CA2545629A1

CA2545629A1 - Aircraft

Info

Publication number: CA2545629A1
Application number: CA002545629A
Authority: CA
Inventors: Joerg Arnold
Original assignee: Individual
Current assignee: IP2H AG
Priority date: 2003-11-16
Filing date: 2004-11-15
Publication date: 2005-06-02
Also published as: PL1685024T3; ES2287795T3; ATE363429T1; DE502004003994D1; JP2007533528A; WO2005049422A2; BRPI0416619A; EP1685024B1; EP1685024A2; WO2005049422A3

Abstract

The invention relates to an aircraft comprising a fuselage (1) and a propulsion device (2) that is coupled to the fuselage (1) and that generates a definable lift. Said propulsion device (2) comprises several impeller blades (3) and the latter (3) can be pivoted through a pre-definable blade angle about a pivoting axis (4). The aircraft has been configured and developed in such a way that the impeller blades (3) are mounted to rotate about a rotational axis (5), the blade angle can be modified during the rotation of the impeller blades to generate lift and the respective pivoting axes (4) of the impeller blades (3) run substantially parallel to the rotational axis (5).

Description

AI RCRAFT
Various embodiments of the present invention relate to an aircraft with a fusE:lage and a propulsion device coupled with the fuselage for the production of a definable lift, whereby the propulsion device includes impeller blades and whereby the impeller blades can be rotated about an axis at a pre-determined blade angle.
Prior art aircraft of this kind exist in various forms and sizes. In particular, helicopters are well-known in which, through the rotation of one or more rotors about a substantially vertical shaft, a force (rotor thrust) is produced. The vertical component of this force provides lift for the helicopter.
Rotor' thrust can be derived from the vertical shaft by controlling the positioning of the impeller blades, which may produce a horizontal component of rotor thrust. This horizontal component of the rotor thrust serves as a propE;lling force that can also serve to move the helicopter sideways or backwards. The blades can be rotated about an axis to a pre-determined angle to produce a feathering effect. A helicopter's rotors are typically radially arranged in relation to their axis of rotation.
Various prior art aircraft have one or more rotors with two or more radially arranged rotors or blades each, each of which are fastened by one end to a rotation shaft and/or a rotor hub. The blades of the rotors circulate over a circular area, and the rotational axis of the rotors (which is perpE:ndicular to the circular area) is parallel and/or coaxial to the common vertical axis along the longitudinal or transverse axis of the helicopter's fuselage. The blades of the rotors are only rotated a few degrees off the vertical axis. This basic construction principle produces various flight-static and flight-dynamic disadvantages in known helicopters.
One disadvantage of prior art helicopters is that the fuselage of these helicopters can only be maneuvered forwards, backwards or sideways, and that i:his movement is coupled with pitch motions or rolling motions of the fuselage. Thus, it is not possible to maneuver the fuselage in all directions while maintaining the fuselage in a vertical orientation (i.e., without tilting the fuselage). When the helicopter is being maneuvered, at least two of the fuselage's situational axis are tilted. In contrast, a satellite can accomplish movement maneuvers in which all three situational axes of the satellite remain parallel. Such "shift maneuvers" are not possible for prior art helicopters.
A further disadvantage of prior art helicopters is that they include the so-called "overhead carrying-rotors" mentioned above, where the impeller blades, and the circular area through which the impeller blades travel, extend well beyond the front and sides of the helicopter. As a result, these helicopters must maintain an adequate distance from obstacles and accordingly can not dock with objects (e.g., to toad people or goods into the helicopter). The loading of people or goods into the helicopters can only take placE: from below in the fuselage. This limits the ability of the helicopters to servE; in rescue and salvage maneuvers.
Also, many prior art helicopters use only one overhead rotor. This results in a torque due to reactive forces that are caused by air resistance against the helicopter's blades. This torque seeks to rotate the helicopter about the vertical axis of the fuselage. A second rotor (for example, a tail rotor) is typically used to compensate for this torque. Such tail rotors are trouble-prone and are frequently the cause of helicopter crashes and crash landings.
Prior art helicopters are also problematic because, during flight, some of the helicopter's blades are moved against the aviation stream while others of the helicopter's blades are moved with the aviation stream, so that the air passes differently over the helicopter's various blades. Thus, changes in airspeed during directional flight changes the helicopter's flight dynamics.
It is especially the case that when helicopters are engaged in very fast flight, the movement of the impeller blades against the aviation stream causes the airflow to start from the front edges of the impeller blades. The rate of motion of impeller blades in the air is a combination of the circular path speed of the impE;ller blade and the speed of the aviation stream.
This limits the possible applicable combination of speed and carrying capacity of the helicopter and the applicable combination of number of blade revolutions and airspeed to a range within which the tips of the impeller blades are not yet in the supersonic speed range and, therefore, cannot be damaged by shock waves.
Blades with movement toward the aviation stream are flowed against -starting from the inside of the rotor circle - partially from the rear edge of the impeller blade. This applies to all parts of the impeller blades whose circular path speed rate toward the aviation stream is smaller than the flow rate of the aviation stream itself. With increasing airspeed, these blades contribute increasingly less to the helicopter's lift and produce an airspeed-dependent rolling moment on the helicopter flight cell and/or on the fuselage, which must be accounted for in the performance of the helicopter.
This problem leads to the limitation of the typical maximum speed of prior art helicopters to around 400 km/h. It also causes energy expenditure to increase as the airspeed of the helicopter increases, which is not favorable to the helicopter's airspeed or carrying capacity. Today's helicopters are therefore very energy-inefficient in their flight performance, and therefore typically only have flight ranges of about 1000 km.
Helicopters are controlled by adjusting the impeller blades' angles of incidence and with some experimental helicopters additionally by tilting the helicopter's rotor axis (e.g., axis of rotation). Unfortunately, since the impeller bladEa must be adjusted both cyclically and collectively, a complex swash plate control and a complicated rotor head construction are necessary. This complicated construction does not currently permit providing rotor heads with any more than 8 blades or load-carrying capacities of the rotor heads over 60 tons.
Typical helicopters are in principle pendulums, in which the fuselage hands from, and oscillates under, the rotor head. The flight attitude of the fuselage is dependent on the dynamic flight condition (e.g., whether the helicopter is engaged in forward -, backwards -, sideways or hovering flight).
The flight attitude of the fuselage can not be set independently from the dynamic flight condition - for example, it is impossible for a helicopter to jackknife. However, there have been attempts to experiment with helicopters that have inclinable or tiltable rotor heads. However, these have resulted in still more fragile and complicated drive constructions.

