend substantially horizontally from the outbound ends of the first wing parts.

The second wing parts are preferably considerably longer than the first wing parts.

Preferably, lines are extrapolated from the planes of the first wing parts of the two wings, and an angle between the lines where they bisect is approximately 60 degrees.

Preferably, each wing has a pair of engines, and the engines are suspended from the wings. In one embodiment of the invention, an aerofoil extends between the engines in each pair, the uplift that the aerofoil creates thereby assisting in supporting the engines.

Each engine pair may be suspended from its associated wing by a single central strut.

Alternatively, each engine pair may be suspended by two struts. A single strut may join the aerofoil. Where Iwo struts are employed, a first strut may join a first engine of the pair, and a second strut may join the second engine.

A fixed-wing twin-body aircraft in accordance with one of the preceding aspects of the invention may comprise a pair of front wings and a rear tail plane mounted on supports which extend upwardly from each body. Alternatively, the aircraft may comprise a pair of rear wings and a pair of front stabilisers. The rear wings and stabilisers may have the same profile (e.g. elbow like) when viewed from the front. (See Figure 29a).

A centre wing may extend for about half the longitudinal length of the aircraft. In another arrangement, the centre wing may extend for the length of the aircraft minus the nose length.

An air flow theory, and an aircraft designed using the theory, will now be described with reference to the accompanying drawings, in which:-Figure Ia is a schematic front view of part of an aircraft in accordance with the invention, Figure lb is a further front view of the part of an aircraft of Figure la, Figure 2 is a plan view of a fluid flow path, Figure 3 is another plan view of a fluid flow path, Figures 4 to 10 are plan views of various flow patterns, Figure 1 la is an isometric view of a flow pattern, Figure 1 lb is an isometric view of a flow pattern, Figure 12a is si isometric view of a flow pattern, Figure 12b is a plan view of a flow pattern, Figure 13 is an isometric view of a flow pattern, Figure 14 is a side view ofa flow pattern, Figure 15 is a frontal view of a flow pattern, Figure 16 is an isometric view of a flow pattern, Figure 17a is a frontal view of a flow pattern, Figure 17b is a plan view of a flow pattern, Figure 18s is a plan view of a flow pattern, Figure 18b is a side view of a flow pattern, Figure 18c is a vertical section, in a plane perpendicular to the direction of flow, of a flow pattern, Figure l9a is a schematic plan view of part of an aircraft, Figure 20 is a view from below of part of an aircraft, Figure 21 a is a side view of a centre wing of an aircraft, in a vertical section through the centre line of the aircraft, Figure 21b is a front view of the wing of Figure 21a, Figure 21c is a vertical section through a trailing edge of the centre wing of Figure 21a, Figure 22a is a side view, in section, of part of a wing, Figure 22b is a side view, in section, of part of a wing, Figure 23 is a further side view, in section, of a wing, Figure 24a, 24b and 24c are side views, in section, of a wing and tail plane, Figure 25a is a front view of part of an aircrafl, Figure 25b is a view from below of part of an aircraft, Figure 26 is an isometric view of an aircraft, Figure 27a is a side view of a flow pattern,

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Figure 27b is a side view, in section, of a wing using the characteristics of the flow pattern of Figure 27a, Figure 28a is an enlarged iew of the wing of Figure 27b, Figure 28b is a variation on the wing of Figure 28a, Figure 29a is a frontal view of part of an aircraft, Figure 29b is a plan view of the aircraft part of Figure 29a, Figure 29c is a side view along the centre line of the aircraft part of Figure 29b, Figure 30 is a view from below of a wing arrangement and a flow pattern resulting from the arrangement, and Figure 31 is an isometric view of an aircraft.

A proposed wing arrangement for a twin-body aircrafI is shown schematically in Figure Ia, which is a front view of two bodies 37a and 37b of circular cross section with a centre wing Ic in the horizontal plane through the centre lines of the bodies. The centre wing Ic has a span equal to the diameter of a body 37a, 37b. This arrangement has been found by private model tests to provide aircraft that are stable over a wide range of pitch values.

In the above arrangement, when a centre wing Ic displaces air downwards between the bodies 37a, 37b, as shown in Figure Ib, in which the bodies and centre wing are shown by broken lines, the upper portions of the bodies behave as a mouthpiece and there is little loss of energy in the downwards displacement of air. Air flows downwards and inwards as shown by the clockwise and anticlockwise arrows 97a and 97b, towards the centre wing Ic in two vortex flows about the bodies 37a, 37b. Air is then discharged substantially vertically downwards, as shown by the flow arrows 99.

Main wings, I a and lb. are raised to a level above the bodies as shown in Figure Ic. The bodies 37a, 37b are then more accessible on the ground. Aerodynamic efficiency is maintained in the arrangement shown, in which the main wings, Ia and lb. are connected by subsidiary wings, 39a and 39b, to the bodies, 37a and 37b. Stream wise vortexes, shown by arrows 35va and 35vb, are generated about the roots Y and Z of the main wings Ia and lb where they join the subsidiary wings 39a and 39b respectively.