An object of the present invention is to provide an aircraft in which at least one of the above problems is eliminated and which is realized through a simple construction.
According to the invention, the task at issue is solved with an aircraft according to patent claim 1. An aircraft with these characteristics is equipped and developed so that the impeller blades are situated in a way that they can be rol:ated about a rotation shaft. While the impeller blades are being turned, the impeller blade angle is also adjustable to produce lift. The respective pivot axes of the impeller blades are also positioned substantially parallel to the rotation shaft.
In an inventive manner, it was recognized that the type of aircraft referE;nced above does not have to be a helicopter equipped with an overhead carrying rotor, where the impeller blades - and thereby also the pivot axes of the impeller blades - are substantially radially arranged about the helicopter's rotation shaft. Furthermore, it was recognized that a particularly simple construction of the propulsion device is viable due to the fact that the impeller blades are situated so that they can be turned about a rotation shaft, whereby the rE;spective pivot axes of the impeller blades are positioned substantially paralilel to the rotation shaft. In other words, the respective pivot axes and the rotation shaft are positioned in such a way that the impeller blades move in a parallel manner about this same rotation shaft while they turn. The blade angles can be changed while the impeller blades are being turned about the rotation shaft to generate a controlled lift. The force and direction of the thrust depends upon the setting of the impeller blade angle.
According to an inventive embodiment of the aircraft, it is, in particular, not necessary to use overhead rotors, which usually extend far beyond the aircraft's fuselage and therefore make accessibility to the fuselage more difficult and also prevent the possibility of docking the fuselage near, for example, a building.
The respective pivot axes of the impeller blades may be positioned substantially equidistantly to each other. This makes it possible for the impeller blades to have an even and balanced course of motion about the rotation shaft. And, along the same lines, the pivot axes of the impeller blades may be positioned so that each is equidistant to the rotation shaft.
The pivot axes of the impeller blades may be positioned not just substantially parallel to the rotation shaft but also substantially parallel to each other. Accordingly, an altogether homogeneous and quasi symmetrical embodiment of the impeller blade arrangement about the rotation shaft is feasik>le.
For a particularly simple and safe calibration of the angle of the impeller blades, the impeller blades' pivot axes may be positioned based on the centroid of the impeller blades. The pivot axis may run exactly through the centroid of the cross sectional profile of the impeller blades.
A neutral position of the impeller blades, where the impeller blades do not produce a thrust or air diversion during their rotation about the rotation shaft., can be produced with a concave curve of the impeller blades' cross-sectional profile in relation to the rotation shaft. The blades' cross-sectional profiles may lie almost completely within a cylindrical wall of an imaginary circular cylinder. Such a turning circular cylinder would produce no thrust and no aiir diversion.
The pivot axis of each blade may stand out perpendicularly from the impeller blade's cross-sectional profile and, thus, run quasi parallel or coaxially to the impeller blade's longitudinal axis.
To safely control and swivel the impeller blades about the pivot axis, the irnpeller blades may include a control shaft at at least one end. These control shafts would each serve as the point of contact for swiveling the impeller blades about the pivot axis. The control shafts may extend perpendicularly from the impeller blade's cross-section profile and be positioned in front of or behind the pivot axis - as seen from the impeller bladEa' direction of rotation about the rotation axis. The blade may be linked, and respectively the impeller blade's angle of attack (or blade angle) may be adju:>ted, by the control shaft. Both positive and negative blade angles can be employed - relative to the neutral position of the impeller blade. As mentioned above, the neutral position of the impeller blade means that when the impeller blades rotate about the rotation shaft, no standing air is diverted from i:he impeller blades. Rather, the impeller blade simply cuts through the air. The distance between the control shaft and the pivot axis defines the translation ratio as the impeller blade angle is adjusted.
For particularly safe storage and movement of the impeller blades, the impeller blades may be rotatably attached at one end to a drive component, or rotatably mounted within the drive component. Thus, the drive component may Ibe rotated in a structurally simple way about the axis of rotation, or rotatably positioned on the axis of rotation. To this end, the drive component may include a bearing shaft or a hollow shaft, which may be aligned with the outer' side of the impeller blades.
In a relatively simple embodiment, the drive component may be constructed as a drive pulley, drive disk or a drive ring, to which the impeller bladEa are pivotably attached.
The pivot axes or blades may be positioned perpendicular to the drive component, the drive pulley, the drive disk or the drive ring. And the impeller blades or pivot axes may be positioned on opposing sides of the bearing shaft. Furthermore, the pivot axes may be arranged in a circle along the edge of the drive component, the drive pulley, the drive disk or drive ring. The pivot axes. are preferably positioned so that they are spaced equidistantly apart.
The arrangement of parallel blades may thereby form a cylindrical rotor assE:mbly.
In principle, there may be as many blades as desired positioned on the drivf: component or on the cylindrical rotor assembly, depending upon the diameter of the drive component and the width of the impeller blades. Each blade may be positioned so that its pivot axis is perpendicular to the drive component and so that it pivots about is pivot axis.
The bearing shaft or hollow shaft of the drive component may extend perpendicularly from the drive component or from the surface of the drive pulley and/or drive disk. To assure that the drive component operates safely, the drive component may be coupled with a toothed belt, a chain or a circular toothed gear. For this, the drive component may include a circular tooth arrangement adjacent the circumference or circular edge of the drive component or on the perimeter of the bearing shaft. Thus, the bearing shaft may I~e constructed in the form of a drive shaft.
In order to safely couple the impeller blades to the drive component, the drive component may include recesses or passages for the storage of the impeller blades' pivot axles. Alternatively or additionally, the drive component may include recesses or passages for the impeller blade's control shafts. The control shafts may be dimensioned to extend through the recesses or passages in the drive component. The recesses or passages may be constructed as cut-outs, holes or slots. In particular, the recesses or passages for the control shafts of the impeller blades may be constructed as long, preferably curved, holes.
In order to limit the weight of the drive component, the drive component may include slots, recesses, passages, cut-outs, holes or slivers, so that the drive component may have a star shaped, circular or spoke-like appearance.
In order to safely adjust the angle of the impeller blades, the drive component may work in conjunction with a control member to adjust the impeller blades about their pivot axis. The control member may be exclusively responsible for the attitude of the impeller blade angle by means of movement of the control shafts. The control member may thereby be decoupled from the rotation of the impeller blades and/or the drive component. In other words, the control member does not rotate with the rotation of the impeller blades about the rotation axis. In a relatively simple embodiment, the control member may be mounted on the axis of rotation.
In order to safely control the pivot shafts, the control member may include a cyclical gear. In principle, the control member may be adjustable in a guide relative to the rotation shaft in order to achieve a safe setting or presetting of the impeller blade angle. The control member may also be mounted and/or directed so that it can be shifted a certain distance or amplitude in all directions perpendicular to the rotation axis.
In a particularly simple embodiment, the guide may include two perpendicularly-arranged linear guides, for the purpose of such cross-table guidance. Alternatively, and in an equally simple embodiment, the guide may _7_ include a rotational guide member that is connected to a linear guide member (e.g., in the form of an extendable and pivotable guide).
As another preferred embodiment, the control assembly may include two rotational control portions in the form of a double eccentric disk control member. The elements of the control member specified above may be described as control disks. The eccentric disk member is advantageous in that it can be positioned directly on, or be supported by, the bearing shaft or hollow shaft of the drive component.
in order to independently and safely control or move the eccentric disks of the eccentric disk control member, each eccentric disk may be associated with an actuator. In particular, two eccentric disks of the eccentric disk control member may each be respectively associated with an actuator.
An eccentric disk control member may include two eccentric disks: an inteUnal eccentric disk, which may have its eccentric cam hole mounted via bearings on the bearing shaft of the drive component; and an outside eccentric disk, which may be mounted on bearings around or adjacent the outside of the internal eccentric disk. The eccentric cam hole of the outside eccentric disk may contain the internal eccentric disk. The control member may be (or run) mounted on bearings on the outside eccentric disk, and/or may be centrally positioned in relation to the outside eccentric disk. In such an arrangement, both eccentric disks can turn freely about each other and/or within one another. If the eccentric disks become twisted, the control member is Engaged to address the issue. The eccentricities of the eccentric disks are selected in such a manner that for a relative angle attitude of the eccentric disks, the fulcrum of the outside eccentric disk corresponds to the fulcrum of thE; bearing shaft of the drive component. If the eccentric disks in this relative angle attitude run smoothly with respect to each other while turning about the bearing shaft of the drive component, the condition of the control member remains unchanged and in a disengaged orientation.
The eccentric disk control member can be used to produce a controlling effect as follows. When the control member is disengaged, due to the mutual angular orientation of the eccentric disks, while not engaging each other, the eccentric disks are rotated into a particular angle that corresponds _g_ to the desired direction of deflection of the rotor assembly's blades.
Subsequently, both eccentric disks are rotated against each other in a manner in which the outside eccentric disk rotates twice as far about an angle as the internal eccentric disk. Accordingly, the outer eccentric disk rotates twice as fast as the internal eccentric disk. Thus, a deflection is produced in the de:>ired direction that is proportional to the eccentric disk angle of rotation -azimuth angle. This double eccentric disk control is thus a vector control in which the deflection direction of the deflection angle is first set and then the amount of the deflection angle is set. Each control command and/or each pre-determined control position - "control stick position" - may be assigned a deflection and/or angle of direction of deflection).
The control may now be done in such a manner that the control is divided into sequential discrete controlling positions. There is a gradual change-over from one controlling position successively into the next controlling positions, as the assigned deflection angle transfers over to the ne~;t assigned deflection angle and assigned angle of amount of deflection transfers over to the next assigned amount of deflection.
The more precise the selected discrete breakdown is, the more precise, and/or simultaneous the available control. The eccentric disks may be adjusted thereby by two actuators, for example two stepper motors. One actuator holds and at the same time rotates the internal eccentric disk and onE~ actuator holds and at the same time rotates the outside eccentric disk.
In addition, each eccentric disk may be provided with a gear rim, in which the actuator may engage a pinion. The eccentric disks do not move while the impeller blades are rotating outside of the control procedure.
In order to safely mount and/or guide the control shafts, the control member may include an annular groove or a circular groove for receiving the impeller blades' control shafts. While the impeller blades are rotating about the rotation axis, the control shafts may revolve within the annular groove or circlular groove. In a further simple embodiment, the control member may be constructed as a control ring or control disk. With this embodiment, an annular groove or a circular groove may be formed in the outer area of the control ring or the control disk.

Now if the control member is moved in a particular direction by the guiding of the control member while the drive component is rotating, then the impeller blades' control shafts (which may revolve in an annular or circle groove) will cyclically follow this deflection. This creates a cyclical blade adjustment. With one revolution, the impeller blades' control shafts go from their neutral position and move into the maximum positive position one time, anti also move into their maximum negative position one time. Between these two extreme deflections, the impeller blades' control shafts pass through their neutral position twice. In the two neutral positions, which are across from each other on the control member's annular groove path, standing air is not diverted from the impeller blades. Because of the impeller blades' direction of motion reverses along the circular path, the standing air is maximally diverted in tl'ne same direction at the two extreme positions.
The extreme positions of the impeller blades are in positions on the shiit axis or in the control member's direction of movement. As a result, the neutral positions are present at positions that are respectively shifted 90 degrees. If the impeller blades' control shafts are positioned in front of the impeller blades' pivot axes at the impeller blade (as seen from the direction of the impeller blades' rotation), then the control member's shift direction is identical to the thrust direction, which is predetermined by air diversion. If the impeller blades' control shafts are positioned behind the pivot axes, then the reverse effect results.
In order to increase the efficiency of the propulsion device, the annular groove or the control ring may deviate from a circular arrangement. Thus, it is not mandatory to have a circular arrangement of the annular groove or the control ring. For example, other arrangements may provide an angle of incidence function that is dependent on the angle of rotation, or superposed angle of incidence functions. Such techniques may be used to change the efficiency of the propulsion device. Furthermore, the angle of incidence function may be proportional to the expression a-cos(x)"', in which "a" is the degree of the impeller blades' angle of incidence and "w" is preferably a whole nunnber, and preferably 11.

In other words, the form of the control ring or the annular groove may be further optimized, in order to increase the efficiency of the impeller blades rotating about a rotation shaft or the efficiency of a cylindrical rotor assembly.
To this end, an alternative embodiment may be provided that supplies an angle of rotation-dependent angle of incidence function, a-cos(x)'". This may be used in place of a circular annular groove or in place of a circular control ring. It is only one example of the ability to optimize the efficiency of the impeller blades rotating about a rotation shaft or the cylindrical rotor assembly.
In reality, annular groove forms are more suitable. They supply superposed angle of incidence functions, that then optimize the efficiency of the cylindrical rotor assembly or efficiency of the impeller blades rotating about a rotation shaft without detracting from the maximum cylindrical rotor assembly or blade thrust.
Such annular groove forms or control ring forms or blade trajectories can be produced from the overlay of a circle - basis circle - with two or four periodic and symmetrical "projections" on the circle. This is done so that the curves of the possible overlapping functions of the basis circle can be inscribed and a square can be circumscribed. Thus, a host of curves is then present. In an informal approximation, it would be conceivable to think of a squiare around the circle.
Simple examples of such overlapping curves are as follows: for two buicaing (convex) ellipses and for four convex epicycloids and asteroids or simple squares or rectangles with rounded corners. If non-circular control rings or annular grooves are used that have a strong deviation to the circular form, the controlling of the impeller blades or the cylindrical rotor assembly should be altered. It would be possible, then, to use two control members that are next to or in front of one another - or to use slot rings of the same or sliglhtly different form may be used at one or both ends of the cylindrical rotor assembly. The blade pivot axles - blade axles of rotation - run around inside the grove ring that is situated next to the drive disk and/or the guide disk and is fixed relative to the disk. Running inside the other slot ring - the control ring - are the impeller blade control shafts or blade coupled shafts. This slot ring can still be rotated and shifted relative to the drive disk or guide disk.

Instead of drilled holes for seating the impeller blade pivot shaft, the drive disk and the guide disk are additionally equipped with radial slots or long holes in which the impeller blade drag pivot shafts can slide radially. The drive disk and the guide disk then have the character of a drive/attachment disk. In this embodiment, which includes pairs of control members and/or slot rings, both the impeller blade pivot shafts and the impeller blade control shafts may be controlled and moved radially. It is apparent then, that in place of a circle-cylindrical rotor assembly or blade arrangement, the impeller blades may move on a cylindrical rotor assembly that may be, for example, ellipsoid, epicyclical, asteroid, or a cross-sectional path of a square with rounded corners or a rectangle with rounded-corners.
In a simple embodiment, the bearing shaft or hollow shaft of the drive component may preferably be configured to run centrically through the control member. The control member may then be positioned concentrically behind or under the drive component. The bearing shaft of the drive component, which can simultaneously function as the drive shaft, may be passed through the control member. The drive component and the control member may thus be positioned parallel to each other. In one embodiment of the drive component as drive disk and of the control member as control ring, the circular area of the drive disk may be positioned parallel or co-planer to the plane in which the control ring lies.
When the drive disk is rotating, the ends of the impeller blades' control shafts may run free of play in the control member's annular grove or circle groove. In this embodiment, the control shafts are inserted through the drive component or the drive disk. This lack of play can be achieved, for example, by an appropriate pulley support, in which the ends of the control shafts are disposed. In the simplest embodiment, this pulley support may be created by means of two radially, easily transferred, antifriction bearings, which are disposed on the end of the control shaft. One antifriction bearing maintains pressurized contact with one of the annular groove's inner walls and the other antifriction bearing maintains a pressurized contact with the other, or opposite, inner wall of the annular groove.