The centre wing Ic, shown schematically in Figures Ia, lb and Ic by a single line, is desirably of considerable thickness relative to the diameter of the bodies, to provide structural strength, and therefore has a large frontal area. The disposition of the centre wing Ic in relation to the main wings la, lb and the nose and tail points of the bodies 37a, 37b has a large influence on the aerodynamic efficiency of the aircraft. The vortex flows downwards between the upper portions of the bodies (and the stream wise vortexes generated by the subsidiary wings) exist within vortexes created by the main wings. Since wings create forced vortexes, this gives rise to vorticity and a design theory for flow of this nature is required to design an efficient aircraft of this kind.

The wings of an aircraft produce lift by displacing air downwards but also displace air to the sides. Air thus moves sideways as well as upwards and downwards as it parts ahead of an aircraft about an aerodynamic centre line. If air descends vertically behind an aircraft and is not dragged forwards with it (including air from the sides), lift is produced with minimum drag and fuel consumption is minimised.

To describe the flow of air past an aircraft as it is displaced to the sides, a volume of air is assumed here to be able to rotate within the body of the flow. Such a volume of air is represented in Figure 2 by a cylinder, 2, shown in a plan view, which has its axis vertical and perpendicular to the direction of flow, the latter being shown by the arrow 4. The cylinder 2 of fluid is forced away from the centre line and is therefore shown to be subjected to a lateral acceleration, represented by an arrow 6. This acceleration causes the cylinder 2 of fluid to move in a circular arc 8 about a centre f. If the motion about the centre f develops as a vortex, a velocity gradient exists along a radius of the vortex and therefore a difference in velocity across the cylinder 2 along a radius of the vortex, a velocity difference which causes rotation of the cylinder 2 as shown by the arrowhead 18.

Two cylinders 2, 10 of fluid flowing side by side in a uniform flow of fluid, as shown in Figure 2 by the cylinder 2 and a second cylinder 10, can separate (at the point e in Figure 2) and follow divergent paths, the second cylinder, 10, following the path Sa in Figure 2.

In Figure 3, a cylindrical envelope 12 is shown which contains the two smaller cylinders, 2 and 10. This cylindrical envelope 12 can behave in the same manner as the cylinder 2 (alone) when subjected to a lateral acceleration and so likewise move along a circular arc 8 about the centre f. Flow along the arc 8 will be accompanied by rotation of the envelope because a velocity difference, represented by the arrows 14 exists on its diameter.

Rotation of the cylindrical envelope 12 can produce, by contact, rotation of each of the smaller cylinders, 2 and 10, contained within it, as shown by the arrowheads 16 and 18.

At the point of contact of the two cylinders, 2 and 10, the flows due to rotation are mutually opposed. This causes shear stress at the point of contact and a tendency for separation of the two cylinders, 2 and 10, to relieve the stress. This can occur at a point if a state of balance is altered, as at the point B shown in Figure 4. At point B, a tangent to the circular arc 8 passes through the centre j on the other side of the centre line c-c, so the flow at point B about the centre f is also radially away from the centre j. A state of balance exists at point B because divergent flow paths can be taken from the radial line through the centre j. The cylinder 2 can flow along a circular arc 20 about the centre f whilst the other cylinder 10 can follow a circular arc 22 about a centre c. The centres j, I and c are equidistant, lying at the apexes of an equilateral triangle with the point B at the mid-point of the side f-c.

If the two cylinders of fluid, 2 and 10, separate initially at the mid-point e between the centres f and j as shown in Figure 2 and again in Figure 5, the rotating cylinder 2 can induce a vortex flow about the point B, as shown in Figure 5, and the other cylinder 10 can induce a vortex about the point A on the other side of the centre line c-c.

The vortexes about A and B are generated by boundary contact with the cylinders 10 and 2 respectively and are of twice the diameter of cylinders 2 and 10. If adjacent cylinders of fluid (2', 2") on the line e-f also flow in circular arcs about the centre f, rotational motion of the cylinders when aligned along the line B-f will produce, by aggregation, a vortex Sb about the point B of a radius equal to the distance from the point B to the centre line c-c.

In the limit, the cylinders of fluid 2 and 10 considered in Figure 2 are infinitely small and at points initially on the centre line (at point e) between centres I and j. The cylinders 2 and 10 then move along circular arcs 42b and 44b, respectively, as shown in Figure 6 to produce vortexes 3b and 5b about the mid-points A and B respectively of the sides jc and f-c of the equilateral triangle f-c-j. Since the volume of fluid in rotation can increase along a circular path from the initial point e to either A or B, the diameter of a rotating volume of fluid can increase from zero at point e to the distance of A from B about points A and B, as shown in Figure 7 by vortexes 3b and Sb about A and B respectively. By the same reasoning, flow in the vortexes 3b and Sb about A and B can generate vortexes gv and hv about the points g and h, respectively, the midpoints of the sides A-c and B-e of an equilateral triangle A-c-B. In turn, just as a vortex about the centre generates a vortex about point A, so a vortex about the centre g can generate a vortex shown by the arrow q2v about the point q2, the mid-point of the side g-O of the equilateral triangle g-O-h.