The cyclic adjustment of the impeller blades may alternatively be crE;ated by means of known cyclical gears for paddle drives derived from marine propulsion technology - for example, the Schneider-Voith drive. These well-known construction principles are less suitable however for fast rotating rotor assemblies, since the paddle control mechanics control masses of high inertia, whose cyclic acceleration leads to high reaction forces and vibrations.
The invention however ensures minimal ground acceleration, since the impeller blades only accelerate cyclically about their longitudinal axis and no other control mechanism must be cyclically accelerated.
For particularly stable flight characteristics, the propulsion device may include at least two arrangements of rotatable impeller blades about a rotation shaft. Thereby unwanted torques about the vertical axis of the aircraft can be avoided.
In order to produce a safe lift, the rotation axis or axes may be po:>itioned in a substantially horizontal plane. This makes it possible for a maximum thrust conversion in a vertical direction.
In a particularly narrow construction of the aircraft, the rotation axis or axEa may be positioned parallel to the fuselage's longitudinal axis along the forVVard flight direction. In an alternative embodiment, the rotation axis or axes may be positioned perpendicular to the fuselage's longitudinal axis along the forward flight direction. In principle, both arrangements of the rotation axis or rotation axes as described above are favorable to the aircraft's flight stability.
In a particularly simple embodiment, impeller blades may form a rotatable rotor about, respectively, one of the rotation axes, whereby the propulsion device may include at least two such rotors.
In order to produce stable flight attitudes, the rotors may be staggered along the longitudinal axis. At least one such rotor may be positioned on each longitudinal side of the fuselage. It is however also conceivable that there may be a plurality of rotors on each longitudinal side of the fuselage. In an embodiment with a plurality of rotors, a stronger lift may be produced, whE;reby heavier loads would be transportable by the aircraft.

In order to avoid unwanted torques, at least two rotors may be rotatable in opposite directions.
In a simple embodiment, at least two rotors may be positioned on each of t:he longitudinal sides of the fuselage, and the rotation axes may align themselves across from the rotors. Thus, finally an arrangement of the rotors is realized with rotational axes that are perpendicular to the fuselage's longitudinal axis in forward flight direction.
For versatile and individual control of the aircraft, each rotor may be separately controllable. In the case of a simplified control, rotors may be controlled together in the same way.
Due to the working principles of a rotor, which operates by rotating blades about a rotation axis, a lift and/or propulsion force is generated in a direction that is perpendicular to the longitudinal axis of the rotor and/or rotational axis. The rotors (as drives) and/or their rotation axes may therefore be positioned parallel and/or coaxial to the transverse axis of the aircraft at the fuselage. However, in order to achieve a statically desirable flight condition and to achieve a particularly stable system, at least two rotors should be used.
ThE: rotors must be staggered in the direction of the longitudinal axis of the fusE:lage - with one on each side of the fuselage. In order to avoid unwanted torques, the two rotor assemblies may rotate in opposite directions. The two rotors, located at the same height of the longitudinal axis, may be positioned at the centroid of the aircraft. Depending upon whether the rotor assemblies turn in the same direction or in opposite directions, a torque may be produced about the transverse axis or about the vertical axis of the fuselage. In such a case, the necessary static stability of flight attitude would not be provided.
When considering the possibility of losing the use of a rotor assembly, a particularly safe embodiment of the invention includes the use of four rotor assemblies. Two rotor assemblies may be provided on both sides of the fuselage, across from each other and/or may form a common rotor assembly longitudinal axis and/or rotational axis. Such a pair of rotor assemblies may each be envisioned for the front and rear portions of the fuselage.
By using two and/or four rotor assemblies in the direction of the longitudinal axis of the fuselage, the aircraft can execute forward and backwards maneuvers, without pitching. To do this, the thrust production of the front and back rotors or rotor pairs would have to be controlled evenly.
On them other hand, a forward and/or backwards maneuver may be accomplished with a varying thrust from the front or rear drive. This then leads again to a torque about the transverse axis of the fuselage, and thus to the well-known pitch motion, by which under force vector analysis the sub-vectors have components in a forwards or backwards direction.
The same that is said of forwards and backwards maneuvers can be said for sideways maneuvers. Lateral maneuvers may be accomplished by varying the thrust of the left or the right rotor assemblies and/or the left or the right drive. A torque thus results about the longitudinal axis of the fuselage and, from the ensuing rolling motion, lateral-directed thrust vectors are then produced.
In a further embodiment, a lateral movement of the aircraft may be produced without a resulting rolling motion. For this, the impeller blades are fE~atherably mounted, by means of a pivot shaft, at their end that is away from the drive component, at or in a guidance device. Such a guidance device can absorb bending moments of the rotary blades that arise from cyclical air diversion and the centrifugal energy of the rotation. The arrangement of impeller blades or the cylindrical rotor assembly can rotate with substantially higher numbers of revolutions with the guidance device.
The guidance device can also absorb substantially more thrust reaction forces andlor produce substantially more forward thrust and lift thrust.
The guidance device essentially may be constructed like the drive component - preferably disk-shaped. In doing this, a guidance device may be in the form of a guide disk.
The guidance device may be supported by a central support shaft or a hollow :haft between the drive component and the guidance device and coupled to the drive component. Therefore, the guidance device may configured to rotate with the drive component. The guidance device and the drive component may be coupled by means of a shaft or a rotation shaft. The longitudinal axis of the support shaft, hollow shaft, axis or rotation axis may run through the center of the drive component and the guidance device, - IS -whereby the longitudinal axis of the support shaft or hollow shaft coincides with the rotor assembly's longitudinal axis and/or the rotation axis.
The support shaft of the guidance device or the guide disk corresponds to 'the bearing shaft of the drive component or the drive disk. One can regard the support shaft as guidance device-lateral extension of the bearing shaft of they drive component. Since the support shaft or hollow shaft carries and/or propels the guidance device, it may also function as the drive shaft for the guidance device.
In order to produce a particularly simple lateral movement of the aircraft without a resulting rolling motion, the guidance device may be provided with a guidance rotor that includes rotor blades. The rotor blades may be linked to the guidance device in a structurally simple way. The rotor blades may be radially positioned between the hub of the guidance device and the edge of the guidance device. For this, the guidance device may include appropriate pa~;sages and/or bearing seats. Such a guidance rotor may correspond to the tail rotor of conventional helicopters and be structurally similar in principle.
The guidance rotor may be propelled about a linkage running through the rotation axis. Such a linkage may include a connecting rod, which is guided by the support shaft and the bearing shaft of the drive component.
They guidance rotor may be attached by connecting rod and by bellcranks.
Guidance rotors of the rotor assemblies make it possible for the aircraft to make lateral movements without a resulting rolling motion. In addition, all guidance rotors may be controlled the same with respect to their thrust.
A turning maneuver of the aircraft about the vertical axis of the fuselage may be accomplished either by an uneven thrust of the rotor assemblies or the guidance rotors. In addition, the thrust would only need to come from two rotor assemblies or guidance rotors that are on opposite sides and not belonging to the same pair of rotor assemblies.
To cause the aircraft to ascend or descend, the lift thrust of the front and the rear drive may be evenly controlled. An increase of climbing thrust may be achieved by increasing the impeller blade angle and/or by increasing the number of revolutions of the rotor assemblies. This is where a substantial difference between this invention and conventional helicopters becomes obvious. Conventional helicopters produce their climbing thrust over a collective increase of the impeller blade angle and/or increase in the number of revolutions of the rotor. With the present invention, however, the climbing thrust is attained by means of the cyclic blade control.
In order to control the fuselage's flight attitude and/or trim, the thrust of thE; front or rear and/or lateral drives may be controlled differently. The invention makes it possible to maintain the trimming of different current flight attitudes independently of the aircraft's dynamic flight condition.
One embodiment of the invention can achieve the same flight-dynamic conditions of a conventional helicopter without the use of a collective blade control. The invention therefore involves a substantially simpler construction.
In addition, this embodiment of the invention can accomplish maneuvers with completely decoupled movements and/or pure translation movements or shift maneuvers without associated pitching and/or rolling motions. Such maneuvers or movements are not possible for conventional helicopters. In addition, the flight attitude of the fuselage can be trimmed. The helicopter can stably assume all positions within the full 360 degrees about the transverse axis of the fuselage. Since the aircraft according to an embodiment of the invE~ntion can have two or four rotor assemblies with essentially an arbitrary number of impeller blades for each rotor assembly, a substantially higher lift thrust can be achieved and can therefore transport substantially higher payloads than is possible with conventional helicopters.
In one embodiment, where the rotor assembly's longitudinal axis and/or the rotation shaft is parallel and/or coaxial to the transverse axis of the fuselage, there is still an apparent disadvantage - the airflow blows against the impeller blades that are moving against the aviation stream and those that are moving toward the aviation stream in an inefficient and varying way depending on the number of revolutions and airspeed. In addition, there is still the disadvantage with this construction that the airspeed is added to the path speE~d of the impeller blades so the maximum airspeed remains very limited.
These disadvantages can be fixed with the following embodiment of the invention. For this, the rotor assembly's longitudinal axes and/or the rotation axe:. are no longer positioned parallel and/or coaxial to the fuselage's transverse axis. Instead, they are positioned parallel and/or coaxial to the fuselage's longitudinal axis. One such propulsion device may then employ at least two rotor assemblies behind one another with a shared rotor assembly longitudinal axis and/or rotation axis to produce the necessary maneuvering torques. When using only two rotor assemblies, the two rotors may be positioned behind one another and overhead, over the fuselage. When using four rotor assemblies, the four rotors may also be arranged laterally, with two rotors behind one another on each side of the fuselage. Each pair of rotor assemblies located behind one another, next to each other, and/or across from each other can then rotate in opposite directions to avoid unwanted torques.
In a particularly efficient embodiment of the invention, blades or rotors that are positioned next to, or behind, one another may be positioned in a quasi mirrored arrangement. One can also speak here of a connected and reflected "behind one another" arrangement of rotor assemblies. For this embodiment, only one guiding element is necessary for both rotor assemblies.
And it is sufficient to only drive one drive component, since the rotor assemblies can be coupled over a common guidance device and/or the reciprocal support axes can be firmly coupled with one another. However, in this embodiment having only one drive, the maneuvering torques may only be achieved for both rotors by the cyclical blade adjustment and no longer by controlling the number of revolutions. Aside from that, the torques, which result from air resistance against the rotor blades, can no longer be compensated for by a separate, opposite propelled rotation of the rotor assemblies.
A more extensive simplification of the aircraft according to an embodiment of the invention may be achieved because at both ends of the impE~ller blades each control member is respectively operable independently from each of the other control members. For example, only one rotor assembly may be used, but that rotor assembly of course includes two control members or control rings. The blades may be attached then at each of their ends either to a drive component or to a control device and have each of their ends. guided in control members or in annular grooves. In this application, maneuvering torques, which run perpendicularly to the rotor assembly longitudinal axis and/or rotation axis, may be achieved by a cyclic torsion of the impeller blades. The torsion takes place via a relative shift of the two control members or control rings in relation to each other. Thus, the impeller blade angle and the selective thrust of the impeller blades constantly changes from one end of the impeller blades to the other end. In the previously described embodiment of the aircraft, maneuvering torques are produced by different gross thrust vectors of the individual rotor assemblies. A rotor assembly with torsion control works like two separate controllable rotor assemblies.
The correlation of the thrust production to the flight maneuvers changes with the alignment of the rotor assemblies' longitudinal axes which are parallel and/or coaxial to the longitudinal axis of the fuselage. Lateral shift maneuvers or turning maneuvers about the vertical axis of the fuselage can no longer be accomplished by means of the guidance rotors, rather they can now only be accomplished by means of the rotor assembly's thrust control. However, fonrvard directed shift maneuvers can no longer by means of the rotor assemblies, rather they can now only be executed through the guidance rotors. The usual way of attaining aircraft propulsion through coupled pitch motion may be avoided. Instead, in a more energy efficient way of attaining aircraft propulsion, the relatively weak guidance rotors (in the context of a true translational movement) may be replaced by strong variable-pitch propellers.
In this embodiment of the invention, a drive train including drive components of the may comprise: a variable-pitch propeller; a shaft powered turbine with fuselage attachment; reciprocating turbine shafts; a front control mennber or front control ring; a drive component or drive disk of the front rotor assE:mbly; front parallel blades; a guidance device or guide disk with reciprocating support shaft; rear parallel blades; a drive component or drive disk of the rear rotor assembly; a rear control member or rear control ring;
and rear bearing shaft intake with fuselage attachment of the drive component or the drive disk of the rear rotor assembly, whereby the variable-pitch propeller sits in front on the turbine shaft, and the bearing shaft of the first drive component or the first drive disk is also coupled with the operated turbine shaft behind the turbine. The turbine in the drive train replaces the otherwise still necessary bearing shaft intake with a fuselage attachment for the front drive component or the front drive disk.
Alternatively, the drive turbine or turbines may be positioned in or on the fuselage and propel the drive train via a transmission. In this alternative embodiment, a drive turbine may propel two lateral drive trains or, when using two drive turbines, these may simply be coupled with a transmission in order to prevent the Loss of a drive turbine.
In an embodiment of the invention, at least one drive turbine may be po sitioned in the fuselage of the aircraft. This would thereby ensure a protected arrangement of the drive turbine.
The arrangement of the drive turbine or turbines in the fuselage would have the further advantage that the turbine exhaust gases may be laterally dirf;cted from the fuselage, directly to or over the rotor. The negative pressure self-produced over and from the rotors may thereby become partly balanced by inflows of the turbine exhaust gases. The necessary self induced drive power somewhat decreases the expenditure of drive power. The hot turbine exhaust gases may prevent a possible freezing of the rotor assembly and, at the same time, the hot turbine exhaust gases may be swirled in such a manner and diverted downwardly so that they cannot reach any other turbine inlets of other turbines and lead to the breakdown of other turbines.
The variable-pitch propeller would have to be laid out in such a way that both forward and backward thrust can be achieved by means of its propeller blade adjustment. In order to further reduce the induced power expE:nditure, a variable-pitch propeller or a propeller may be positioned both in front of a rotor assembly, as well as behind this rotor assembly. The front propeller may be laid out then as a draft propeller, and the rear propeller as a pressure propeller. Beiween the variable-pitch propellers or propellers - for example between the draft propeller and the pressure propeller - two or more rotor assemblies may be positioned behind one another.
In the hovering flight of the aircraft, one may operate the two propellers above so that they push against each other. Both propellers may be adjusted for hovering flight in such a way that their propulsion effect mutually compensates for itself, but additional air mass of the rotor assembly or the respective rotor assemblies is nevertheless applied.
The propeller can also be used to partially compensate for the torque.
It substantially compensates for a remainder torque, that comes from the rotation of the rotor assembly and/or reactive air resistance. This torque, which is generally uncompensated, arises for example when using only one torsion-controlled rotor assembly - see above. In addition, the propeller rotation can be set to rotate in the opposite direction of the rotor assembly rotation by means of a reversing gear mechanism. This torque reconciliation is particularly interesting for smaller aircraft according to an embodiment of the invention that has only one drive train.
In a further embodiment of the invention, at the fuselage, one could include at least one airfoil or auxiliary airfoil to which the rotor assembly or assemblies are attached, or from which the rotor assembly or assemblies are hung.
A torque that comes from air resistance that arises or is uncompensated, does not have the same serious effect on the aircraft according to the embodiment of the invention as it does on a conventional helicopter. In a conventional helicopter, this torque must absolutely be compensated, usually by a second rotor, for example a tail rotor, in order to prevent a continuing rotation of the fuselage about its vertical axis. In one embodiment of the invention, such torques about the vertical axis do not arise.
Corresponding torques arise only about the longitudinal axis of the fuselage and all lead at the most to a lateral pendulum deflection of the fuselage. An aircraft according to an embodiment of the invention is flight-statically substantially more stable than conventional helicopters.
Due to the parallel and/or coaxial orientation of the rotor assembly's longitudinal axes to the longitudinal axis of the fuselage, the speed component of the helicopter's drive speed is canceled along with the rotor-path speed during the forward movement of the helicopter. This is because both speed components are perpendicular to each other.
The aircraft according to an embodiment of the invention makes it therefore possible to attain substantially higher maximum airspeeds than conventional helicopters. In principle, the maximum airspeeds of turboprop aircrafts are possible also with the aircraft according to an embodiment of the invention. There is also the expectation that the aircraft according to an embodiment of the invention can also fly even faster than a turboprop aircraft having the same drive power, because the aircraft according to an embodiment of the invention does not include an airplane tail unit or airplane wings, which produce substantial additional air resistance, which is not the case with the aircraft according to an embodiment of the invention. The aircraft according to an embodiment of the invention has the flight dynamics of both a conventional helicopter and a conventional airplane and can therefore be flown (from the perspective of flight-dynamics) like a helicopter or an airplane. The flight operation of the aircraft according to an embodiment of the invention is substantially more energy-efficient than conventional helicopters, since no increases in rolling moments arise with increases in airspeed. Thus, the aircraft according to an embodiment of the invention avoids all well-known disadvantages of conventional helicopters specified in the introduction. In addition, because of the ability to arrange the rotor assemblies parallel and/or coaxially to the longitudinal axis of the fuselage, two or more rotor assemblies may be positioned laterally behind one another and laterally next to one another adjacent the fuselage. Thus, large fuselage constructions with load capacities of around the 200 tons, or alternatively of more than 200 passengers, are possible.
Docking maneuvers and thereby also difficult transport, rescue and salvage maneuvers are possible with the aircraft according to an embodiment of the invention, because of the aforementioned advantages and because of the lack of over-hanging, overhead carrying rotors. For difficult transport, rescue and salvage maneuvers, the fuselage may be equipped with a docking assE:mbly, which may be used for the loading or unloading of transported goods and/or to allow people to embark or disembark from the aircraft. In an embodiment of the invention of relatively simple construction, the docking assembly may include a tunnel, a bridge or a basket. Based on easily-obse~rved docking assemblies of aircraft cockpits, the docking assembly may be positioned at the front end of the fuselage.