Referring to Figure 8, vortexes about the points A and B can, by symmetry (about the line A-B), generate vortexes about the points G' and H' (aswell as g and h), as shown in Figure 8. These vortexes, G'v and U'v about points G' and H' respectively, are compatible with flow in vortexes about the centres J" and F" (which have a similar relationship to the centres G' and H' as do centresj and f to centres g and h). As shown in Figure 8, a particle of fluid may therefore flow along the circular arc 32, in a vortex about j, transfer at the point g to flow along the circular arc 36 about the vortex centre at B to point G' and then transfer to flow in a circular arc 46 about the centre J". The arcuate path (through the circular arcs 32, 36 and 46) can be followed by a body of fluid. Flowing along the arc 32, the body of fluid is caused to rotate about the centre g. Between the points g and 0', as the body of fluid flows along the arc 36, the direction of rotation is reversed. Reversal of the direction of rotation occurs again at the point G' so that when the flow crosses the line J"-c, a cycle is completed. There is then no residual rotation as a result of the flow through the arc wave described from section j-f to section J"-F". Since such a flow path can also exist on the opposite side of the centre line e-c, through an arc wave comprising the circular arcs 34 about centre 1, arc 38 about centre A and arc 48 about centre F', flow paths on each side of the centre line c-c can diverge from the linej-f and converge again before passing through the line J"-F". Flow paths that divide at the centre a and recombine at the centre c can also follow paths of circular arcs. Thus, as shown in Figure 9, a particle of fluid at point e can follow a circular path 50 about a centre a to point g, then follow a circular arc 54 about the mid-point 0, passing through point A to point 0', at which the flow can transfer to the circular arc 58 about centre A" to bring the particle of fluid back to the centre line at c. From the same position at e, another particle of fluid can flow along the circular arc 52 about centre b, arc 56 about the mid-point 0 and then arc 60 about centre B", passing through points h, B and H' in following this path.

The flow paths described with reference to Figs. 8 and 9 coexist, but both involve rotation at points g and G' on one side of the centre line c-c and points h and H' on the other side. Bending stresses arise in the actuate wave flow described (which bending stresses can be considered to cause rotation) because the flows described are within a body of fluid and such flows will be sustained only if viscosity is relatively high. This is not true in considering the flow of air past an aircraft. A series of waves will therefore be formed as waves of smaller arc curvature are formed to reduce flow bending stresses.

An ordered series of such waves is illustrated in Figure 10, in which the flow is shown to diverge ahead of a pressure field centered on the point c. In consequence, a series of pressure arcs is established, shown by the arc 24 between centers j and f, arc 26 between centers A and B, and the arc 28 between centers G' and H'. Subsidiary pressure centers exist on the centre line, at the mid-point 0 between points A and B, and at the midpoint e between points j and 1, and these set up subsidiary pressure arcs, such as the arc 30 between points g and Ii. Vortex flows arise at the end points of the pressure arcs, these being the positions of vortex centers previously described. One such vortex, gv about the point g, is shown for purposes of illustration in Figure 0.

In the flow past an aircraft, stresses can be relieved by rotation in a plane perpendicular to the direction of flight. Figure ha shows isometrically the vortex gv in the horizontal plane through points j, c and f. A vortex gr of the same diameter as the vortex gv and about the same centre g is also shown, but the vortex gr is in a vertical plane perpendicular to the centre line e-c. Flow in the vortex gr develops and diminishes as the body of the flow in the general direction of the centre line e-c passes through the perpendicular vertical plane through e to the perpendicular vertical plane through 0. The growth and decay of the vortex gr is represented, by broken lines, in the plan view in Figure 10 by the projection on the horizontal plane of a "bicone", Rg. The word "bicone" is used here, for brevity, to describe two identical cones arranged back-to-back with their bases coinciding, each cone having an included angle of 60 degrees at the apex in a plane through the centre line, being the volume of revolution of two identical back-to-back equilateral triangles with a common side about an axis through the remote apexes. The bicone Rg shown in plan in Figure 10 and isometrically in Figure Jib, is a rotation bicone, since it replaces (in the three-dimensional flow described) the vortex gv generated by rotation in the two-dimensional flow pattern of Figure 10.

Since rotation has been shown to arise about the centre G' also in the arcuate flow pattern previously described, a rotation bicone RG' will develop in the three-dimensional flow pattern also, as shown isometrically in Figure 1 2a. The rotation bicones Rg and RG' have a common apex and their axes are in line. The flow arc 36 (shown previously in Figure 8) passes between the centres g and Gin a volume defined by a flow bicone having apexes at the points g and G'. This flow bicone FgG' which contains the flow arc 36 is shown in a plan view in Figure 1 2b, together with the rotation bicones Rg and RO', and also the flow centres previously shown in Figure 8 for the flow on one side (only) of the centre line e-c.

Flow bicones are formed sequentially between the centres of rotation bicones, so fluid can progress through flow bicones in orderly flow. This is illustrated in Figure 10, in which two sequential bicones, FhH' on an axis hH' and FH'h' on an axis H'h' are shown in plan. The sequence can extend forwards and backwards, as illustrated in Figure 10, in which a series of bicones is shown along an axis yH -yH.

In the flow described with reference to Figure 2, cylinders of air of indefinite length were considered and these cylinders may be replaced by bicones (as previously defined) with the axes vertical and having a maximum diameter equal to the diameter of the cylinders considered. Two such bicones, 2c and lOc are shown isometrically in Figure 13. As elements of the cylinders previously considered, these bicones behave similarly so that when moving in a horizontal plane and separated at the centre point e on the line j-f can follow circular paths, 8 and 8a, respectively. In following these paths, like the cylinders of which they are elements, the bicones 2c and lOc are caused to spin about their axes, as shown by the arrows lSb and ISa, respectively, in Figure 13. The bicones may though move out of the horizontal plane.

As a bicone moves out of the horizontal plane, the axis of a bicone will be caused to rotate about its centre, as illustrated in Figure 14 in a side view of the bicone lOc. The bicone is shown to move upwards along a circular arc 8t about a centre t. Rotation of the bicone about a horizontal axis through its centre in doing so can cause the rotation of a volume of air in a bicone lOo that has a maximum diameter equal to the length of the axis of the contained bicone lOc. The rotation of the larger bicone lOo is shown in Fig 14 by the arrow 1 8t. While moving upwards, the centre of the bicone lOc also moves sideways away from the vertical centre plane through e-c and likewise the bicone 2c moves away to the other side as the bicones separate. The movement of the bicones 2c, lOc when viewed in the direction of flow of the body of air is then as shown in Figure 15 with the centres of the bicones following circular arcs. The centre of the bicone lOc follows a path along a circular arc 8s. Following this pathcauses the rotation, shown by the arrow 18s, of an internal bicone, lOi, within it. This internal bicone lOi has an axis of length equal to the maximum diameter of the bicone lOc that contains it. A bicone 2i within the biconc 2c can be created in a similar way as shown also in Figure 15 from the front and in an isometric view in Figure 16.

The bicone 2c is shown again in a view from the front in Figure 17a. Also shown in Figure 1 7a are adjacent bicones, 2c' and 2c", the centres of the three bicones lying on the line e-f(shown in Figure 15), each being an element of one of the aligned cylinders 2, 2' and 2" shown previously in Figure 5 (on the line e-f). If the three bicones 2c, 2c' and 2c" rotate together as one about the line e-f (as an axis), a combined bicone 2C' can be formed. By application of the same reasoning as was applied to the movement of bicones lOc and 2c, it is deduced that movement upwards (or downwards) and sideways of the bicone 2C' can lead to the creation of a bicone 2Ci' within it and a bicone 2Co' (shown in part only) about it. The bicone 2Ci' within the bicone 2C' is shown in a plan view in Fig 1 7b, together with adjacent bicones, 2Ci and 2Ci". The axes of these three bicones (2Ci, 2Ci' and 2Ci") lie in the same horizontal plane and together form a group like the group of bicones 2c, 2c' and 2c" shown in Figure 1 7a. Just as the latter group can form a larger bicone (2C'), so the group made up of bicones 2Ci, 2Ci' and 2Ci" can be contained within the larger bicone 2Co' shown again, in part, in Figure 1 7b.

A corresponding group of three bicones lOCi, lOCi' and lOCi", on the left hand side of the centre plane, is shown in plan in Figure 18a and in a view from the side in Figure 18b.

The horizontal axes lie in a plane Pm, the plane in which lie the centres f, j and c shown previously in Figure 13. The bicone lOc', with a vertical axis, is also shown in the view from the side in Figure 1 8b.

A vertical section in a plane perpendicular to the direction of flow (of the body of air) through the axis of the bicone lOCi' is shown in Figure 18c. On the right hand side of the centre plane are shown the bicones 2Ci', 2c' and 2i'. On the left hand side of the centre plane are shown the bicone lOc' and the adjacent bicone lOc, the axes of which lie in the same vertical section (perpendicular to the direction of flow) as the axis of bicone 2Ci'. It can be seen from Figure 18c that a bicone lOd can sit below the bicones lOc and lOc' (with the axes of the three bicones, lOc, lOc' and lOd in the same vertical plane) so that there is line contact between the bicone lOd and both the bicone lOc and the bicone lOc'.

Each of the three bicones tOe, lOc' and lOd can rotate about a horizontal axis through its centre in the direction of flow of the body of air. Just as an internal bicone, 21, can be created within the bicone 2c, as illustrated in Figure 16, by rotation, so an internal bicone 101 can be created by rotation in the bicone lOc. Similarly, an internal bicone 101' can be created within the b'cone lOc' and an internal bicone lOf within the bicone lOd. If rotation is sustained in the direction of flow of the body of air, a helical vortex will be produced having a diameter equal to the length of the axis of the rotated bicone. Thus, a vortex lOfv having a diameter equal to the length of the axis of the bicone lOf can be generated about the centre f. Similarly, a vortex lOpv can be generated about the centre p by rotation of the bicone l0i and a vortex lOqv about the centre q by rotation of the bicone 101'.

Referring to Figures 19a and 19b, an aircraft comprises a two substantially similar bodies 37a, 37b. The bodies 37a, 37b are of substantially circular cross section. A centre wing Ic connects the two bodies 37a, 37b together. The longitudinal axis of a first body 37a is parallel to the longitudinal axis of a second body 37b. The vortexes lOfv, lOpv and lOqv shown above are shown again in Figure 19a in a schematic view from the front of an aircraft wing arrangement. A main wing Ia in the plane Pm forms a junction with a subsidiary wing 39a along the axis of vortex lOqv, which vortex can be sustained by the subsidiary wing. The vortex lOfv can form about a body 37a centred on the vortex lOfv axis through point f. The vortex lOpv can be centred, at a point p. along an edge of an upper surface (in the plane Pm) of a centre wing Ic. On the right hand side in Figure 1 9a is shown a front view of the bicone 2Ci'. This bicone is shown in p'an view n Figure 19b, together with the corresponding bicone lOCi' on the left hand side.