The aircraft according to invention could be equipped, for example, with an escape tube or a catch box at the nose of the aircraft, which may be used to allow rescued people, animals or goods to enter the aircraft. It may also facilitate the deplaning of auxiliary or rescue forces and/or goods after docking. This is an enormous advantage, because, for example, in multistoried buildings, the auxiliary and rescue forces may not use elevators for safety reasons and, if necessary, are therefore forced to transport equipment between many floors using the stairways.
Regarding safe docking of the aircraft with, for example, a building, a docking assembly could include a preferably funnel-shaped guide. A guide of this type may be attached to a building and be configured for coupling to an aircraft for use as an entryway into the aircraft. This could facilitate a useful docking of the aircraft.
Regarding a particularly stable coupling of the aircraft to, for example, a building or to the entryway, the docking assembly could include a locking mechanism. Such a locking mechanism could include, for example, male locking device members on the docking assembly and female locking device mernbers on the entryway.
In other words, an aircraft could include a docking assembly, which fits exactly into guides or entryways or locking devices that are intended for this purposes. These guides, entryways, or locking devices may be disposed, for example, in the escape windows, emergency doors, escape portals, or any other form of emergency exits that are on the outside of, for example, tall buildings. The aircraft can dock and anchor itself at these places, open the appropriate emergency or escape exits, load people, animals, or goods, and then depart with the loaded people, animal, or goods on board.
The docking assembly may include a funnel-shaped guide that is adapted to entirely receive a docking nose. When the docking nose is slid into the funnel-shaped guide with its male locking members, the docking nose can be guided into the female locking members found at the end of the funnel shaped guide. The female locking members may, for example, be in a symrnetrical 3-point arrangement or in a 4-point arrangement around the perimeter of the entry opening at the end of the funnel-shaped guide. The locking and/or unlocking of the locking mechanism could be operated by the inventive aircraft electromechanically, or mechanically through connecting rods. The locking device may be, for example, a hook lock, a spreading locking device, a channel lock, a bolt lock, or a transverse rod lock in the form of, for example, a roller, pin, or fork lock.
The entryway of the funnel-shaped guide can be closed by a hatch, a door or a windowpane, which can be opened from the outside. The guide could be transferred into the inside of a building so that it bindingly locks outside of the building with the building's front and is therefore not visually offE:nsive. In high-rise buildings such guides could be, for example, installed on all sides of the building and could be installed, for example, on every pre-determined number of floors. The guide could likewise be mounted on fixed or rotatable arms. One such arrangement could be more favorable, for example, on offshore drilling platforms, mining platforms, manufacturing plants or Large ships at sea. For the normal transport of people or goods, these guic9es could likewise be made available for use in high buildings, or towers, as airport terminals.
The above-referenced technical advantages of aircraft according to the invention result in economic, logistical and strategic advantages in regard to conventional civilian and military flying operations.
Since the aircraft according to the invention can, in principle, transport the same freight weight or the same number of passengers as middle or long-range aircraft and also exhibits comparably high airspeeds and ranges, the aircraft according to invention should serve as a substantial competitor in the area of middle and long-range aircraft flight, while at the same time offering substantial ecological, economic and logistical advantages. The aircraft according to the invention can land on runways from a substantial height through vertical descending flight, and also take off in this manner. This avoids the well-known noise pollution of conventional airplanes in neighborhoods near runways.
The aircraft according to invention does not need a cost-intensive infrastructure such as, for example, airfields with expansive airplane runways.
Thus, the costs associated with aircraft according to the invention are reduced and the aircraft can fly into any desired city and its city center directly, even if they city doesn't include an airport. Networks of air routes could be developed with minimal infrastructure. This is particularly favorable for the economic development of countries that do not have the means to develop an infrastructure that includes conventional airports.
In long-distance flying operations (for example, over oceans), an aircraft according to the invention can, unlike conventional airplanes, land on motherships at sea for maintenance, refueling, or for an emergency landing.
The inventive aircraft can accomplish emergency landings at low speed on water or land and thereby prevent the usual destruction of airplanes that is associated with water-based emergency landings or the frequent destruction of airplanes that is associated with forced landings on land. The inventive aircraft is substantially safer than conventional middle and long-range airplanes.
Due to the high load, maneuverability and dockability of the inventive aircraft, salvage and rescue actions can be accomplished, which are not possible with conventional helicopters. With the invention, victims of the attack on the World Trade Center in New York could have been saved from the then-inaccessible floors of the building. With the invention, supplying or evacuating crisis or disaster areas can be accomplished better and faster than with currently available means of transportation.
In military applications, the inventive aircraft makes new, substantially more efficient operations and strategies possible. For example, large material or troop movements, which are currently possible only over slow, combined routEa of transportation using, for example, ships and/or large airplanes and/or land transportation, could be accomplished with aircraft according to the invention substantially faster via direct transport into the military target areas. The saving of time and money made possible by the inventive aircraft are of enormous military-strategic importance. For example, in the framework of miilitary operations, airfields will no longer need to be occupied and secured. High sea-going vessels can be supplied at any point on the open sea, without having to wait for, or initiate crossovers with, supply ships.