The axes along which the vortexes IOpv and lOqv lie are also shown in the plan view of Figure 1 9b, by the chain lines yp-yp and yq-yq respectively. Another vortex axis, yg-yg, lies through the apex g of the bicone lOCi'. Figure 19b also shows the two circular bodies, 37a and 37b, with nose sections 37an and 37bn about which vortexes can form.

The bicone I OCI' has an axis that extends between an apex in the centre plane (i.e. The vertical plane through ec) and an apex at the point g, which is also shown in Figure 10.

Referring to Figure 10, the bicones IOCI' and 2Ci' together span the entire space between the centres g and h, the axes lying along a straight line connecting the points g and h. The bicone IOC? is shown again in Figure 20, in a view from below of a wing Ia and fuselage 37a on one side of a twin-body aircraft. The arrangement shown in Figure 20 is that shown schematically in the front view in Figure Ic. The span of the centre wing Ic is shown as 2/3 of a unit in the arrangement shown in Figure 20. The breadth of the centre wing Ic, i.e. the distance of the wing Ic, in the direction of flight, is then 3/3 units.

In Figure 20, point A is an apex of principal flow bicones ahead of the wing and behind it, one-half of each being shown by its horizontal projection on the plane of the wing.

Thus, the equilateral triangle eAj is the horizontal projection of one-half of a bicone formed ahead of the wing and the equilateral triangle cAJ" is the horizontal projection of one-half of a bicone formed behind the wing. If wing flaps (not shown to preserve conciseness and clarity) extend outwards to the line yA-yA, when the flaps are operated, say, for landing the vortex produced at the tip of the flap will lie in plan view along the axis yA-yA of the principal bicones. A series of smaller bicones is shown in Figure 20 along the axis yq-yq, with apexes at the points ql, q2, q3, and q4. Only one of these points is shown in Figure 7 (the point q2) but the other points (qI, q3, and q4) could be shown additionally in Figure 7 as potential vortex centres.

A side view of the centre wing ic is shown in Figure 21a in a vertical section through the centre line. The centre wing Ic, of aerofoil section, has a flat underside 62 and a pointed nose 64. The upper surface is formed of a circular arc 66 (about the centre 66c) from the nose 64 to a point at which the upper surface follows the line of a tangent (to the circular arc 66) to intersect the underside 62 at the trailing edge 68. A front view of the wing is shown in Figure 21b and a vertical section through the trailing edge in Figure 21c.

Referring to Figure 2 Ia, the upper surface of the wing 1 c rises to the horizontal plane, Pm, of the main wings (shown in Figure 2 Ia), at which section the top edges of the wing ic are spaced apart a distance equal to the distance between the bodies. Sides 70a and 70b (shown in Figure 21b) of the wing Ic, are inclined at 60 degrees to the horizontal as shown in Figure 21b. Below a horizontal plane througb the centre lines of the bodies, the centre wing Ic has a span equal to the distance between the bodies 37a, 37b, as shown in Figure 21b. Air is drawn downwards towards the trailing edge 68 as shown by the vertical flow 99 in Figure 21c, and this resists the tendency for vortex flows 97a and 97b above the upper surfaces of the bodies to be maintained at the rear. Suction on the tails of the bodies 37a, 37b caused by vortex flows about them increases drag on the aircraft.

Flow over the upper surface of the wing Ic can follow the circular arc 66, as shown by the half-arrow 66w in Figure 21a. The flow of air in such a circular arc causes bending stresses in the body of the flow (as previously mentioned in deducing Figure 10 from Figs. 8 and 9) and air therefore tends to flow along a circular arc 72 of smaller curvature above the wing Ic, about the centre 72c as shown in Figure 21a.

Air flows through arcuate waves of diminishing curvature as the height above the wing Ic increases. A set of geometrically related waves therefore exists above the wing 1 c, and also below it as the flow divides ahead of the wing and recombines behind it.

Referring to Figure 22a, the flow pattern changes as the pitch of the wing Ic is increased, and also the lift, until a critical value of the pitch is reached, at which a vortex 74v develops about the centre 74c. The centre of the vortex 74c lies in a horizontal plane through a nose point 64 of the wing Ic. A boundary 74 of the vortex is determined by the distance of its centre 74c from its trailing edge 68' at this pitch.

Development of the vortex 74v causes the flow pattern over the upper surface of the wing lc to be disrupted. Air tends to flow along an arc of reduced radius of curvature from the nose 64, along an arc 86 about a centre 86c. The wing then behaves as the aerofoil section shown in Figure 22b. A wing Ic' of this aerofoil section has the arc 86 about the centre 86c incorporated in its upper surface from the nose 64, followed by a straight section 82 that connects an end point 88 of the arc 86 to the trailing edge 68'. The flow past the wing Ic' resembles that shown in Figure 23, which is a vertical section on the centre line that shows the primary flows.

Referring to Figure 23, lift on a wing lc is produced largely by suction on the upper surface. A wave arc 84 about a centre 84c arises from the deflection of air upwards ahead of the wing Ic due to increase in pressure as the wing approaches the air. As the air parts ahead of the wing Ic, a flow path along an arc 76 about a centre 76c is followed below the wing Ic by an arc 78 about a centre 78c. At this pitch, the body of air within the circular boundary 74 is likely to rotate as a vortex. A wave constituted of arcs of smaller radius of curvature is formed, commencing with the arc 86 about the centre 86c. The initial arc 86 is followed by sequential arcs 90, 92, 94 and 96. Flow separation occurs at the upper surface of the wing I c at the point 88 where the flow arc 86 (along the top surface of the wing) joins the flow arc 90, which diverges from the surface of the wing.

Flow separation leads to loss of lift and aerodynamic stall of the wing Ic.

Rotation of the body of air within the boundary 74 as a vortex does not provide a stable flow pattern because at the trailing edge 68' of the wing the flows above and below the wing are in opposition. Referring to Figure 22b, the formation of vortexes, 100 and 102 within the vortex boundary 74 enables air drawn downwards and forwards towards the trailing edge 68' to be diverted both backwards with the flow past the underside of the wing Ic (joining the flow along the arc 106 about centre I 06c shown in Figure 22b) and also forwards into a suction region above the wing Ic, towards the point of flow separation 88. The flow forwards over the wing ic surface opposes the downwards flow over the upper surface of the wing (from the nose) and leads to the formation of a vortex 104 about a centre I 04c.

The flow pattern shown in Figure 23 is that prior to stall but at a critical pitch of the wing Ic when stall is imminent. Air can flow along an arc 98 (shown by broken line) about a centre 84d with no flow separation. Prior to rotation of air within the boundary 74 as a vortex, air can also flow along a wave formed by the arcs 86, 90, 92 and 94, shown in Figure 24a.

Referring to Figure 24, a tail plane 80' which is geometrically similar to a wing ic' in section is arranged so that its nose point 108 coincides with an end point of the flow arc 94. The tail plane 80' acts in concert with the wing Ic' and thereby assists in sustaining the unstalled pattern of flow over the wing Ic'. The wing Ic' can be replaced by a wing Ic of the aerofoil section shown previously in Figure 21 a and the tail plane similarly modified to be geometrically similar without changing their relative positions. When the wing Ic and a tail plane 80 are in a normal attitude they are arranged as shown in Figure 24b.

When the wing Ic and tail plane 80 are incorporated in an aircraft they appear as shown in Figure 24c.

Referring to Figures 25a and 25b, the aircraft is shown in part (the main wing and subsidiary wing on the right hand side being omitted) in, respectively, a front view and from below. The tail plane 80 is mounted on vertical tail fins, I 24a and 1 24b, mounted in turn on the tail sections of the bodies 37a and 37b. The main wing Ia and the body 37a are as shown in Figure 20.

The aircraft is shown isometrically in Figure 26.

Referring to Figures 25 and 26, the aircraft has two engines mounted as a pair beneath each wing. Referring to Figure 25, two engines, 118' and 118", are shown mounted as a pair below wing 1 a. The engines are connected by an engine wing le. The engine wing Ic at least partially supports the engines 118', 118" and reduces the structural load imposed on the wing, particularly during takeoff. A flap may be incorporated in the engine wing Ic to increase the lift in takeoff. Such a flap may also be used to balance the aircraft if asymmetrically loaded or used as an aileron for flight control.

Referring to Figure 25a, the engines are individually supported by vertical struts 119' and 119'. An engine pair may be supported instead by a single strut connected to the centre of the engine wing Ic.

The lift produced by the engine wing Ic may be supplemented by the use of additional wings, attached to the free sides of the engine casings.

The aerofoil section shown in Figure 21 is sensitive to change in pitch because of its pointed nose. The nose can be slightly rounded to reduce this sensitivity. Generally, ills advantageous to use a centre wing which has a nose formed in part from a sector of a cylinder. For this purpose, the vortex 104 shown in Figure 22b can be incorporated in the design of a nose of an aerofoil section.

The vortex 104 about the centre 104c shown in Figure 22b is shown again in Figure 27a.

In Figure 23, the arc 86 is shown as the first arc in a sequence, being followed by arcs 90, 92, 94 and 96. This series of arcs is also shown in Figure 27a, together with the arc 84 about the centre 84c (also shown in Figure 23). Referring to Figure 27a, the arc 84 extends over the wing between the end points 84t and 84t'. This overarching arc 84 is a characteristic of flow past the wing prior to the onset of flow separation, when air flows smoothly past the wing with little turbulence and drag. Maintaining the overarching flow arc 84 is therefore an aim in designing a round-nosed aerofoil section for use in a centre wing. This can be achieved by locating the trailing edge of the wing 68" in Figure 27a.

The chain line 108 that connects the vortex centre I 04c with the trailing edge 68" determines the pitch of the wing. With this pitch, flows above and below the wing divide ahead of the wing and merge behind it smoothly with little loss of energy. The flow is well ordered because vortex flows can exist about the centres 84t and 84t' above the level of the wing and about the centres 86c and 96c below the level of the wing.