There are different possibilities of applying and expanding upon the teachings of the current invention in favorable ways. To this end, reference is made, on the one hand, to the enclosed claims, and on the other hand, to the following explanation of preferential examples of an aircraft according to the invention in view of the figures. In connection with the explanation of the preferred embodiments of the inventive aircraft, generally preferred embodiments and further aspects of the teachings are described in reference to the figures. The figures show:
Fig. 1 depicts schematic front -, back -, top- and side views of a first exemplary embodiment of an aircraft according to the invention, Fig. 2 depicts schematic front -, back -, top- and side views of a second exemplary embodiment of an aircraft according to the invention, Fig. 3 depicts schematic front -, top- and side views of the aircraft of Fig. 2 with a docking assembly disposed at the front part of the fuselage, Fig. 4 is a schematic cross section of an impeller blade of the propulsion device, Fig. 5 is a plan view of a drive disk for the impeller blades, Fig. 6 is a plan view of a control ring with a circular annular groove and an Eccentric disk guide, Fig. 7 is a plan view of the drive disk with the control ring shown (by broN;en lines) in its neutral position, Fig. 8 is a plan view of the drive disk with the control ring shown (also by broken lines) in an operating position, whereby the thrust runs in the direction of the arrow, Fig. 9 is a schematic side view of a further exemplary embodiment of an aircraft according to the invention, and Fig. 10 is a schematic front and rear view of the exemplary embodiment of Fig. 9.
Fig. 1 shows a first exemplary embodiment of an aircraft according to the invention in a schematic front -, back -, top and side views. The aircraft includes a fuselage 1 and a propulsion device 2 coupled with the fuselage 1 for the production of a definable lift. The propulsion device 2 includes a plurality of impeller blades 3, which are rotatable about a pivot axis 4 into a predetermined blade angle. The blades 3 are mounted to rotated about a rotation shaft 5 and the impeller blade angle is changeable during rotation for the production of the lift. Moreover, the respective pivot axes 4 of the impeller blades 3 are positioned substantially parallel to the rotation axis 5. The pivot axEa 4 of the impeller blades 3 are also positioned substantially parallel to each other.
Furthermore, the pivot axes 4 are positioned equidistant to each other and are disposed the same distance from the rotation axis 5.
The blades 3 are rotatably attached via their pivot axle 4 at one end to a drive component 7, or rotatably mounted within the drive component 7.
Each blade 3 includes a control shaft 6 as a point of attack for pivoting the impeller blades 3 around the pivot axle 4. The drive component 7 includes a bearing shaft 8.
In the exemplary embodiment shown here, impeller blades 3 form a rotor assembly 15 that is, respectively, rotatabfe about one of the rotation axes 5, whereby the propulsion device 2 includes a total of four such rotor assemblies 15. Two rotor assemblies 15 are positioned on each longitudinal sides of the fuselage 1. Thereby, the rotation axes 5 are aligned opposite the rotor assemblies 15.
Fig. 2 shows schematic front -, back -, top and side views of a second exemplary embodiment of an aircraft according to the invention. In this exemplary embodiment, rotor assemblies 15 are positioned parallel to a longitudinal axis of the fuselage 1 that extends in the forward flight direction.
The blades 3 of the rotor assemblies 15 are torsion-controlled and include a control member at both ends for controlling the control shaft.
There is a guidance rotor 17 positioned adjacent the guidance device 16 for forward movements or backward motions. The guidance rotor 17 is comprised of rotor blades 18. With the exemplary embodiment shown here, two wave achievement drive turbines are positioned overhead.
Fig. 3 shows schematic front -, top and side views of the exemplary embodiment of an aircraft shown in Fig. 2, whereby a docking assembly 19 is attached to the fuselage 1 for the loading or unloading of cargo and/or for the loading or unloading of people. The docking assembly 19 is designed as an escape tube.
Fig. 4 shows a schematic view of the cross-section profile of an impeller blade 3. The pivot axis 4 is recognizable on the one side, and the control shaft 6 is recognizable on the other side.
Fig. 5 shows a schematic plan view of a drive component 7 designed as a drive disk, which includes a bearing shaft 8. The drive component 7 inclludes passages 9 for receiving the pivot axles 4 of the impeller blades 3.
Moreover, the drive component 7 includes passages 10 for the control shafts 6 of the impeller blades 3. The passages 10 are designed as curved elongate holes. The drive component 7 features passages 11 for weight conservation.
Fig. 6 is a schematic plan view of a control member 12 that is in the forrn of a control ring having an annular groove 14 running within the area of the outside edge of the control ring for controlling the control shaft 6 of an impeller blade 3. The control member 12 is moveable within a guide that is in the form of an eccentric disk guide. This facilitates the movement of the control element 12 relative to the axis of rotation 5. In one exemplary embodiment, the control element 12 is positioned parallel to the drive component 7 so that the control shaft 6 of an impeller blade 3 extends through the passage 10 in the drive component 7 and into the annular groove 14 in the control element.
Fig. 7 shows a schematic plan view of an arrangement of the drive component 7 with the control member 12 of a rotor assembly 15 positioned behind the rotor assembly 15. The control member 12 is only represented by broken lines and only in its outer boundary region. The control member 12 in Fig. 7 is in its neutral position, whereby no thrust and no air diversion are produced by the impeller blades 3. The cross sectional profile of the impeNer blade 3 is concavely curved toward the rotation shaft 5. The blades 3 are essE;ntially arranged in an imaginary circular cylinder, which is produced by the curvature of the impeller blades 3.
In Fig. 8, the control member 12 is shifted relative to the rotation shaft 5 by means of the guide. A thrust is produced in the thrust direction 20. One can recognize the principle of the cyclic blade adjustment in Fig. 8 by means of the control member 12, whereby the impeller blades 3 are feathered between their extreme deflections one time during one rotation of the drive component 7 relative to the control member 12. The two extreme deflection positions of the impeller blades 3, are virtually located on a line running through the rotation axis 5 and defined by the thrust direction 20. When the positions are moved by 90 degrees, the impeller blades 3 are again in their neutral position, in which they do not produce a thrust or any air diversion.
The direction of rotation of the impeller blades 3 in the exemplary embodiment shown in Fig. 8 is clockwise.
Fig. 9 shows a schematic side view of a further exemplary embodiment of an aircraft according to invention with a fuselage 1, whereby a draft propeller 21 is positioned in front of a rotor assembly 15 and a pressure propeller 22 is positioned behind another a rotor assembly 15. Moreover, turbine outlets 23 are positioned adjacent the rotor assemblies 15.
Fig. 10 shows a schematic front and back view of the exemplary embodiment from Fig. 9, whereby airfoils 24 or auxiliary airfoils are positioned adjacent the fuselage 1. The rotor assemblies 15 are attached to or hung from the airfoils 24. The draft propellers 21 are positioned in front of the rotor assemblies 15 and the pressure propellers 22 are positioned behind the rotor assemblies 15.
Regarding further favorable embodiments of the inventive aircraft, in order to avoid repetition of the general parts of the description, one can refer to the enclosed patent claims.
Finally, it is expressly asserted that the examples of the aircraft according to invention described above are only for the purposes of discussion of the claimed device. However, these examples should not be regarded as limiting.

Claims

1. Aircraft with a fuselage (1) and a propulsion device (2) coupled to the fuselage (1) for the production of a definable lift, whereby the propulsion device (2) includes impeller blades (3) and whereby the impeller blades (3) are mounted so that they can be pivoted about a pivot axis (4) to a pre-determined blade angle and are rotatable about a rotation axis (5), whereby the impeller blade angle is adjustable during rotation to produce lift, whereby respective pivot axes (4) of the impeller blades (3) are positioned substantially parallel to the rotation axis (5), whereby the impeller blades (3) form a rotor assembly (15) that is rotatable about one of the rotation axes (5), and whereby the propulsion device (2) includes at least two such rotors, characterized in that a variable pitch propeller or propeller (21, 22) is positioned in front of as well as behind a rotor assembly.

2. The aircraft of Claim 1, characterized in that the respective pivot axes (4) of they impeller blades (3) are positioned substantially equidistant to one another.

3. The aircraft of Claims 1 or 2, characterized in that the pivot axes (4) of the impeller blades (3) are respectively positioned substantially the same distance to the axis of rotation (5).