A practical aerofoil section based on the flow pattern described with reference to Figure 27a is obtained by thickening above and below the pitch line 108, as illustrated in Figure 27b. The nose of the wing Ic (shown hatched) is formed in part from a cylinder of a diameter equal to that of the vortex 104, joined by a transition section below to a horizontal section to form the underside 62d to the trailing edge 68". This aerofoil section for a centre wing is shown to a larger scale in Figure 28a, which shows the arcs 86, 90, 92, 94 and 96, shown previously in Figure 27a, superimposed in relation to the underside of the wing section. It will be noted that arcs 92 and 96 (shown by broken lines) protrude into the aerofoil section. Whilst the flow cannot penetrate the section boundary, the pressure variation associated with a wave can travel along the underside of the wing, and thus sustain the flow regime in the body of the air flow past the wing ic.

Relative to the breadth of 313 units (in the direction of flight) of the centre wing Ic shown in the aircraft in Figure 25b, a wing of the section shown in Fig. 28a has a breadth of eight and two thirds units. This approximates to 5.13 units. Substantially the same wing section as shown in Figure 28a is shown in Figure 28b, but the wing section shown in Figure 28b has a breadth of 5.13 units in the direction of flight. A wing Ic of the section shown in Figure 28b can be mounted between two bodies, each of the same diameter (2/3 of a unit) and of the same length (6/3 units) as the body shown previously in Figure 20. Such an arrangement is shown in side view in Figure 29c, i.e. a body 37b and a centre wing Ic. The trailing edge 68" of the centre wing lc lies in the vertical plane through the tail points of the bodies 37a, 37b.

The upper surface of the centre wing Ic shown in Figure 29c can rise from its nose to the plane Pm of the main wing of the aircraft shown in Figure 28b. Such an aircraft is shown, in part, in Figure 29a and in Fig 29b (one of the main wings and its associated subsidiary wing being omitted in both figures).

In the aircraft, the span of the centre wing ic at the trailing edge 68" is 2 units (the overall width of the fuselage assembly comprising the two bodies and the centre wing).

The axis of a stream wise vortex generated at the tip of the trailing edge 68" of the centre wing ic lies vertically below the side of a body. This assists in maintaining a flow vertically downwards at the rear of the aircraft (providing stability and efficiency) when combined with the use of a subsidiary wing that generates a stream wise vortex along an axis vertically above the side of the body.

Referring to Figures 28a, 29a and 29b, a vertical tail fin 124 is mounted centrally on the rear section of the centre wing Ic.

The main wing Ia of the aircraft is mounted towards the rear of the body 37a and is connected to the body 37a by a subsidiary wing 39a. An engine 11 8a is mounted in the main wing Ia.

A horizontal stabiliser wing I I 6a towards the front of the aircraft is connected to the body 37a by a subsidiary wing 1 17a. When the bodies 37a, 37b are horizontal the main wing Ia and the stabiliser wing I 16a lie in the same horizontal plane above the bodies.

Furthermore, the subsidiary wing 39a is in the same plane as the subsidiary wing I 17a.

The flow pattern for the wings of the aircraft is shown in Figure 30, which is a view from below like that shown in Figure 20. The main wing Ia and forward stabiliser wing 1 16a are shown hatched. The outline of the engine liSa, mounted in the wing, is shown by broken line. The principal flow bicones shown in Figure 20 are shown in part by the projections of the equilateral triangles eAj and cAJ" on the horizontal plane. In this aircraft, with a rear mounted wing Ia, the axis yA-yA intersects the main wing Ia behind the point A, at the point A', so a rotation bicone centred on A' is formed on an axis A-A" of radius at its greatest section equal to the distance of point G' from its centre at A'.

The centre line of the engine 11 8a lies on the axis yg-yg, as does the tip of the forward stabiliser wing 1 16a. A bicone series is established ahead of the main wing Ia by the horizontal stabiliser wing 1 16a, which is of rectangular plan form, and has the leading edge of its tip at the point g and a trailing edge through point q2.

The body 37a shown in Figure 29 with a nose point fn and a tail point ft is at a lower level than the plane of the wings shown and is represented in Fig 30 by a chain line between the points fn and ft.

The aircraft described with reference to Figs. 29 and 30 is shown isometrically in Figure 31.

Claims

Claims I. A fixed-wing dual-fuselage aircraft comprising two fuselages

and a pair of wings, one wing being attached to one fuselage, the other wing being attached to the other fuselage, wherein the two fuselages are separated by a centre wing, and the centre wing extends for approximately only half the longitudinal length of the aircraft.

2. A fixed-wing dual-fuselage aircraft comprising two fuselages and a pair of wings, one wing being attached to one fuselage, the other wing being attached to the other fuselage, wherein the two fuselages are separated by a centre wing, and the centre wing extends for approximately the length of the aircraft minus the nose length.

3. A fixed-wing dual-fuselage aircraft according to Claim 2, wherein the centre wing extends from the inbound end of the aircraft nose to the rear of the aircraft.

4. A fixed-wing dual-fuselage aircraft according to any of Claims I to 3, wherein the aircraft comprises sustaining means for sustaining a stream wise vortex adjacent a root of each wing.

5. A fixed-wing dual-fuselage aircraft comprising two fuselages, each fuselage has a respective outer wing extending from it, wherein the fuselages are spaced by a centre wing, and the centre wing extends for approximately only half the longitudinal length of the aircraft.

6. A fixed-wing dual-fuselage aircraft comprising two fuselages, each fuselage has a respective outer wing extending from it, wherein the fuselages are spaced by a centre wing, and the centre wing extends for approximately the length of the aircraft minus the nose length.

7. A fixed-wing dual-fuselage aircraft according to Claim 6, wherein the centre wing extends from the inbound end of the aircraft nose to the rear of the aircraft.

8. A fixed-wing dual-fuselage aircraft according to any of Claims 5 to 7, wherein the outer wings are inclined upwardly where they extend from the fuselages.

9. A fixed-wing dual-fuselage aircraft according to any preceding claim, wherein the minimum distance between the fuselages is substantially equal to the diameter of a fuselage.

10. A fixed-wing dual-fuselage aircraft according to any preceding claim, wherein the aircraft comprises a tail fin mounted centrally on the rear section of the centre wing, and the tail fin extends upwardly from the centre wing.

11. A fixed-wing dual-fuselage aircraft according to any preceding claim, wherein the wings are a pair of rear wings, and the aircraft also comprises a pair of front stabilisers.

12. A fixed-wing dual-fuselage aircraft according to Claim 4, or any of Claims 9 to Ii when dependent on Claim 4, wherein the fixed- wing dual-fuselage aircraft is arranged so that a vortex can be generated about a nose of each fuselage.

13. A fixed-wing dual-fuselage aircraft according to Claim 4, or any of Claims 9 to 11 when dependent on Claim 4, or Claim 12, wherein the wings are dimensionally designed in the region of the root to allow the sustaining of a stream wise vortex at or adjacent each wing root.

14. A fixed-wing dual-fuselage aircraft according to Claim 4, or any of Claims 9 to 11 when dependent on Claim 4, or Claim 12, or Claim 13, wherein the sustaining means comprises two subsidiary wings, a first subsidiary wing being connected to the main wing part of a first wing, and a second subsidiary wing being connected to the main wing part of the second wing, and each subsidiary wing is inclined relative to its respective main wing part.

15. A fixed-wing dual-fuselage aircraft according to Claim 14, wherein the sustaining means is arranged to sustain streamwise vortexes contiguous junctions of subsidiary wings and their associated fuselages.

16. A fixed-wing dual-fuselage aircraft according to Claim 14 or 15, wherein the subsidiary wings are inclined downwards from the inbound ends of the associated main wing parts.

17. A fixed-wing dual-fuselage aircraft according to Claim 14, 15 or 16, wherein each subsidiary wing is of a length equal in diameter to approximately half the diameter of a body.

18. A fixed-wing dual-fuselage aircraft according to any of Claims 14 to 17, wherein the subsidiary wings are oriented at between 45 degrees to 75 degrees to the horizontal.

19. A fixed-wing dual-fuselage aircraft according to any of Claims 14 to 18, wherein the subsidiary wings extend outwards radially from the fuselage at angles of +30 degrees and -30 degrees respectively to the vertical plane through the longitudinal axis of the respective fuselage.

20. A fixed-wing dual-fuselage aircraft according to any of Claims 14 to 19, wherein the subsidiary wings are of aerofoil section.

21. A fixed-wing dual-fuselage aircraft according to any of Claims 14 to 19, wherein the subsidiary wings are of aerodynamic section.

22. A fixed-wing dual-fuselage aircraft according to any of Claims 14 to 21, wherein flaps are provided on the subsidiary wings for flight control.

23. A fixed-wing dual-fuselage aircraft according to any of Claims 14 1o22, wherein the main wing part, comprising the wing tip, is substantially longer than the subsidiary wing.

24. A fixed-wing dual-fuselage aircraft according to any of Claims 14 to 23, wherein the main wing part extends in a first plane and the subsidiary wing extends at least partly in a second plane at an angle to the first plane, the angle being between 105 and 135 degrees.

25. A fixed-wing dual-fuselage aircraft according to Claim 24, wherein the angle is approximately 120 degrees.

26. A fixed-wing dual-fuselage aircraft according to Claim 4, or any claim when dependent on Claim 4, wherein the sustaining means is arranged to fix stream wise vortexes at or adjacent the leading edge of the main wing parts of the two wings.

27. A fixed-wing dual-fuselage aircraft according to any preceding claim, wherein each wing has a pair of engines, and the engines are suspended from the wings.

28. A fixed-wing dual-fuselage aircraft according to Claim 27, wherein an aerofoil extends between the engines in each pair, and the uplift that the aerofoil creates thereby assisting in supporting the engines.

29. A fixed-wing dual-fuselage aircraft according to Claim 27 or 28, wherein each engine pair is suspended from its associated wing by a single central strut.

30. A fixed-wing dual-fuselage aircraft according to Claim 27 or 28, wherein each engine pair is suspended by two struts.

31. A fixed-wing dual-fuselage aircraft according to any preceding claim, wherein the wings are a pair of front wings, the aircraft also comprises a rear tail plane mounted on supports which extend upwardly from each fuselage.

32. A fixed-wing dual-fuselage aircraft according to Claim 11, wherein the rear wings and stabilisers have the same profile when viewed from the front.

33. A fixed-wing dual-fuselage aircraft substantially as described or shown herein.