4. The aircraft of Claims 1 to 3, characterized in that the pivot axes (4) of the impeller blades (3) are positioned substantially parallel to each other.

5. The aircraft of Claims 1 to 4, characterized in that the pivot axis (4) of each respective impeller blade (3) is positioned to pass through the centroid of the impelller blade (3).

6. The aircraft of Claims 1 to 5, characterized in that the cross sectional profile of the impeller blades (3) is concavely curved toward the axis of rotation (5).

7. The aircraft of Claims 1 to 6, characterized in that the impeller blades (3) comprise a control shaft (6) at at least one end that serves as the point of contact for swiveling the impeller blades (3) about the pivot axis (4).

8. The aircraft of Claims 1 to 7, characterized in that the impeller blades (3) are at rotatably mounted at one end adjacent or in a drive component (7).

9. The aircraft of Claim 8, characterized in that the drive component (7) is rotatable about the axis of rotation (5) or are rotatably mounted on the axis of rotation (5).

10. The aircraft of Claims 8 or 9, characterized in that the drive component (7) comprises a bearing shaft (8) or hollow shaft.

11. The aircraft of Claims 8 to 10, characterized in that the drive component (7) is a, drive pulley, drive disk or drive ring.

12. The aircraft of Claims 8 to 11, characterized in that the pivot axes (4) are arranged in a circle at the edge of the drive component (7) or the drive pulley or the drive disk or at the edge of the drive ring.

13. The aircraft of Claims 8 to 12, characterized in that the drive component (7) comprises recesses or passages (9) for receiving the pivot axes (4) of the impeller blades.

14. The aircraft of Claims 8 to 13, characterized in that the drive component (7) comprises recesses or passages (10) for the control shafts (6) of the impeller blades (3).

15. The aircraft of Claim 14, characterized in that the recesses or passages (10) for the control shafts (6) of the impeller blades (3) are constructed as long, preferably curved, holes.

16. The aircraft of Claims 8 to 15, characterized in that the drive component (7) comprises recesses, cutouts or passages (11).

17. The aircraft of Claims 8 to 16, characterized in that the drive component (7) works in conjunction with a control member (12) to adjust the impeller blades (3) about their pivot axis (4).

18. The aircraft of Claim 17, characterized in that the control member ( 12) is disengaged by the rotation of the impeller blades (3) and/or the drive component (7).

19. The aircraft of Claim 17 or 18, characterized in that the control member (12) is mounted on the axis of rotation (5).

20. The aircraft of Claims 17 to 19, characterized in that the control member (12) comprises a cyclical gear.

21. The aircraft of Claims 17 to 20, characterized in that the control member (12) is movable within a guide relative to the axis of rotation (5).

22. The aircraft of Claim 21, characterized in that the guide comprises two perpendicularly-arranged linear guides, in the form of a cross-table guide.

23. The aircraft of Claim 21, characterized in that the guide comprises a rotational guide that is connected to a linear guide in the form of an extendable and pivotable guide.

24. The aircraft of Claim 21, characterized in that the guide comprises two rotational control portions in the form of a double eccentric disk control member (13).

25. The aircraft of Claim 24, characterized in that two eccentric disks of the eccentric disk control member (13) are each respectively associated with an actuator.

26. The aircraft of Claims 17 to 25, characterized in that the control member (12) comprises an annular groove (14) or circular groove for receiving the impeller blades' (3) control shafts (6).

27. The aircraft of Claims 17 to 26, characterized in that the control member (12) is a control ring or a control disk.

28. The aircraft of Claims 26 or 27, characterized in that annular groove (14) or the control ring comprises a non-circular portion.

29. The aircraft of Claim 28, characterized in that the non-circular portion provides an angle of incidence function that is dependent on the angle of rotation, or provides a superposed angle of incidence function.

30. -The aircraft of Claim 29, characterized in that the angle of incidence function is proportional to the expression a-cos(x)w, in which "a" is the degree of the impeller blades' (3) angle of incidence and "w" is preferably a whole number, and preferably 11.

31. The aircraft of Claims 17 to 30, characterized in that the bearing shaft (8) or hollow shaft of the drive component (7) is preferably configured to run centrically through the control member (12).

32. The aircraft of Claims 17 to 31, characterized in that the drive component (7) and the control member (12) are positioned parallel to each other.

33. The aircraft of Claims 1 to 32, characterized in that the drive device (2) comprises at least two arrangements of rotatable impeller blades (3) about an axis of rotation (5).

34. The aircraft of Claims 1 to 33, characterized in that the axis of rotation (5) or axes of rotations (5) is or are positioned in a substantially horizontal plane.

35. The aircraft of Claims 1 to 34, characterized in that the axis of rotation (5) or axes of rotations (5) is/are positioned parallel to the longitudinal axis of the fuselage's (1) forward flight direction.

36. The aircraft of Claims 1 to 34, characterized in that the axis of rotation (5) or axes of rotations (5) is/are positioned perpendicular to a longitudinal axis of the fuselage's (1) forward flight direction.

37. The aircraft of Claim 1 to 36, characterized in that the rotors (15) are staggered along the longitudinal axis.

38. The aircraft of Claim 1 to 37, characterized in that at least one rotor (15) is positioned on each longitudinal side of the fuselage (1).

39. The aircraft of Claims 1 to 38, characterized in that at least two rotors (15) rotate in opposite directions from one another.

40. The aircraft of Claims 1 to 39, characterized in that at least two rotors (15) are located on each longitudinal side of the fuselage (1) and the rotation axes (5) of corresponding rotors (15) on respective longitudinal sides of the aircraft are in alignment.

41. The aircraft of Claims 1 to 40, characterized in that each rotor (15) is separately controllable.

42. The aircraft of Claims 1 to 41, characterized in that rotors (15) are controllable together in the same manner.

43. The aircraft of Claims 1 to 42, characterized in that the impeller blades (3) are featherably mounted, by means of a pivot shaft (4), at their end that is away from the drive component (7), at or in a guidance device (16).

44. The aircraft of Claim 43, characterized in that the guidance device (16) is in substantially the same form as the drive component (7) - preferably disk-shaped.

45. The aircraft of Claim 43 or 44, characterized in that the guidance device (16) is positioned to rotate with the drive component (7).

46. The aircraft of Claims 43 to 45, characterized in that the guidance device (16) and the drive component (7) are coupled by means of a shaft or the rotation shaft (5).

47. The aircraft of Claims 43 to 46, characterized in that a guidance device (16) is associated with a guidance rotor (17) comprising a plurality of rotor blades (18).

48. The aircraft of Claim 47, characterized in that the rotor blades (18) are linked to the guidance device (16).

49. The aircraft of Claim 47 or 48, characterized in that the guidance rotor (17) is propelled about a linkage running through the rotation axis (5).

50. The aircraft of Claims 1 to 49, characterized in that impeller blades (3) or rotors (15) that are positioned next to, or behind one another are positioned in a virtually mirrored arrangement.

51. The aircraft of Claims 17 to 50, characterized in that at both ends of the impeller blades (3) each control member (12) is operable independently of each of the other control members (12).

52. The aircraft of Claims 1 to 51, characterized in that at least one drive turbine is positioned in the fuselage (1) of the aircraft.

53. The aircraft of Claim 52, characterized in that the turbine exhaust gases are laterally directed from the fuselage (1), directly toward or over the rotor (15).

54. The aircraft of Claim 1 to 53, characterized in that between the variable-pitch propellers or propellers (21, 22) two or more rotor assemblies (15) are positioned one behind the other.

55. The aircraft of Claims 1 to 54, characterized in that an airfoil or auxiliary airfoil is attached adjacent the fuselage (1), and the rotor assembly (15) or assemblies are attached to, or hung from, the airfoil or auxiliary airfoil.